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Sample records for 22-foot subsonic tunnel

  1. Data Reduction Functions for the Langley 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Boney, Andy D.

    2014-01-01

    The Langley 14- by 22-Foot Subsonic Tunnel's data reduction software utilizes six major functions to compute the acquired data. These functions calculate engineering units, tunnel parameters, flowmeters, jet exhaust measurements, balance loads/model attitudes, and model /wall pressures. The input (required) variables, the output (computed) variables, and the equations and/or subfunction(s) associated with each major function are discussed.

  2. Langley 14- by 22-foot subsonic tunnel test engineer's data acquisition and reduction manual

    NASA Technical Reports Server (NTRS)

    Quinto, P. Frank; Orie, Nettie M.

    1994-01-01

    The Langley 14- by 22-Foot Subsonic Tunnel is used to test a large variety of aircraft and nonaircraft models. To support these investigations, a data acquisition system has been developed that has both static and dynamic capabilities. The static data acquisition and reduction system is described; the hardware and software of this system are explained. The theory and equations used to reduce the data obtained in the wind tunnel are presented; the computer code is not included.

  3. 14- by 22-Foot Subsonic Tunnel Laser Velocimeter Upgrade

    NASA Technical Reports Server (NTRS)

    Meyers, James F.; Lee, Joseph W.; Cavone, Angelo A.; Fletcher, Mark T.

    2012-01-01

    A long-focal length laser velocimeter constructed in the early 1980's was upgraded using current technology to improve usability, reliability and future serviceability. The original, free-space optics were replaced with a state-of-the-art fiber-optic subsystem which allowed most of the optics, including the laser, to be remote from the harsh tunnel environment. General purpose high-speed digitizers were incorporated in a standard modular data acquisition system, along with custom signal processing software executed on a desktop computer, served as the replacement for the signal processors. The resulting system increased optical sensitivity with real-time signal/data processing that produced measurement precisions exceeding those of the original system. Monte Carlo simulations, along with laboratory and wind tunnel investigations were used to determine system characteristics and measurement precision.

  4. The Langley 14- by 22-Foot Subsonic Tunnel: Description, Flow Characteristics, and Guide for Users

    NASA Technical Reports Server (NTRS)

    Gentry, Garl L., Jr.; Quinto, P. Frank; Gatlin, Gregory M.; Applin, Zachary T.

    1990-01-01

    The Langley 14- by 22-foot Subsonic Tunnel is a closed circuit, single-return atmospheric wind tunnel with a test section that can be operated in a variety of configurations (closed, slotted, partially open, and open). The closed test section configuration is 14.5 ft high by 21.75 ft wide and 50 ft long with a maximum speed of about 338 ft/sec. The open test section configuration has a maximum speed of about 270 ft/sec, and is formed by raising the ceiling and walls, to form a floor-only configuration. The tunnel may be configured with a moving-belt ground plane and a floor boundary-layer removal system at the entrance to the test section for ground effect testing. In addition, the tunnel had a two-component laser velocimeter, a frequency modulated (FM) tape system for dynamic data acquisition, flow visualization equipment, and acoustic testing capabilities. Users of the 14- by 22-foot Subsonic Tunnel are provided with information required for planning of experimental investigations including test hardware and model support systems.

  5. Free-Stream Turbulence Intensity in the Langley 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Neuhart, Dan H.; McGinley, Catherine B.

    2004-01-01

    An investigation was conducted using hot-wire anemometry to determine the turbulence intensity levels in the test section of the Langley 14- by 22-Foot Subsonic Tunnel in the closed or walls-down configuration. This study was one component of the three-dimensional High-Lift Flow Physics experiment designed to provide code validation data. Turbulence intensities were measured during two stages of the study. In the first stage, the free-stream turbulence levels were measured before and after a change was made to the floor suction surface of the wind tunnel s boundary layer removal system. The results indicated that the new suction surface at the entrance to the test section had little impact on the turbulence intensities. The second stage was an overall flow quality survey of the empty tunnel including measurements of the turbulence levels at several vertical and streamwise locations. Results indicated that the turbulence intensity is a function of tunnel dynamic pressure and the location in the test section. The general shape of the frequency spectrum is fairly consistent throughout the wind tunnel, changing mostly in amplitude (also slightly with frequency) with change in condition and location.

  6. Characterization of the Test Section Walls at the 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Lunsford, Charles B.; Graves, Sharon S.

    2003-01-01

    The test section walls of the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel are known to move under thermal and pressure loads. Videogrammetry was used to measure wall motion during the summer of 2002. In addition, a laser distancemeter was used to measure the relative distance between the test section walls at a single point. Distancemeter and videogrammetry results were consistent. Data were analyzed as a function of temperature and pressure to determine their effects on wall motion. Data were collected between 50 and 100 F, 0 and 0.315 Mach, and dynamic pressures of 0 and 120 psf. The overall motion of each wall was found to be less than 0.25 in. and less than facility personnel anticipated. The results show how motion depends on the temperature and pressure inside the test section as well is the position of the boundary layer vane. The repeatability of the measurements was +/-0.06 in. This report describes the methods used to record the motion of the test section walls and the results of the data analysis. Future facility plans include the development of a suitable wall restraint system and the determination of the effects of the wall motion on tunnel calibration.

  7. Trapezoidal Wing Experimental Repeatability and Velocity Profiles in the 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Hannon, Judith A.; Washburn, Anthony E.; Jenkins, Luther N.; Watson, Ralph D.

    2012-01-01

    The AIAA Applied Aerodynamics Technical Committee sponsored a High Lift Prediction Workshop held in June 2010. For this first workshop, data from the Trapezoidal Wing experiments were used for comparison to CFD. This paper presents long-term and short-term force and moment repeatability analyses for the Trapezoidal Wing model tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel. This configuration was chosen for its simplified high-lift geometry, publicly available set of test data, and previous CFD experience with this configuration. The Trapezoidal Wing is a three-element semi-span swept wing attached to a body pod. These analyses focus on configuration 1 tested in 1998 (Test 478), 2002 (Test 506), and 2003 (Test 513). This paper also presents model velocity profiles obtained on the main element and on the flap during the 1998 test. These velocity profiles are primarily at an angle of attack of 28 degrees and semi-span station of 83% and show confluent boundary layers and wakes.

  8. Development of a Microphone Phased Array Capability for the Langley 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Humphreys, William M.; Brooks, Thomas F.; Bahr, Christopher J.; Spalt, Taylor B.; Bartram, Scott M.; Culliton, William G.; Becker, Lawrence E.

    2014-01-01

    A new aeroacoustic measurement capability has been developed for use in open-jet testing in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22 tunnel). A suite of instruments has been developed to characterize noise source strengths, locations, and directivity for both semi-span and full-span test articles in the facility. The primary instrument of the suite is a fully traversable microphone phased array for identification of noise source locations and strengths on models. The array can be mounted in the ceiling or on either side of the facility test section to accommodate various test article configurations. Complementing the phased array is an ensemble of streamwise traversing microphones that can be placed around the test section at defined locations to conduct noise source directivity studies along both flyover and sideline axes. A customized data acquisition system has been developed for the instrumentation suite that allows for command and control of all aspects of the array and microphone hardware, and is coupled with a comprehensive data reduction system to generate information in near real time. This information includes such items as time histories and spectral data for individual microphones and groups of microphones, contour presentations of noise source locations and strengths, and hemispherical directivity data. The data acquisition system integrates with the 14x22 tunnel data system to allow real time capture of facility parameters during acquisition of microphone data. The design of the phased array system has been vetted via a theoretical performance analysis based on conventional monopole beamforming and DAMAS deconvolution. The performance analysis provides the ability to compute figures of merit for the array as well as characterize factors such as beamwidths, sidelobe levels, and source discrimination for the types of noise sources anticipated in the 14x22 tunnel. The full paper will summarize in detail the design of the instrumentation

  9. User's manual for the model interface and plugboard cabinets in the 14- by 22-foot subsonic tunnel

    NASA Technical Reports Server (NTRS)

    Askew, Robert B.; Quinto, P. Frank

    1994-01-01

    The primary method of connection between the wind tunnel model instrumentation and the data acquisition system in the 14- by 22-Foot Subsonic Tunnel is through the Model Interface (MIF) and Plugboard cabinets. The MIF and Plugboard cabinets allow versatility in the connection of the instrumentation to the different data systems in the facility. The User's Manual describes the components inside the MIF cabinet, the input and output of the MIF, and the MIF patchboard, and the Plugboard cabinets. There are examples of standard connections for most of the instrumentation used in the facility.

  10. User's manual for the Langley Research Center 14- by 22- foot subsonic tunnel static data acquisition system

    NASA Technical Reports Server (NTRS)

    Orie, Nettie M.; Quinto, P. Frank

    1993-01-01

    The Static Data Acquisition System (SDAS) components primarily responsible for acquiring data at the 14- by 22-Foot Subsonic Tunnel are the NEFF 620/600 Data Acquisition Unit (DAU) and the PSI 780B electronically scanned pressure (ESP) measurement system. A 9250 Modcomp computer is used to process the signal, to do all aerodynamic calculation, and to control the output of data. All of the tasks required to support a wind tunnel investigation are menu driven. The purpose of this report is to acquaint users of this system with the wide range of capabilities that exist with the available hardware and software and provide them with the proper procedures to follow when setting up or running individual tests.

  11. Space Launch System Liftoff and Transition Aerodynamic Characterization in the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Pinier, Jeremy T.; Erickson, Gary E.; Paulson, John W.; Tomek, William G.; Bennett, David W.; Blevins, John A.

    2015-01-01

    A 1.75% scale force and moment model of the Space Launch System was tested in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel to quantify the aerodynamic forces that will be experienced by the launch vehicle during its liftoff and transition to ascent flight. The test consisted of two parts: the first was dedicated to measuring forces and moments for the entire range of angles of attack (0deg to 90deg) and roll angles (0 deg. to 360 deg.). The second was designed to measure the aerodynamic effects of the liftoff tower on the launch vehicle for ground winds from all azimuthal directions (0 deg. to 360 deg.), and vehicle liftoff height ratios from 0 to 0.94. This wind tunnel model also included a set of 154 surface static pressure ports. Details on the experimental setup, and results from both parts of testing are presented, along with a description of how the wind tunnel data was analyzed and post-processed in order to develop an aerodynamic database. Finally, lessons learned from experiencing significant dynamics in the mid-range angles of attack due to steady asymmetric vortex shedding are presented.

  12. Acoustic Data Processing and Transient Signal Analysis for the Hybrid Wing Body 14- by 22-Foot Subsonic Wind Tunnel Test

    NASA Technical Reports Server (NTRS)

    Bahr, Christopher J.; Brooks, Thomas F.; Humphreys, William M.; Spalt, Taylor B.; Stead, Daniel J.

    2014-01-01

    An advanced vehicle concept, the HWB N2A-EXTE aircraft design, was tested in NASA Langley's 14- by 22-Foot Subsonic Wind Tunnel to study its acoustic characteristics for var- ious propulsion system installation and airframe con gurations. A signi cant upgrade to existing data processing systems was implemented, with a focus on portability and a re- duction in turnaround time. These requirements were met by updating codes originally written for a cluster environment and transferring them to a local workstation while en- abling GPU computing. Post-test, additional processing of the time series was required to remove transient hydrodynamic gusts from some of the microphone time series. A novel automated procedure was developed to analyze and reject contaminated blocks of data, under the assumption that the desired acoustic signal of interest was a band-limited sta- tionary random process, and of lower variance than the hydrodynamic contamination. The procedure is shown to successfully identify and remove contaminated blocks of data and retain the desired acoustic signal. Additional corrections to the data, mainly background subtraction, shear layer refraction calculations, atmospheric attenuation and microphone directivity corrections, were all necessary for initial analysis and noise assessments. These were implemented for the post-processing of spectral data, and are shown to behave as expected.

  13. Wing pressure distributions from subsonic tests of a high-wing transport model. [in the Langley 14- by 22-Foot Subsonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Applin, Zachary T.; Gentry, Garl L., Jr.; Takallu, M. A.

    1995-01-01

    A wind tunnel investigation was conducted on a generic, high-wing transport model in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data that document effects of various model configurations and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a part-span, double-slotted trailing-edge flap system. The trailing-edge flap was tested at four different deflection angles (20 deg, 30 deg, 40 deg, and 60 deg). Four wing configurations were tested: cruise, flaps only, Krueger flap only, and high lift (Krueger flap and flaps deployed). Tests were conducted at free-stream dynamic pressures of 20 psf to 60 psf with corresponding chord Reynolds numbers of 1.22 x 10(exp 6) to 2.11 x 10(exp 6) and Mach numbers of 0.12 to 0.20. The angles of attack presented range from 0 deg to 20 deg and were determined by wing configuration. The angle of sideslip ranged from minus 20 deg to 20 deg. In general, pressure distributions were relatively insensitive to free-stream speed with exceptions primarily at high angles of attack or high flap deflections. Increasing trailing-edge Krueger flap significantly reduced peak suction pressures and steep gradients on the wing at high angles of attack. Installation of the empennage had no effect on wing pressure distributions. Unpowered engine nacelles reduced suction pressures on the wing and the flaps.

  14. Development of a Large Field-of-View PIV System for Rotorcraft Testing in the 14- x 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Jenkins, Luther N.; Yao, Chung-Sheng; Bartram, Scott M.; Harris, Jerome; Allan, Brian; Wong, Oliver; Mace, W. Derry

    2009-01-01

    A Large Field-of-View Particle Image Velocimetry (LFPIV) system has been developed for rotor wake diagnostics in the 14-by 22-Foot Subsonic Tunnel. The system has been used to measure three components of velocity in a plane as large as 1.524 meters by 0.914 meters in both forward flight and hover tests. Overall, the system performance has exceeded design expectations in terms of accuracy and efficiency. Measurements synchronized with the rotor position during forward flight and hover tests have shown that the system is able to capture the complex interaction of the body and rotor wakes as well as basic details of the blade tip vortex at several wake ages. Measurements obtained with traditional techniques such as multi-hole pressure probes, Laser Doppler Velocimetry (LDV), and 2D Particle Image Velocimetry (PIV) show good agreement with LFPIV measurements.

  15. Infrared Images of Boundary Layer Transition on the D8 Transport Configuration in the LaRC 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Mason, Michelle L.; Gatlin, Gregory M.

    2015-01-01

    Grit, trip tape, or trip dots are routinely applied on the leading-edge regions of the fuselage, wings, tails or nacelles of wind tunnel models to trip the flow from laminar to turbulent. The thickness of the model's boundary layer is calculated for nominal conditions in the wind tunnel test to determine the effective size of the trip dots, but the flow over the model may not transition as intended for runs with different flow conditions. Temperature gradients measured with an infrared camera can be used to detect laminar to turbulent boundary layer transition on a wind tunnel model. This non-intrusive technique was used in the NASA Langley 14- by 22-Foot Subsonic Tunnel to visualize the behavior of the flow over a D8 transport configuration model. As the flow through the wind tunnel either increased to or decreased from the run conditions, a sufficient temperature difference existed between the air and the model to visualize the transition location (due to different heat transfer rates through the laminar and the turbulent boundary layers) for several runs in this test. Transition phenomena were visible without active temperature control in the atmospheric wind tunnel, whether the air was cooler than the model or vice-versa. However, when the temperature of the model relative to the air was purposely changed, the ability to detect transition in the infrared images was enhanced. Flow characteristics such as a wing root horseshoe vortex or the presence of fore-body vortical flows also were observed in the infrared images. The images of flow features obtained for this study demonstrate the usefulness of current infrared technology in subsonic wind tunnel tests.

  16. Laser velocimeter data acquisition system for the Langley 14- by 22-foot subsonic tunnel. Software reference guide version 3.3

    NASA Technical Reports Server (NTRS)

    Jumper, Judith K.

    1994-01-01

    The Laser Velocimeter Data Acquisition System (LVDAS) in the Langley 14- by 22-Foot Tunnel is controlled by a comprehensive software package. The software package was designed to control the data acquisition process during wind tunnel tests which employ a laser velocimeter measurement system. This report provides detailed explanations on how to configure and operate the LVDAS system to acquire laser velocimeter and static wind tunnel data.

  17. Subsonic aerodynamic characteristic of semispan commercial transport model with wing-mounted advanced ducted propeller operating in reverse thrust. [conducted in the Langley 14 by 22 foot subsonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Applin, Zachary T.; Jones, Kenneth M.; Gile, Brenda E.; Quinto, P. Frank

    1994-01-01

    A test was conducted in the Langley 14 by 22 Foot Subsonic Tunnel to determine the effect of the reverse-thrust flow field of a wing-mounted advanced ducted propeller on the aerodynamic characteristics of a semispan subsonic high-lift transport model. The advanced ducted propeller (ADP) model was mounted separately in position alongside the wing so that only the aerodynamic interference of the propeller and nacelle affected the aerodynamic performance of the transport model. Mach numbers ranged from 0.14 to 0.26; corresponding Reynolds numbers ranged from 2.2 to 3.9 x 10(exp 6). The reverse-thrust flow field of the ADP shielded a portion of the wing from the free-stream airflow and reduced both lift and drag. The reduction in lift and drag was a function of ADP rotational speed and free-stream velocity. Test results included ground effects data for the transport model and ADP configuration. The ground plane caused a beneficial increase in drag and an undesirable slight increase in lift. The ADP and transport model performance in ground effect was similar to performance trends observed for out of ground effect. The test results form a comprehensive data set that supports the application of the ADP engine and airplane concept on the next generation of advanced subsonic transports. Before this investigation, the engine application was predicted to have detrimental ground effect characteristics. Ground effect test measurements indicated no critical problems and were the first step in proving the viability of this engine and airplane configuration.

  18. Deployment of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades in Simulated Forward Flight: Preliminary PSP Results from Test 581 in the 14- by 22-Foot Subsonic Tunnel

    NASA Technical Reports Server (NTRS)

    Watkins, Anthony Neal; Leighty, Bradley D.; Lipford, William E.; Wong, Oliver D.; Goodman, Kyle Z.; Crafton, James; Forlines, Alan; Goss, Larry; Gregory, James W.; Juliano, Thomas J.

    2011-01-01

    This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. This report will also detail possible improvements to the system.

  19. SACCON Forced Oscillation Tests at DNW-NWB and NASA Langley 14x22-Foot Tunnel

    NASA Technical Reports Server (NTRS)

    Vicroy Dan D.; Loeser, Thomas D.; Schuette, Andreas

    2010-01-01

    A series of three wind tunnel static and forced oscillation tests were conducted on a generic unmanned combat air vehicle (UCAV) geometry. These tests are part of an international research effort to assess the state-of-the-art of computational fluid dynamics (CFD) methods to predict the static and dynamic stability and control characteristics. The experimental dataset includes not only force and moment time histories but surface pressure and off body particle image velocimetry measurements as well. The extent of the data precludes a full examination within the scope of this paper. This paper provides some examples of the dynamic force and moment data available as well as some of the observed trends.

  20. 7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  1. 5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  2. 1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  3. 3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  4. SUBSONIC WIND TUNNEL PERFORMANCE ANALYSIS SOFTWARE

    NASA Technical Reports Server (NTRS)

    Eckert, W. T.

    1994-01-01

    This program was developed as an aid in the design and analysis of subsonic wind tunnels. It brings together and refines previously scattered and over-simplified techniques used for the design and loss prediction of the components of subsonic wind tunnels. It implements a system of equations for determining the total pressure losses and provides general guidelines for the design of diffusers, contractions, corners and the inlets and exits of non-return tunnels. The algorithms used in the program are applicable to compressible flow through most closed- or open-throated, single-, double- or non-return wind tunnels or ducts. A comparison between calculated performance and that actually achieved by several existing facilities produced generally good agreement. Any system through which air is flowing which involves turns, fans, contractions etc. (e.g., an HVAC system) may benefit from analysis using this software. This program is an update of ARC-11138 which includes PC compatibility and an improved user interface. The method of loss analysis used by the program is a synthesis of theoretical and empirical techniques. Generally, the algorithms used are those which have been substantiated by experimental test. The basic flow-state parameters used by the program are determined from input information about the reference control section and the test section. These parameters were derived from standard relationships for compressible flow. The local flow conditions, including Mach number, Reynolds number and friction coefficient are determined for each end of each component or section. The loss in total pressure caused by each section is calculated in a form non-dimensionalized by local dynamic pressure. The individual losses are based on the nature of the section, local flow conditions and input geometry and parameter information. The loss forms for typical wind tunnel sections considered by the program include: constant area ducts, open throat ducts, contractions, constant

  5. Investigation of a Technique for Measuring Dynamic Ground Effect in a Subsonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Graves, Sharon S.

    1999-01-01

    To better understand the ground effect encountered by slender wing supersonic transport aircraft, a test was conducted at NASA Langley Research Center's 14 x 22 foot Subsonic Wind Tunnel in October, 1997. Emphasis was placed on improving the accuracy of the ground effect data by using a "dynamic" technique in which the model's vertical motion was varied automatically during wind-on testing. This report describes and evaluates different aspects of the dynamic method utilized for obtaining ground effect data in this test. The method for acquiring and processing time data from a dynamic ground effect wind tunnel test is outlined with details of the overall data acquisition system and software used for the data analysis. The removal of inertial loads due to sting motion and the support dynamics in the balance force and moment data measurements of the aerodynamic forces on the model is described. An evaluation of the results identifies problem areas providing recommendations for future experiments. Test results are validated by comparing test data for an elliptical wing planform with an Elliptical wing planform section with a NACA 0012 airfoil to results found in current literature. Major aerodynamic forces acting on the model in terms of lift curves for determining ground effect are presented. Comparisons of flight and wind tunnel data for the TU-144 are presented.

  6. Acoustic measurement study 40 by 80 foot subsonic wind tunnel

    NASA Technical Reports Server (NTRS)

    1974-01-01

    An acoustical study conducted during the period from September 1, 1973 to April 30, 1974 measured sound pressure levels and vibration amplitudes inside and outside of the subsonic tunnel and on the tunnel structure. A discussion of the technical aspects of the study, the field measurement and data reduction procedures, and results are presentd, and conclusions resulting from the study which bear upon near field and far field tunnel noise, upon the tunnel as an acoustical enclosure, and upon the sources of noise within the tunnel drive system are given.

  7. Analysis of Post-Support and Wind-Tunnel Wall Interference on Flow Field About Subsonic High-Lift High-Speed Research Configuration

    NASA Technical Reports Server (NTRS)

    Lessard, Wendy B.

    2000-01-01

    The present study was performed to determine how significant the interference effects of the wind-tunnel model support system and tunnel walls can be for a high-speed configuration during takeoff and landing conditions. A 5-percent scale model of the Technology Concept Airplane was recently tested in the Langley 14- by 22-Foot Sub-sonic Tunnel. The model was numerically modeled with and without the support and tunnel walls and compared with experimental data. Detailed analysis of the flow provided additional insight concerning what effects the post support and tunnel walls had on the flow field. This study revealed that although the overall forces and moments could be experimentally accounted for, the detailed flow features, such as the surface pressure distributions, could not be accurately simulated without including the post support in the computations.

  8. The requirements for a new full scale subsonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Kelly, M. W.; Mckinney, M. O.; Luidens, R. W.

    1972-01-01

    Justification and requirements are presented for a large subsonic wind tunnel capable of testing full scale aircraft, rotor systems, and advanced V/STOL propulsion systems. The design considerations and constraints for such a facility are reviewed, and the trades between facility test capability and costs are discussed.

  9. Mitigation of wind tunnel wall interactions in subsonic cavity flows

    SciTech Connect

    Wagner, Justin L.; Casper, Katya Marie; Beresh, Steven J.; Henfling, John F.; Spillers, Russell Wayne; Pruett, Brian Owen Matthew

    2015-03-06

    In this study, the flow over an open aircraft bay is often represented in a wind tunnel with a cavity. In flight, this flow is unconfined, though in experiments, the cavity is surrounded by wind tunnel walls. If untreated, wind tunnel wall effects can lead to significant distortions of cavity acoustics in subsonic flows. To understand and mitigate these cavity–tunnel interactions, a parametric approach was taken for flow over an L/D = 7 cavity at Mach numbers 0.6–0.8. With solid tunnel walls, a dominant cavity tone was observed, likely due to an interaction with a tunnel duct mode. Furthermore, an acoustic liner opposite the cavity decreased the amplitude of the dominant mode and its harmonics, a result observed by previous researchers. Acoustic dampeners were also placed in the tunnel sidewalls, which further decreased the dominant mode amplitudes and peak amplitudes associated with nonlinear interactions between cavity modes. This then indicates that cavity resonance can be altered by tunnel sidewalls and that spanwise coupling should be addressed when conducting subsonic cavity experiments. Though mechanisms for dominant modes and nonlinear interactions likely exist in unconfined cavity flows, these effects can be amplified by the wind tunnel walls.

  10. Mitigation of wind tunnel wall interactions in subsonic cavity flows

    DOE PAGES

    Wagner, Justin L.; Casper, Katya Marie; Beresh, Steven J.; ...

    2015-03-06

    In this study, the flow over an open aircraft bay is often represented in a wind tunnel with a cavity. In flight, this flow is unconfined, though in experiments, the cavity is surrounded by wind tunnel walls. If untreated, wind tunnel wall effects can lead to significant distortions of cavity acoustics in subsonic flows. To understand and mitigate these cavity–tunnel interactions, a parametric approach was taken for flow over an L/D = 7 cavity at Mach numbers 0.6–0.8. With solid tunnel walls, a dominant cavity tone was observed, likely due to an interaction with a tunnel duct mode. Furthermore, anmore » acoustic liner opposite the cavity decreased the amplitude of the dominant mode and its harmonics, a result observed by previous researchers. Acoustic dampeners were also placed in the tunnel sidewalls, which further decreased the dominant mode amplitudes and peak amplitudes associated with nonlinear interactions between cavity modes. This then indicates that cavity resonance can be altered by tunnel sidewalls and that spanwise coupling should be addressed when conducting subsonic cavity experiments. Though mechanisms for dominant modes and nonlinear interactions likely exist in unconfined cavity flows, these effects can be amplified by the wind tunnel walls.« less

  11. Effect of collector configuration on test section turbulence levels in an open-jet wind tunnel

    NASA Technical Reports Server (NTRS)

    Manuel, G. S.; Molloy, John K.; Barna, P. Stephen

    1992-01-01

    Flow quality studies in the Langley 14- by 22-Foot Subsonic Tunnel indicated periodic flow pulsation at discrete frequencies in the test section when the tunnel operated in an open-jet configuration. To alleviate this problem, experiments were conducted in a 1/24-scale model of the full-scale tunnel to evaluate the turbulence reduction potential of six collector configurations. As a result of these studies, the original bell-mouth collector of the 14- by 22-Foot Subsonic Tunnel was replaced by a collector with straight walls, and a slot was incorporated between the trailing edge of the collector and the entrance of the diffuser.

  12. A review of technologies applicable to low-speed flight of high-performance aircraft investigated in the Langley 14- x 22-foot subsonic tunnel

    NASA Technical Reports Server (NTRS)

    Paulson, John W., Jr.; Quinto, P. Frank; Banks, Daniel W.; Kemmerly, Guy T.; Gatlin, Gregory M.

    1988-01-01

    An extensive research program has been underway at the NASA Langley Research Center to define and develop the technologies required for low-speed flight of high-performance aircraft. This 10-year program has placed emphasis on both short takeoff and landing (STOL) and short takeoff and vertical landing (STOVL) operations rather than on regular up and away flight. A series of NASA in-house as well as joint projects have studied various technologies including high lift, vectored thrust, thrust-induced lift, reversed thrust, an alternate method of providing trim and control, and ground effects. These technologies have been investigated on a number of configurations ranging from industry designs for advanced fighter aircraft to generic wing-canard research models. Test conditions have ranged from hover (or static) through transition to wing-borne flight at angles of attack from -5 to 40 deg at representative thrust coefficients.

  13. A lumped parameter mathematical model for simulation of subsonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Krosel, S. M.; Cole, G. L.; Bruton, W. M.; Szuch, J. R.

    1986-01-01

    Equations for a lumped parameter mathematical model of a subsonic wind tunnel circuit are presented. The equation state variables are internal energy, density, and mass flow rate. The circuit model is structured to allow for integration and analysis of tunnel subsystem models which provide functions such as control of altitude pressure and temperature. Thus the model provides a useful tool for investigating the transient behavior of the tunnel and control requirements. The model was applied to the proposed NASA Lewis Altitude Wind Tunnel (AWT) circuit and included transfer function representations of the tunnel supply/exhaust air and refrigeration subsystems. Both steady state and frequency response data are presented for the circuit model indicating the type of results and accuracy that can be expected from the model. Transient data for closed loop control of the tunnel and its subsystems are also presented, demonstrating the model's use as a control analysis tool.

  14. Unsteady two dimensional airloads acting on oscillating thin airfoils in subsonic ventilated wind tunnels

    NASA Technical Reports Server (NTRS)

    Fromme, J.; Golberg, M.

    1978-01-01

    The numerical calculation of unsteady two dimensional airloads which act upon thin airfoils in subsonic ventilated wind tunnels was studied. Neglecting certain quadrature errors, Bland's collocation method is rigorously proved to converge to the mathematically exact solution of Bland's integral equation, and a three way equivalence was established between collocation, Galerkin's method and least squares whenever the collocation points are chosen to be the nodes of the quadrature rule used for Galerkin's method. A computer program displayed convergence with respect to the number of pressure basis functions employed, and agreement with known special cases was demonstrated. Results are obtained for the combined effects of wind tunnel wall ventilation and wind tunnel depth to airfoil chord ratio, and for acoustic resonance between the airfoil and wind tunnel walls. A boundary condition is proposed for permeable walls through which mass flow rate is proportional to pressure jump.

  15. Optimized aerodynamic design process for subsonic transport wing fitted with winglets. [wind tunnel model

    NASA Technical Reports Server (NTRS)

    Kuhlman, J. M.

    1979-01-01

    The aerodynamic design of a wind-tunnel model of a wing representative of that of a subsonic jet transport aircraft, fitted with winglets, was performed using two recently developed optimal wing-design computer programs. Both potential flow codes use a vortex lattice representation of the near-field of the aerodynamic surfaces for determination of the required mean camber surfaces for minimum induced drag, and both codes use far-field induced drag minimization procedures to obtain the required spanloads. One code uses a discrete vortex wake model for this far-field drag computation, while the second uses a 2-D advanced panel wake model. Wing camber shapes for the two codes are very similar, but the resulting winglet camber shapes differ widely. Design techniques and considerations for these two wind-tunnel models are detailed, including a description of the necessary modifications of the design geometry to format it for use by a numerically controlled machine for the actual model construction.

  16. Low-Speed Wind Tunnel Tests of Two Waverider Configuration Models

    NASA Technical Reports Server (NTRS)

    Pegg, Robert J.; Hahne, David E.; Cockrell,Charles E., Jr.

    1995-01-01

    A definitive measurement of the low-speed flight characteristics of waverider-based aircraft is required to augment the overall design database for this important class of vehicles which have great potential for efficient high-speed flight. Two separate waverider-derived vehicles were tested; one in the 14- by 22-Foot Tunnel and the other in the 12-Foot Low Speed Tunnel at Langley Research Center. These tests provided measurements of moments and forces about all three axes, control effectiveness, flow field characteristics and the effects of configuration changes. The results of these tunnel tests are summarized and the subsonic aerodynamic characteristics of the two configurations are shown.

  17. Two dimensional aerodynamic interference effects on oscillating airfoils with flaps in ventilated subsonic wind tunnels. [computational fluid dynamics

    NASA Technical Reports Server (NTRS)

    Fromme, J.; Golberg, M.; Werth, J.

    1979-01-01

    The numerical computation of unsteady airloads acting upon thin airfoils with multiple leading and trailing-edge controls in two-dimensional ventilated subsonic wind tunnels is studied. The foundation of the computational method is strengthened with a new and more powerful mathematical existence and convergence theory for solving Cauchy singular integral equations of the first kind, and the method of convergence acceleration by extrapolation to the limit is introduced to analyze airfoils with flaps. New results are presented for steady and unsteady flow, including the effect of acoustic resonance between ventilated wind-tunnel walls and airfoils with oscillating flaps. The computer program TWODI is available for general use and a complete set of instructions is provided.

  18. Wind tunnel investigations of forebody strakes for yaw control on F/A-18 model at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Murri, Daniel G.

    1993-01-01

    Wind tunnel investigations have been conducted of forebody strakes for yaw control on 0.06-scale models of the F/A-18 aircraft at free-stream Mach numbers of 0.20 to 0.90. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center and the Langley 7- by 10-Foot High-Speed Tunnel. The principal objectives of the testing were to determine the effects of the Mach number and the strake plan form on the strake yaw control effectiveness and the corresponding strake vortex induced flow field. The wind tunnel model configurations simulated an actuated conformal strake deployed for maximum yaw control at high angles of attack. The test data included six-component forces and moments on the complete model, surface static pressure distributions on the forebody and wing leading-edge extensions, and on-surface and off-surface flow visualizations. The results from these studies show that the strake produces large yaw control increments at high angles of attack that exceed the effect of conventional rudders at low angles of attack. The strake yaw control increments diminish with increasing Mach number but continue to exceed the effect of rudder deflection at angles of attack greater than 30 degrees. The character of the strake vortex induced flow field is similar at subsonic and transonic speeds. Cropping the strake planform to account for geometric and structural constraints on the F-18 aircraft has a small effect on the yaw control increments at subsonic speeds and no effect at transonic speeds.

  19. Subsonic Cascade Wind Tunnel Tests Using a Compressor Configuration of DCA Blades.

    DTIC Science & Technology

    1981-06-01

    1399 - .257 1 4 -I,. ’- 𔃾 1 4 4930 1432 -. 1653 15:;7 .5669 1412 -. 1295 - 7746407 13.3𔄃a 0 5-3,’ 175139 -. 8 , .. 7 4 * ,r I t - 7146 .1289 . 0833 ...EDoNa OF I NOV 61 IS OBSOLETE (Page 1 ) s’w4 0 10 2-014- 6601 r1NCT.AA.9T~’TE S SECURITY CLASSIFICATION OF THIS PAGE (She..n DotaEnter*ed) UNCLASSIFIED...6 LIST OF FIGURES Figure 1 . Subsonic Cascade Facility ------------------------ 43 2. Photograph of Subsonic Cascade and Test Section -- 44 3

  20. Wind tunnel investigation of the interaction and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    1991-01-01

    The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.

  1. Comparisons of subsonic drag estimates derived from Pioneer Venus probes flight data with wind-tunnel results

    NASA Technical Reports Server (NTRS)

    Blanchard, R. C.; Phillips, W. P.; Kelly, G. M.; Findlay, J. T.

    1980-01-01

    Subsonic drag coefficients have been obtained from flight data for the Pioneer Venus multiprobes. The technique used to extract the information from the data consisted of utilizing in situ pressure and temperature measurements. Analysis of the major model parameter error sources indicates overall error levels of five percent or less in the flight values of the drag coefficient. Comparisons of the flight coefficients with preflight wind-tunnel test data showed generally good agreement except for the Sounder descent probe configuration. To preclude atmospheric phenomena as a possible explanation of this difference, additional wind-tunnel tests were performed on the Sounder descent probe. Special attempts were made to duplicate the probe geometry for tests in a high Reynolds number environment in order to achieve as realistic model and flight conditions as practical. Preliminary results from this testing in the NASA LaRC Low Turbulence Pressure Tunnel produced a drag coefficient of 0.68 at 0 deg angle of attack which is within the expected accuracy limits of the flight derived drag coefficient value of 0.72 + or - 0.04, thus eliminating atmospheric phenomena as the explanation for the initial difference.

  2. Wind tunnel investigation of vortex flows on F/A-18 configuration at subsonic through transonic speed

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    1991-01-01

    A wind tunnel experiment was conducted in the David Taylor Research Center 7- by 10-Foot Transonic Tunnel of the wing leading-edge extension (LEX) and forebody vortex flows at subsonic and transonic speeds about a 0.06-scale model of the F/A-18. The primary goal was to improve the understanding and control of the vortical flows, including the phenomena of vortex breakdown and vortex interactions with the vertical tails. Laser vapor screen flow visualizations, LEX, and forebody surface static pressures, and six-component forces and moments were obtained at angles of attack of 10 to 50 degrees, free-stream Mach numbers of 0.20 to 0.90, and Reynolds numbers based on the wing mean aerodynamic chord of 0.96 x 10(exp 6) to 1.75 x 10(exp 6). The wind tunnel results were correlated with in-flight flow visualizations and handling qualities trends obtained by NASA using an F-18 High-Alpha Research Vehicle (HARV) and by the Navy and McDonnell Douglas on F-18 aircraft with LEX fences added to improve the vertical tail buffet environment. Key issues that were addressed include the sensitivity of the vortical flows to the Reynolds number and Mach number; the reduced vertical tail excitation, and the corresponding flow mechanism, in the presence of the LEX fence; the repeatability of data obtained during high angle-of-attack wind tunnel testing of F-18 models; the effects of particle seeding for flow visualization on the quantitative model measurements; and the interpretation of off-body flow visualizations obtained using different illumination and particle seeding techniques.

  3. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Faceted Missile Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2004-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the vortex-induced surface static pressures on a slender, faceted missile model at subsonic and transonic speeds. Global PSP calibrations were obtained using an in-situ method featuring the simultaneous electronically-scanned pressures (ESP) measurements. Both techniques revealed the significant influence leading-edge vortices on the surface pressure distributions. The mean error in the PSP measurements relative to the ESP data was approximately 0.6 percent at M(sub infinity)=0.70 and 2.6 percent at M(sub infinity)=0.90 and 1.20. The vortex surface pressure signatures obtained from the PSP and ESP techniques were correlated with the off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The on-surface and off-surface techniques were complementary, since each provided details of the vortex-dominated flow that were not clear or apparent in the other.

  4. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Faceted Missile Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2004-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the vortex-induced surface static pressures on a slender, faceted missile model at subsonic and transonic speeds. Satisfactory global calibrations of the PSP were obtained at =0.70, 0.90, and 1.20, angles of attack from 10 degrees to 20 degrees, and angles of sideslip of 0 and 2.5 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at 57 discrete locations on the model. Both techniques clearly revealed the significant influence on the surface pressure distributions of the vortices shed from the sharp, chine-like leading edges. The mean error in the PSP measurements relative to the ESP data was approximately 0.6 percent at M infinity =0.70 and 2.6 percent at M infinity =0.90 and 1.20. The vortex surface pressure signatures obtained from the PSP and ESP techniques were correlated with the off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The on-surface and off-surface techniques were complementary, since each provided details of the vortex-dominated flow that were not clear or apparent in the other.

  5. Aerodynamic design guidelines and computer program for estimation of subsonic wind tunnel performance

    NASA Technical Reports Server (NTRS)

    Eckert, W. T.; Mort, K. W.; Jope, J.

    1976-01-01

    General guidelines are given for the design of diffusers, contractions, corners, and the inlets and exits of non-return tunnels. A system of equations, reflecting the current technology, has been compiled and assembled into a computer program (a user's manual for this program is included) for determining the total pressure losses. The formulation presented is applicable to compressible flow through most closed- or open-throat, single-, double-, or non-return wind tunnels. A comparison of estimated performance with that actually achieved by several existing facilities produced generally good agreement.

  6. Comparison of Drop and Wind-Tunnel Experiments on Bomb Drag at High Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Gothert, B.

    1948-01-01

    The drag coefficients of bombs at high velocities velocity of fall was 97 percent of the speed of sound) (the highest are determined by drop tests and compared with measurements taken in the DVL high-speed closed wind tunnel and the open jet at AVA - Gottingen.

  7. Measurement of Separated Flow Structures Using a Multiple-Camera DPIV System. [conducted in the Langley Subsonic Basic Research Tunnel

    NASA Technical Reports Server (NTRS)

    Humphreys, William M., Jr.; Bartram, Scott M.

    2001-01-01

    A novel multiple-camera system for the recording of digital particle image velocimetry (DPIV) images acquired in a two-dimensional separating/reattaching flow is described. The measurements were performed in the NASA Langley Subsonic Basic Research Tunnel as part of an overall series of experiments involving the simultaneous acquisition of dynamic surface pressures and off-body velocities. The DPIV system utilized two frequency-doubled Nd:YAG lasers to generate two coplanar, orthogonally polarized light sheets directed upstream along the horizontal centerline of the test model. A recording system containing two pairs of matched high resolution, 8-bit cameras was used to separate and capture images of illuminated tracer particles embedded in the flow field. Background image subtraction was used to reduce undesirable flare light emanating from the surface of the model, and custom pixel alignment algorithms were employed to provide accurate registration among the various cameras. Spatial cross correlation analysis with median filter validation was used to determine the instantaneous velocity structure in the separating/reattaching flow region illuminated by the laser light sheets. In operation the DPIV system exhibited a good ability to resolve large-scale separated flow structures with acceptable accuracy over the extended field of view of the cameras. The recording system design provided enhanced performance versus traditional DPIV systems by allowing a variety of standard and non-standard cameras to be easily incorporated into the system.

  8. Subsonic wind-tunnel measurements of a slender wing-body configuration employing a vortex flap

    NASA Technical Reports Server (NTRS)

    Frink, Neal T.

    1987-01-01

    A wind tunnel study at Mach 0.4 was conducted for a slender wing-body configuration with a leading edge vortex flap of curved planform that is deflectable about a 74 degree swept hinge line. The basic data consist of a unique combination of longitudinal aerodynamic, surface pressure, and vortex flap hinge-moment measurements on a common model. The longitudinal aerodynamic, pressure and hinge-moment data are presented without analysis in tabular format. Plots of the tabulated pressure data are also given.

  9. Subsonic Investigation of a Leading-Edge Boundary Layer Control Suction System on a High-Speed Civil Transport Configuration

    NASA Technical Reports Server (NTRS)

    Campbell, Bryan A.; Applin, Zachary T.; Kemmerly, Guy T.; Coe, Paul L., Jr.; Owens, D. Bruce; Gile, Brenda E.; Parikh, Pradip G.; Smith, Don

    1999-01-01

    A wind tunnel investigation of a leading edge boundary layer control system was conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.08 to 0.27, with corresponding chord Reynolds numbers of 1.79 x 10(exp 6) to 5.76 x 10(exp 6). Variations in the amount of suction, as well as the size and location of the suction area, were tested with outboard leading edge flaps deflected 0 and 30 deg and trailing-edge flaps deflected 0 and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein.

  10. Spin-Tunnel Investigation of a 1/28-Scale Model of a Subsonic Attack Airplane

    NASA Technical Reports Server (NTRS)

    Lee, Henry A.; Healy, Frederick M.

    1964-01-01

    An investigation has been made of a 1/28-scale model of the Grumman A-6A airplane in the Langley spin tunnel. The erect spin and recovery characteristics of the model were determined for the flight design gross weight loading and for a loading with full internal fuel and empty external wing fuel tanks. The effects of extending slats and deflecting flaps were investigated. Inverted-spin and recovery characteristics of the model were determined for the flight design gross weight loading. The size of the spin-recovery tail parachute necessary to insure satisfactory spin-recovery was determined, and the effect of firing wing-mounded rockets during spins was investigated.

  11. Fiber-optic-based laser vapor screen flow visualization system for aerodynamic research in larger scale subsonic and transonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Inenaga, Andrew S.

    1994-01-01

    Laser vapor screen (LVS) flow visualization systems that are fiber-optic based were developed and installed for aerodynamic research in the Langley 8-Foot Transonic Pressure Tunnel and the Langley 7- by 10-Foot High Speed Tunnel. Fiber optics are used to deliver the laser beam through the plenum shell that surrounds the test section of each facility and to the light-sheet-generating optics positioned in the ceiling window of the test section. Water is injected into the wind tunnel diffuser section to increase the relative humidity and promote condensation of the water vapor in the flow field about the model. The condensed water vapor is then illuminated with an intense sheet of laser light to reveal features of the flow field. The plenum shells are optically sealed; therefore, video-based systems are used to observe and document the flow field. Operational experience shows that the fiber-optic-based systems provide safe, reliable, and high-quality off-surface flow visualization in smaller and larger scale subsonic and transonic wind tunnels. The design, the installation, and the application of the Langley Research Center (LaRC) LVS flow visualization systems in larger scale wind tunnels are highlighted. The efficiency of the fiber optic LVS systems and their insensitivity to wind tunnel vibration, the tunnel operating temperature and pressure variations, and the airborne contaminants are discussed.

  12. Computation of aerodynamic interference effects on oscillating airfoils with controls in ventilated subsonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Fromme, J. A.; Golberg, M. A.

    1979-01-01

    Lift interference effects are discussed based on Bland's (1968) integral equation. A mathematical existence theory is utilized for which convergence of the numerical method has been proved for general (square-integrable) downwashes. Airloads are computed using orthogonal airfoil polynomial pairs in conjunction with a collocation method which is numerically equivalent to Galerkin's method and complex least squares. Convergence exhibits exponentially decreasing error with the number n of collocation points for smooth downwashes, whereas errors are proportional to 1/n for discontinuous downwashes. The latter can be reduced to 1/n to the m+1 power with mth-order Richardson extrapolation (by using m = 2, hundredfold error reductions were obtained with only a 13% increase of computer time). Numerical results are presented showing acoustic resonance, as well as the effect of Mach number, ventilation, height-to-chord ratio, and mode shape on wind-tunnel interference. Excellent agreement with experiment is obtained in steady flow, and good agreement is obtained for unsteady flow.

  13. Subsonic Aerodynamic Assessment of Vortex Flow Management Devices on a High-Speed Civil Transport Configuration

    NASA Technical Reports Server (NTRS)

    Campbell, Bryan A.; Applin, Zachary T.; Kemmerly, Guy T.

    1999-01-01

    An experimental investigation of the effects of leading-edge vortex management devices on the subsonic performance of a high-speed civil transport (HSCT) configuration was conducted in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.14 to 0.27, with corresponding chord Reynolds numbers of 3.08 x 10 (sup 6) to 5.47 x 10 (sup 6). The test model was designed for a cruise Mach number of 2.7. During the subsonic high-lift phase of flight, vortical flow dominates the upper surface flow structure, and during vortex breakdown, this flow causes adverse pitch-up and a reduction of usable lift. The experimental results showed that the beneficial effects of small leading-edge vortex management devices located near the model reference center were insufficient to substantially affect the resulting aerodynamic forces and moments. However, devices located at or near the wiring apex region demonstrated potential for pitch control with little effect on overall lift.

  14. Comparison of wind tunnel test results at free stream Mach 0.7 with results from the Boeing TEA-230 subsonic flow method. [wing flow method tests

    NASA Technical Reports Server (NTRS)

    Mohn, L. W.

    1975-01-01

    The use of the Boeing TEA-230 Subsonic Flow Analysis method as a primary design tool in the development of cruise overwing nacelle configurations is presented. Surface pressure characteristics at 0.7 Mach number were determined by the TEA-230 method for a selected overwing flow-through nacelle configuration. Results of this analysis show excellent overall agreement with corresponding wind tunnel data. Effects of the presence of the nacelle on the wing pressure field were predicted accurately by the theoretical method. Evidence is provided that differences between theoretical and experimental pressure distributions in the present study would not result in significant discrepancies in the nacelle lines or nacelle drag estimates.

  15. Subsonic Wake Characterization of the Orion Capsule Using PIV in the Ames UPWT 11-foot Wind Tunnel (Invited)

    NASA Technical Reports Server (NTRS)

    Heineck, James T.; Ross, James C.; Yamauchi, Gloria K.

    2015-01-01

    The subsonic regime of Crew Capsule reentry has a very turbulent waker through which the Drogue Chutes must deploy. This presentation describes the particle image velocimetry measurement campaign used to help retire the risk.

  16. Theoretical and Experimental Investigation of the Effect of Tunnel Walls on the Forces on an Oscillating Airfoil in Two-Dimensional Subsonic Compressible Flow

    NASA Technical Reports Server (NTRS)

    Runyan, Harry L; Woolston, Donald S; Rainey, A Gerald

    1956-01-01

    This report presents a theoretical and experimental investigation of the effect of wind-tunnel walls on the air forces on an oscillating wing in two-dimensional subsonic compressible flow. A method of solving an integral equation which relates the downwash on a wing to the unknown loading is given, and some comparisons are made between the theoretical results and the experimental results. A resonance condition, which was predicted by theory in a previous report (NACA report 1150), is shown experimentally to exist. In addition, application of the analysis is made to a number of examples in order to illustrate the influence of walls due to variations in frequency of oscillation, Mach number , and ratio of tunnel height to wing semichord.

  17. Aeroacoustic Characterization of the NASA Ames Experimental Aero-Physics Branch 32- by 48-Inch Subsonic Wind Tunnel with a 24-Element Phased Microphone Array

    NASA Technical Reports Server (NTRS)

    Costanza, Bryan T.; Horne, William C.; Schery, S. D.; Babb, Alex T.

    2011-01-01

    The Aero-Physics Branch at NASA Ames Research Center utilizes a 32- by 48-inch subsonic wind tunnel for aerodynamics research. The feasibility of acquiring acoustic measurements with a phased microphone array was recently explored. Acoustic characterization of the wind tunnel was carried out with a floor-mounted 24-element array and two ceiling-mounted speakers. The minimum speaker level for accurate level measurement was evaluated for various tunnel speeds up to a Mach number of 0.15 and streamwise speaker locations. A variety of post-processing procedures, including conventional beamforming and deconvolutional processing such as TIDY, were used. The speaker measurements, with and without flow, were used to compare actual versus simulated in-flow speaker calibrations. Data for wind-off speaker sound and wind-on tunnel background noise were found valuable for predicting sound levels for which the speakers were detectable when the wind was on. Speaker sources were detectable 2 - 10 dB below the peak background noise level with conventional data processing. The effectiveness of background noise cross-spectral matrix subtraction was assessed and found to improve the detectability of test sound sources by approximately 10 dB over a wide frequency range.

  18. Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail Interactions on a Slender Wing at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2013-01-01

    A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed. 15.

  19. Considerations on the effect of wind-tunnel walls on oscillating air forces for two-dimensional subsonic compressible flow

    NASA Technical Reports Server (NTRS)

    Runyan, Harry L; Watkins, Charles E

    1953-01-01

    This report treats the effect of wind-tunnel walls on the oscillating two-dimensional air forces in a compressible medium. The walls are simulated by the usual method of placing images at appropriate distances above and below the wing. An important result shown is that, for certain conditions of wing frequency, tunnel height, and Mach number, the tunnel and wing may form a resonant system so that the forces on the wing are greatly changed from the condition of no tunnel walls. It is pointed out that similar conditions exist for three-dimensional flow in circular and rectangular tunnels and apparently, within certain Mach number ranges, in tunnels of nonuniform cross section or even in open tunnels or jets.

  20. Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Lamar, John E.; Obara, Clifford J.; Fisher, Bruce D.; Fisher, David F.

    2001-01-01

    Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

  1. Hybrid Wing Body Model Identification Using Forced-Oscillation Water Tunnel Data

    NASA Technical Reports Server (NTRS)

    Murphy, Patrick C.; Vicroy, Dan D.; Kramer, Brian; Kerho, Michael

    2014-01-01

    Static and dynamic testing of the NASA 0.7 percent scale Hybrid Wing Body (HWB) configuration was conducted in the Rolling Hills Research Corporation water tunnel to investigate aerodynamic behavior over a large range of angle-of-attack and to develop models that can predict aircraft response in nonlinear unsteady flight regimes. This paper reports primarily on the longitudinal axis results. Flow visualization tests were also performed. These tests provide additional static data and new dynamic data that complement tests conducted at NASA Langley 14- by 22-Foot Subsonic Tunnel. HWB was developed to support the NASA Environmentally Responsible Aviation Project goals of lower noise, emissions, and fuel burn. This study also supports the NASA Aviation Safety Program efforts to model and control advanced transport configurations in loss-of-control conditions.

  2. Development of pneumatic test techniques for subsonic high-lift and in-ground-effect wind tunnel investigations

    NASA Technical Reports Server (NTRS)

    Englar, Robert J.

    1994-01-01

    Wind tunnel evaluations of two-dimensional high-lift airfoils and of vehicles operating in ground effect near the tunnel floor require special test facilities and procedures. These are needed to avoid errors caused by proximity to the walls and interference from the wall boundary layers. Pneumatic test techniques and facilities were developed for GTRI aerodynamic research tunnels and calibrated to verify that these wall effects had been removed. The modified facilities were then employed to evaluate the aerodynamic characteristics of blown very-high-lift airfoils and of racing hydroplanes operating in ground effect at various levels above the floor. The pneumatic facilities, techniques and calibrations are discussed and typical aerodynamic data recorded both with and without the test-section blowing systems are presented.

  3. Space Shuttle Orbiter trimmed center-of-gravity extension study. Volume 7: Effects of configuration modifications on the subsonic aerodynamic characteristics of the 1140 A/B orbbiter at high Reynolds numbers. [Langley low turbulence pressure tunnel

    NASA Technical Reports Server (NTRS)

    Phillips, W. P.

    1981-01-01

    Subsonic longitudinal andd laternal directional characteristics were obtained for several modified configurations of the 140 A/B orbiter (0.010 scale). These modifications, designed to extend longitudinal trim capability forward of the 65 percent fuselage length station, consisted of modified wing planform fillet and a canard. Tests were performed in the Langley Low Turbulence Pressure Tunnel at Reynolds numbers from about 4.2 million to 14.3 million based on the fuselage reference length.

  4. Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Double Delta Wing Model at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Gonzalez, Hugo A.

    2006-01-01

    A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to study the effect of wing fillets on the global vortex induced surface static pressure field about a sharp leading-edge 76 deg./40 deg. double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M(sub infinity) = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an insitu method featuring the simultaneous acquisition of electronically scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M(sub infinity) = 0.50 to 0.85 but increased to several percent at M(sub infinity) =0.95 and 1.20. The PSP pressure distributions and pseudo-colored, planform-view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having parabolic or diamond planforms situated at the strake-wing intersection were respectively designed to manipulate the vortical flows by removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

  5. Three-dimensional aerodynamic analysis of a subsonic transport high-lift configuration and comparisons with wind-tunnel test results

    NASA Technical Reports Server (NTRS)

    Edge, D. Christian; Perkins, John N.

    1995-01-01

    The sizing and efficiency of an aircraft is largely determined by the performance of its high-lift system. Subsonic civil transports most often use deployable multi-element airfoils to achieve the maximum-lift requirements for landing, as well as the high lift-to-drag ratios for take-off. However, these systems produce very complex flow fields which are not fully understood by the scientific community. In order to compete in today's market place, aircraft manufacturers will have to design better high-lift systems. Therefore, a more thorough understanding of the flows associated with these systems is desired. Flight and wind-tunnel experiments have been conducted on NASA Langley's B737-100 research aircraft to obtain detailed full-scale flow measurements on a multi-element high-lift system at various flight conditions. As part of this effort, computational aerodynamic tools are being used to provide preliminary flow-field information for instrumentation development, and to provide additional insight during the data analysis and interpretation process. The purpose of this paper is to demonstrate the ability and usefulness of a three-dimensional low-order potential flow solver, PMARC, by comparing computational results with data obtained from 1/8 scale wind-tunnel tests. Overall, correlation of experimental and computational data reveals that the panel method is able to predict reasonably well the pressures of the aircraft's multi-element wing at several spanwise stations. PMARC's versatility and usefulness is also demonstrated by accurately predicting inviscid three-dimensional flow features for several intricate geometrical regions.

  6. A high subsonic speed wind tunnel investigation of winglets on a representative second-generation jet transport wing

    NASA Technical Reports Server (NTRS)

    Flechner, S. G.; Jacobs, P. F.; Whitcomb, R. T.

    1976-01-01

    The effects of winglets on the aerodynamic forces and moments, loads, and crossflow velocities behind the wing tip are discussed. The results of the investigation indicate that winglets significantly reduce the drag coefficient at lifting conditions. The experiments were conducted in an 8-foot transonic pressure tunnel at Mach numbers from 0.70 to 0.83 and over a lift coefficient range up to 0.65. A semispan model was used.

  7. Low subsonic aerodynamic characteristics of five irregular planform wings with systematically varying wing fillet geometry tested in the NASA/Ames 12 foot pressure tunnel (LA65)

    NASA Technical Reports Server (NTRS)

    Ball, J. W.; Watson, D. B.

    1976-01-01

    An experimental and analytical aerodynamic program to develop predesign guides for irregular planform wings (also referred to as cranked leading edge or double delta wings is reported; the benefits are linearization of subsonic lift curve slope to high angles of attack and avoidance of subsonic pitch instabilities at high lift by proper tailoring of the planform-fillet-wing combination while providing the desired hypersonic trim angle and stability. Because subsonic and hypersonic conditions were the two prime areas of concern in the initial application of this program to optimize shuttle orbiter landing and entry characteristics, the study was designated the Subsonic/Hypersonic Irregular Planforms Study (SHIPS).

  8. Fundamental Aeronautics Program. Subsonic Rotary Wing Project: SRW Aeromechanics Overview/UH-60 Airloads Wind Tunnel Test Summary

    NASA Technical Reports Server (NTRS)

    Norman, Thomas R.

    2011-01-01

    Objectives: a) Advance the understanding of phenomena in aerodynamics, dynamics, and active control of rotorcraft. b) Develop and validate first-principles tools. c) Acquire data for tool validation from small and large-scale testing of existing and novel rotorcraft configurations. Recent Accomplishments include: (CFD) - Made significant improvements in structured and unstructured rotorcraft CFD methods (OVERFLOW and FUN3D). (Icing) - a) Continued development of high-fidelity icing analysis tools. b) Completed test of oscillating airfoil in Icing Research Tunnel (IRT). c) Developed plans and began detailed preparations for subscale rotor test in IRT.

  9. A Wind-tunnel Test Technique for Measuring the Dynamic Rotary Stability Derivatives at Subsonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Beam, Benjamin H

    1956-01-01

    A method is described for measuring the dynamic stability derivatives of a model airplane in a wind tunnel. The characteristic features of this system are that single-degree-of-freedom oscillations were used to obtain combinations of rolling, yawing and pitching motions; that the oscillations were excited and controlled by velocity feedback which permitted operation under conditions unfavorable for more conventional types of oscillatory testing; and that data processing was greatly simplified by using analog computer elements in the strain-gage circuitry. A small number of experimental data are included to illustrate the general scope of results obtainable with this system.

  10. Wind-Tunnel Investigation of the Effect of Porous Spoilers on the Wake of a Subsonic Transport Model

    NASA Technical Reports Server (NTRS)

    Corsiglia, V. R.; Rossow, V. J.

    1976-01-01

    Tests were conducted in the Ames Research Center 40- by 80-Foot Wind Tunnel to determine how porosity of wing spoilers on a B-747 airplane would affect the rolling moments imposed on an aircraft following in the wake. It was found that spoilers with 40 percent porosity and hole diameter to thickness ratio of 1.1 were just as effective in reducing the rolling moment imposed on the follower as solid spoilers, for the case of two spoilers per wing panel (6.4 percent semispan each) with a following model whose span was 20 percent of the span of the generator. When a larger following model was tested, whose span was 50 percent of that of the generator, the effectiveness of the two spoilers per wing was substantially reduced.

  11. Wind-Tunnel Investigation of the Static Lateral Stability Characteristics of Wing-Fuselage Combinations at High Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Wiggins, James W.; Fournier, Paul G.

    1953-01-01

    A systematic research program is being carried out in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic characteristics of various arrangements of the component parts of research-type airplane models, including some complete model configurations. Data are being obtained on characteristics in pitch, sideslip, and during steady roll at Mach numbers from 0.40 to about 0.95. This paper presents results which show the effect of taper ratio on the aerodynamic characteristics in sideslip of wing-fuselage combinations having wings with a sweep of 45 degrees at the quarter-chord line, an aspect ratio of 4, and a NACA 65A006 airfoil section.

  12. Subsonic wind-tunnel tests of a trailing-cone device for calibrating aircraft static pressure systems

    NASA Technical Reports Server (NTRS)

    Jordan, F. L., Jr.; Ritchie, V. S.

    1973-01-01

    A trailing-cone device for calibrating aircraft static-pressure systems was tested in a transonic wind tunnel to investigate the pressure-sensing characteristics of the device including effects of several configuration changes. The tests were conducted at Mach numbers from 0.30 to 0.95 with Reynolds numbers from (0.9 x one million to 4.1 x one million per foot). The results of these tests indicated that the pressures sensed by the device changed slightly but consistently as the distance between the device pressure orifices and cone was varied from 4 to 10 cone diameters. Differences between such device-indicated pressures and free-stream static pressure were small, however, and corresponded to Mach number differences of less than 0.001 for device configurations with pressure orifices located 5 or 6 cone diameters ahead of the cone. Differences between device-indicated and free-stream static pressures were not greatly influenced by a protection skid at the downstream end of the pressure tube of the device nor by a 2-to-1 change in test Reynolds number.

  13. Wind tunnel investigation of an all flush orifice air data system for a large subsonic aircraft. [conducted in a Langley 8 foot transonic pressure tunnel

    NASA Technical Reports Server (NTRS)

    Larson, T. J.; Flechner, S. G.; Siemers, P. M., III

    1980-01-01

    The results of a wind tunnel investigation on an all flush orifice air data system for use on a KC-135A aircraft are presented. The investigation was performed to determine the applicability of fixed all flush orifice air data systems that use only aircraft surfaces for orifices on the nose of the model (in a configuration similar to that of the shuttle entry air data system) provided the measurements required for the determination of stagnation pressure, angle of attack, and angle of sideslip. For the measurement of static pressure, additional flush orifices in positions on the sides of the fuselage corresponding to those in a standard pitot-static system were required. An acceptable but less accurate system, consisting of orifices only on the nose of the model, is defined and discussed.

  14. Pressure Distributions from Subsonic Tests of a NACA 0012 Semispan Wing Model

    NASA Technical Reports Server (NTRS)

    Applin, Zachary T.

    1995-01-01

    An unswept, semispan wing model incorporating a NACA 0012 airfoil section was tested in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data which document effects of wing configuration and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a full-span, single-slotted trailing-edge flap. The trailing-edge flap was tested at a deflection angle of 40 degrees and the Krueger flap at a deflection of 55 degrees. Three wing configurations were tested: cruise, trailing-edge flap only, and Knueger flap and trailing-edge flap deployed. Tests were conducted at free-stream dynamic pressures of 15, 30 and 60 psf, with corresponding chord Reynolds numbers of 1.22 to 2.11 million and Mach numbers of 0.12 to 0.20. Angles of attack presented range from 0 to 20 degrees, depending on wing configuration. The data are presented without analysis.

  15. A Wall Correction Program Based on Classical Methods for the National Transonic Facility (Solid Wall or Slotted Wall) and the 14x22-Ft Subsonic Tunnel at NASA LaRC

    NASA Technical Reports Server (NTRS)

    Hemsch, Michael J. (Technical Monitor); Iyer, Venkit

    2004-01-01

    A Fortran subroutine CMWALL is described, which is an implementation of the collective information from classical methods-based wall corrections. These methods use established closed-form expressions which were developed based on simple linear potential-based methods. This is a simple and rapid tool to calculate corrections due to wall interference in the National Transonic Facility (Solid Wall or Slotted Wall) or the 14x22-Ft Subsonic Tunnel at NASA LaRC. It is designed to be easily implemented in the existing tunnel data reduction programs, either as real-time or post-point. It is however important to realize that the method is based on the simplifying assumptions of linearity, small model and attached flow. The computed results are thus to be viewed as first-cut estimates, to be refined further using more complex methods based on measured wall pressures (known as wall signature methods).

  16. Wind tunnel test results of a 1/8-scale fan-in-wing model

    NASA Technical Reports Server (NTRS)

    Wilson, John C.; Gentry, Garl L.; Gorton, Susan A.

    1996-01-01

    A 1/8-scale model of a fan-in-wing concept considered for development by Grumman Aerospace Corporation for the U.S. Army was tested in the Langley 14- by 22-Foot Subsonic Tunnel. Hover testing, which included height above a pressure-instrumented ground plane, angle of pitch, and angle of roll for a range of fan thrust, was conducted in a model preparation area near the tunnel. The air loads and surface pressures on the model were measured for several configurations in the model preparation area and in the tunnel. The major hover configuration change was varying the angles of the vanes attached to the exit of the fans for producing propulsive force. As the model height above the ground was decreased, there was a significant variation of thrust-removed normal force with constant fan speed. The greatest variation was generally for the height-to-fan exit diameter ratio of less than 2.5; the variation was reduced by deflecting fan exit flow outboard with the vanes. In the tunnel angles of pitch and sideslip, height above the tunnel floor, and wind speed were varied for a range of fan thrust and different vane angle configurations. Other configuration features such as flap deflections and tail incidence were evaluated as well. Though the V-tail empennage provided an increase in static longitudinal stability, the total model configuration remained unstable.

  17. Subsonic Investigation of Leading-Edge Flaps Designed for Vortex- and Attached-Flow on a High-Speed Civil Transport Configuration

    NASA Technical Reports Server (NTRS)

    Campbell, Bryan A.; Kemmerly, Guy T.; Kjerstad, Kevin J.; Lessard, Victor R.

    1999-01-01

    A wind tunnel investigation of two separate leading-edge flaps, designed for vortex and attached-flow, respectively, were conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.12 to 0.27, with corresponding chord Reynolds numbers of 2.50 x 10 (sup 6) to 5.50 x 10 (sup 6). Variations of the leading-edge flap deflection angle were tested with outboard leading-edge flaps deflected 0 deg. and 26.4 deg. Trailing-edge flaps were deflected 0 deg., 10 deg., 12.9 deg., and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein. The data associated with each deflected leading-edge flap indicate L/D improvements over the undeflected configuration. These improvements may be instrumental in providing the necessary lift augmentation required by an actual HSCT during the climb-out and landing phases of the flight envelope. However, further tests will have to be done to assess their full potential.

  18. Low-Speed Wind-Tunnel Test of an Unpowered High-Speed Stoppable Rotor Concept in Fixed-Wing Mode

    NASA Technical Reports Server (NTRS)

    Lance, Michael B.; Sung, Daniel Y.; Stroub, Robert H.

    1991-01-01

    An experimental investigation of the M85, a High Speed Rotor Concept, was conducted at the NASA Langley 14 x 22 foot Subsonic Tunnel, assisted by NASA-Ames. An unpowered 1/5 scale model of the XH-59A helicopter fuselage with a large circular hub fairing, two rotor blades, and a shaft fairing was used as a baseline configuration. The M85 is a rotor wing hybrid aircraft design, and the model was tested with the rotor blade in the fixed wing mode. Assessments were made of the aerodynamic characteristics of various model rotor configurations. Variation in configurations were produced by changing the rotor blade sweep angle and the blade chord length. The most favorable M85 configuration tested included wide chord blades at 0 deg sweep, and it attained a system lift to drag ratio of 8.4.

  19. High-speed Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Ackeret, J

    1936-01-01

    Wind tunnel construction and design is discussed especially in relation to subsonic and supersonic speeds. Reynolds Numbers and the theory of compressible flows are also taken into consideration in designing new tunnels.

  20. Pitot pressure measurements in flow fields behind circular-arc nozzles with exhaust jets at subsonic free-stream Mach numbers. [langley 16 foot transonic tunnel

    NASA Technical Reports Server (NTRS)

    Mason, M. L.; Putnam, L. E.

    1979-01-01

    The flow field behind a circular arc nozzle with exhaust jet was studied at subsonic free stream Mach numbers. A conical probe was used to measure the pitot pressure in the jet and free stream regions. Pressure data were recorded for two nozzle configurations at nozzle pressure ratios of 2.0, 2.9, and 5.0. At each set of test conditions, the probe was traversed from the jet center line into the free stream region at seven data acquisition stations. The survey began at the nozzle exit and extended downstream at intervals. The pitot pressure data may be applied to the evaluation of computational flow field models, as illustrated by a comparison of the flow field data with results of inviscid jet plume theory.

  1. Wind Tunnel Aerodynamic Characteristics of a Transport-type Airfoil in a Simulated Heavy Rain Environment

    NASA Technical Reports Server (NTRS)

    Bezos, Gaudy M.; Dunham, R. Earl, Jr.; Gentry, Garl L., Jr.; Melson, W. Edward, Jr.

    1992-01-01

    The effects of simulated heavy rain on the aerodynamic characteristics of an NACA 64-210 airfoil section equipped with leading-and trailing-edge high-lift devices were investigated in the Langley 14- by 22-Foot Subsonic Tunnel. The model had a chord of 2.5 ft, a span of 8 ft, and was mounted on the tunnel centerline between two large endplates. Aerodynamic measurements in and out of the simulated rain environment were obtained for dynamic pressures of 30 and 50 psf and an angle-of-attack range of 0 to 20 degrees for the cruise configuration. The rain intensity was varied to produce liquid water contents ranging from 16 to 46 gm/cu m. The results obtained for various rain intensity levels and tunnel speeds showed significant losses in maximum lift capability and increases in drag for a given lift as the liquid water content was increased. The results obtained on the landing configuration also indicate a progressive decrease in the angle of attack at which maximum lift occurred and an increase in the slope of the pitching-moment curve as the liquid water content was increased. The sensitivity of test results to the effects of the water surface tension was also investigated. A chemical was introduced into the rain environment that reduced the surface tension of water by a factor of 2. The reduction in the surface tension of water did not significantly alter the level of performance losses for the landing configuration.

  2. Case Studies for the Statistical Design of Experiments Applied to Powered Rotor Wind Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Overmeyer, Austin D.; Tanner, Philip E.; Martin, Preston B.; Commo, Sean A.

    2015-01-01

    The application of statistical Design of Experiments (DOE) to helicopter wind tunnel testing was explored during two powered rotor wind tunnel entries during the summers of 2012 and 2013. These tests were performed jointly by the U.S. Army Aviation Development Directorate Joint Research Program Office and NASA Rotary Wing Project Office, currently the Revolutionary Vertical Lift Project, at NASA Langley Research Center located in Hampton, Virginia. Both entries were conducted in the 14- by 22-Foot Subsonic Tunnel with a small portion of the overall tests devoted to developing case studies of the DOE approach as it applies to powered rotor testing. A 16-47 times reduction in the number of data points required was estimated by comparing the DOE approach to conventional testing methods. The average error for the DOE surface response model for the OH-58F test was 0.95 percent and 4.06 percent for drag and download, respectively. The DOE surface response model of the Active Flow Control test captured the drag within 4.1 percent of measured data. The operational differences between the two testing approaches are identified, but did not prevent the safe operation of the powered rotor model throughout the DOE test matrices.

  3. 27. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    27. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH LABORATORY, BUILDING 25C, WHICH REPLACED THE 10-FOOT WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

  4. 28. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    28. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH LABORATORY, BUILDING 25C, WHICH REPLACED THE 10-FOOT WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

  5. 26. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    26. VIEW OF EXHAUST AND DEFLECTOR FOR SUBSONIC AERODYNAMICS RESEARCH LABORATORY, BUILDING 25C, WHICH REPLACED THE 10-FOOT WIND TUNNEL (1991). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

  6. A Subsonic Wind-Tunnel Study to Determine the Buffet and Static Aerodynamic Characteristics of a Systematic Series of Wings. Phase 1

    NASA Technical Reports Server (NTRS)

    Ray, Edward J.; Taylor, Robert T.

    1968-01-01

    A wind-tunnel investigation has been conducted in the Langley High-Speed 7- by 10-Foot Tunnel to determine the buffet and static aerodynamic characteristics of a systematic wing series at Mach numbers ranging from 0.23 to 0.94. The results have indicated that for a given Mach number, the wings which display superior aerodynamic efficiency characteristics generally display the highest buffet free lift coefficient. The characteristics exhibited by the wings which were considered have indicated that correlations can be made between the onset of buffet and selected divergences in the static aerodynamic characteristics. Axial force has been found to be the most sensitive static component to the onset of buffeting.

  7. Subsonic and transonic hinge moment and wing bending/torsion characteristics of .015 scale space shuttle models 49-0 and 67-TS in the Rockwell International trisonic wind tunnel (IA70), volume 1

    NASA Technical Reports Server (NTRS)

    Hughes, M. T.; Mennell, R. C.

    1974-01-01

    Experimental aerodynamic investigations were conducted on an 0.015-scale representation of the integrated space shuttle launch vehicle in the trisonic wind tunnel. The primary test objective was to obtain subsonic and transonic elevon and bodyflap hinge moments and wing bending-torsion moments in the presence of the launch vehicle. Wing pressures were also recorded for the upper and lower right wing surfaces at two spanwise stations. The hinge moment, wing bending/torsion moments and wing pressure data were recorded over an angle-of-attack (alpha) range from -8 deg to +8 deg, and angle-of-sideslip (beta) range from -8 deg to +8 deg and at Mach numbers of 0.90, 1.12, 1.24 and 1.50. Tests were also conducted to determine the effects of the orbiter rear attach cross beam and the forward attach wedge and strut diameter. The orbiter alone was tested at 0.90 and 1.24 Mach number only.

  8. Aero-acoustic experimental verification of optimum configuration of variable-pitch fans for 40 x 80 foot subsonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Lown, H.

    1977-01-01

    The aerodynamic and acoustic performance of two drive fan configurations (low-speed and high-speed variable pitch design) for a 40 x 80 foot wind tunnel were monitored. A 1/7-scale model was utilized. The necessary aero-acoustic data reduction computer program logic was supplied. Test results were evaluated, and the optimum configuration to be employed in the 40 foot full scale fan was recommended.

  9. Flight and wind-tunnel measurements showing base drag reduction provided by a trailing disk for high Reynolds number turbulent flow for subsonic and transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Powers, Sheryll Goecke; Huffman, Jarrett K.; Fox, Charles H., Jr.

    1986-01-01

    The effectiveness of a trailing disk, or trapped vortex concept, in reducing the base drag of a large body of revolution was studied from measurements made both in flight and in a wind tunnel. Pressure data obtained for the flight experiment, and both pressure and force balance data were obtained for the wind tunnel experiment. The flight test also included data obtained from a hemispherical base. The experiment demonstrated the significant base drag reduction capability of the trailing disk to Mach 0.93 and to Reynolds numbers up to 80 times greater than for earlier studies. For the trailing disk data from the flight experiment, the maximum decrease in base drag ranged form 0.08 to 0.07 as Mach number increased from 0.70 to 0.93. Aircraft angles of attack ranged from 3.9 to 6.6 deg for the flight data. For the trailing disk data from the wind tunnel experiment, the maximum decrease in base and total drag ranged from 0.08 to 0.05 for the approximately 0 deg angle of attack data as Mach number increased from 0.30 to 0.82.

  10. Wind-Tunnel Investigation of Subsonic Longitudinal Aerodynamic Characteristics of a Tiltable-Wing Vertical-Take-Off-and-Landing Supersonic Bomber Configuration Including Turbojet Power Effects

    NASA Technical Reports Server (NTRS)

    Thompson, Robert F.; Vogler, Raymond D.; Moseley, William C., Jr.

    1959-01-01

    Jet-powered model tests were made to determine the low-speed longitudinal aerodynamic characteristics of a vertical-take-off and-landing supersonic bomber configuration. The configuration has an unique engine-wing arrangement wherein six large turbojet engines (three on each side of the fuselage) are buried in a low-aspect-ratio wing which is tilted into the vertical plane for take-off. An essentially two-dimensional variable inlet, spanning the leading edge of each wing semispan, provides air for the engines. Jet flow conditions were simulated for a range of military (nonafterburner) and afterburner turbojet-powered flight at subsonic speeds. Three horizontal tails were tested at a station down-stream of the jet exit and at three heights above the jet axes. A semi-span model was used and test parameters covered wing-fuselage incidence angles from 0 deg to 15 deg, wing angles of attack from -4 deg to 36 deg, a variable range of horizontal-tail incidence angles, and some variations in power simulation conditions. Results show that, with all horizontal tails tested, there were large variations in static stability throughout the lift range. When the wing and fuselage were alined, the model was statically stable throughout the test range only with the largest tail tested (tail span of 1.25 wing span) and only when the tail was located in the low test position which placed the tail nearest to the undeflected jet. For transition flight conditions, none of the tail configurations provided satisfactory longitudinal stability or trim throughout the lift range. Jet flow was destabilizing for most of the test conditions, and varying the jet-exit flow conditions at a constant thrust coefficient had little effect on the stability of this model. Wing leading-edge simulation had some important effects on the longitudinal aerodynamic characteristics.

  11. Investigation of space shuttle orbiter subsonic stability and control characteristics and determination of control surface hinge moments in the Rockwell International low speed wind tunnel (OA37)

    NASA Technical Reports Server (NTRS)

    Hughes, T.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a string-mounted 0.030 scale representation of the 140A/B space shuttle orbiter in the 7.75- by 11-foot low speed wind tunnel. The primary test objectives were to establish basic longitudinal and lateral directional stability and control characteristics for the basic configuration plus control surface hinge moments. Aerodynamic force and moment data were measured in the body axis system by an internally mounted, six-component strain gage balance. Additional configurations investigated were sealed rudder hingeline gaps, sealed elevon gaps and compartmentized speedbrakes.

  12. Flow quality measurements in compressible subsonic flows

    NASA Technical Reports Server (NTRS)

    Stainback, P. Calvin; Johnson, Charles B.

    1987-01-01

    The purpose is to re-examine the heat transfer from a hot-wire probe in the compressible subsonic flow regime; describe the three-wire hot-wire probe calibration and data reduction techniques used to measure the velocity, density, and total temperature fluctuation; and present flow quality results obtained in the Langley 0.3 meter Transonic Cryogenic Wind Tunnel and in flight with the NASA JetStar from the same three-wire hot-wire probe.

  13. X-33 Metal Model Testing In Low Turbulence Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The countrys next generation of space transportation, a reusable launch vehicle (RLV), continues to undergo wind tunnel testing at NASA Langley Research Center, Hampton, Va. All four photos are a metal model of the X-33 reusable launch vehicle (about 15 inches long by 15 inches wide) being tested for Lockheed Martin Skunk Works in the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center. Tests are being conducted by members of the Aerothermodynamics Branch. According to Kelly Murphy of Langleys Aerothermodynamics Branch, the aluminum and stainless steel model of the X-33 underwent aerodynamic testing in the tunnel. *The subsonic tests were conducted at the speed of Mach 25,* she said. *Force and moment testing and measurement in this tunnel lasted about one week.* Future testing of the metal model is scheduled for Langleys 16-Foot Transonic Tunnel, from the end of March to mid-April 1997, and the Unitary Wind Tunnel, from mid-April to the beginning of May. Other tunnel testing for X-33 models are scheduled from the present through June in the hypersonic tunnels, and the 14- by 22-Foot Tunnel from about mid-June to mid-July. Since 1991 Marshall Space Flight Center in Huntsville, Ala. has been the lead center for coordinating the Agencys X-33 Reusable Launch Vehicle (RLV) Program, an industry-led effort, which NASA Administrator Daniel S. Goldin has declared the agency's highest priority new program. The RLV Technology Program is a partnership among NASA, the United States Air Force and private industry to develop world leadership in low-cost space transportation. The goal of the program is to develop technologies and new operational concepts that can radically reduce the cost of access to space. The RLV program also hopes to speed the commercialization of space and improve U.S. economic competitiveness by making access to space as routine and reliable as today's airline industry, while reducing costs and enhancing safety and reliability. The RLV

  14. X-33 Metal Model Testing In Low Turbulence Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    1997-01-01

    The countrys next generation of space transportation, a reusable launch vehicle (RLV), continues to undergo wind tunnel testing at NASA Langley Research Center, Hampton, Va. All four photos are a metal model of the X-33 reusable launch vehicle (about 15 inches long by 15 inches wide) being tested for Lockheed Martin Skunk Works in the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center. Tests are being conducted by members of the Aerothermodynamics Branch. According to Kelly Murphy of Langleys Aerothermodynamics Branch, the aluminum and stainless steel model of the X-33 underwent aerodynamic testing in the tunnel. *The subsonic tests were conducted at the speed of Mach .25,* she said. *Force and moment testing and measurement in this tunnel lasted about one week.* Future testing of the metal model is scheduled for Langleys 16-Foot Transonic Tunnel, from the end of March to mid-April 1997, and the Unitary Wind Tunnel, from mid-April to the beginning of May. Other tunnel testing for X-33 models are scheduled from the present through June in the hypersonic tunnels, and the 14- by 22-Foot Tunnel from about mid-June to mid-July. Since 1991 Marshall Space Flight Center in Huntsville, Ala. has been the lead center for coordinating the Agencys X-33 Reusable Launch Vehicle (RLV) Program, an industry-led effort, which NASA Administrator Daniel S. Goldin has declared the agency's highest priority new program. The RLV Technology Program is a partnership among NASA, the United States Air Force and private industry to develop world leadership in low-cost space transportation. The goal of the program is to develop technologies and new operational concepts that can radically reduce the cost of access to space. The RLV program also hopes to speed the commercialization of space and improve U.S. economic competitiveness by making access to space as routine and reliable as today's airline industry, while reducing costs and enhancing safety and reliability. The RLV

  15. High Lift Common Research Model for Wind Tunnel Testing: An Active Flow Control Perspective

    NASA Technical Reports Server (NTRS)

    Lin, John C.; Melton, Latunia P.; Viken, Sally A.; Andino, Marlyn Y.; Koklu, Mehti; Hannon, Judith A.; Vatsa, Veer N.

    2017-01-01

    This paper provides an overview of a research and development effort sponsored by the NASA Advanced Air Transport Technology Project to achieve the required high-lift performance using active flow control (AFC) on simple hinged flaps while reducing the cruise drag associated with the external mechanisms on slotted flaps of a generic modern transport aircraft. The removal of the external fairings for the Fowler flap mechanism could help to reduce drag by 3.3 counts. The main challenge is to develop an AFC system that can provide the necessary lift recovery on a simple hinged flap high-lift system while using the limited pneumatic power available on the aircraft. Innovative low-power AFC concepts will be investigated in the flap shoulder region. The AFC concepts being explored include steady blowing and unsteady blowing operating in the spatial and/or temporal domain. Both conventional and AFC-enabled high-lift configurations were designed for the current effort. The high-lift configurations share the cruise geometry that is based on the NASA Common Research Model, and therefore, are also open geometries. A 10%-scale High Lift Common Research Model (HL-CRM) is being designed for testing at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel during fiscal year 2018. The overall project plan, status, HL-CRM configurations, and AFC objectives for the wind tunnel test are described.

  16. Wind Tunnel Testing of Powered Lift, All-Wing STOL Model

    NASA Technical Reports Server (NTRS)

    Collins, Scott W.; Westra, Bryan W.; Lin, John C.; Jones, Gregory S.; Zeune, Cal H.

    2008-01-01

    Short take-off and landing (STOL) systems can offer significant capabilities to warfighters and, for civil operators thriving on maximizing efficiencies they can improve airspace use while containing noise within airport environments. In order to provide data for next generation systems, a wind tunnel test of an all-wing cruise efficient, short take-off and landing (CE STOL) configuration was conducted in the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) 14- by 22-foot Subsonic Wind Tunnel. The test s purpose was to mature the aerodynamic aspects of an integrated powered lift system within an advanced mobility configuration capable of CE STOL. The full-span model made use of steady flap blowing and a lifting centerbody to achieve high lift coefficients. The test occurred during April through June of 2007 and included objectives for advancing the state-of-the-art of powered lift testing through gathering force and moment data, on-body pressure data, and off-body flow field measurements during automatically controlled blowing conditions. Data were obtained for variations in model configuration, angles of attack and sideslip, blowing coefficient, and height above ground. The database produced by this effort is being used to advance design techniques and computational tools for developing systems with integrated powered lift technologies.

  17. Wind Tunnel Test of a Risk-Reduction Wing/Fuselage Model to Examine Juncture-Flow Phenomena

    NASA Technical Reports Server (NTRS)

    Kegerise, Michael A.; Neuhart, Dan H.

    2016-01-01

    A wing/fuselage wind-tunnel model was tested in the Langley 14- by 22-foot Subsonic Wind Tunnel in preparation for a highly-instrumented Juncture Flow Experiment to be conducted in the same facility. This test, which was sponsored by the NASA Transformational Tool and Technologies Project, is part of a comprehensive set of experimental and computational research activities to develop revolutionary, physics-based aeronautics analysis and design capability. The objectives of this particular test were to examine the surface and off-body flow on a generic wing/body combination to: 1) choose a final wing for a future, highly instrumented model, 2) use the results to facilitate unsteady pressure sensor placement on the model, 3) determine the area to be surveyed with an embedded laser-doppler velocimetry (LDV) system, 4) investigate the primary juncture corner- flow separation region using particle image velocimetry (PIV) to see if the particle seeding is adequately entrained and to examine the structure in the separated region, and 5) to determine the similarity of observed flow features with those predicted by computational fluid dynamics (CFD). This report documents the results of the above experiment that specifically address the first three goals. Multiple wing configurations were tested at a chord Reynolds number of 2.4 million. Flow patterns on the surface of the wings and in the region of the wing/fuselage juncture were examined using oil- flow visualization and infrared thermography. A limited number of unsteady pressure sensors on the fuselage around the wing leading and trailing edges were used to identify any dynamic effects of the horseshoe vortex on the flow field. The area of separated flow in the wing/fuselage juncture near the wing trailing edge was observed for all wing configurations at various angles of attack. All of the test objectives were met. The staff of the 14- by 22-foot Subsonic Wind Tunnel provided outstanding support and delivered

  18. Investigation of space shuttle orbiter subsonic stability and control characteristics in the NAAL low speed wind tunnel (0A62b), volume 1

    NASA Technical Reports Server (NTRS)

    Mennell, R.; Hughes, T.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a sting-mounted 0.0405 scale representation of the 140A/B space shuttle orbiter in a 7.75 ft by 11 ft low speed wind tunnel during the time period from November 14, 1973, to December 6, 1973, with the primary test objectives being to establish basic longitudinal stability characteristics in and out of ground effect, as well as lateral-directional stability characteristics in free air. Two dual podded nacelle configurations were also tested, one with three dual podded nacelles on the lower wing surface, and the other with a single dual nacelle on the lower centerline with dual nacelle pylons mounted above each wing. Stability and control characteristics were investigated at nominal elevon, rudder, aileron, and body flap deflections. Pressure bugs were used to determine pressures on the vertical tail at spanwise stations, and aerodynamic force and moment data were measured in the stability axis system by an internally mounted, six component strain gage balance.

  19. Investigation of space shuttle orbiter subsonic stability and control characteristics in the NAAL low speed wind tunnel (OA62B), volume 2

    NASA Technical Reports Server (NTRS)

    Mennell, R.; Hughes, T.

    1974-01-01

    Experimental aerodynamic investigations were conducted on a sting-mounted 0.0405 scale representation of the 140A/B space shuttle orbiter in a 7.75 ft by 11 ft low speed wind tunnel during the period from November 14, 1973 to December 6, 1973. Establishment of basic longitudinal stability characteristics in and out of ground effect, and the establishment of lateral-directional stability characteristics in free air were the primary test objectives. The following effects and configurations were tested: (1) two dual podded nacelle configurations; (2) stability and control characteristics at nominal elevon deflections, rudder deflections, airleron deflections, rudder flare angles, and body flap deflections; (3) effects of various elevon and elevon/fuselage gaps on longitudinal stability and control; (4) pressures on the vertical tail at spanwise stations using pressure bugs; (5) aerodynamic force and moment data measured in the stability axis system by an internally mounted, six-component strain gage balance. For Vol. 1, see N74-32324.

  20. Refined methods of aeroelastic analysis and optimization. [swept wings, propeller theory, and subsonic flutter

    NASA Technical Reports Server (NTRS)

    Ashley, H.

    1984-01-01

    Graduate research activity in the following areas is reported: the divergence of laminated composite lifting surfaces, subsonic propeller theory and aeroelastic analysis, and cross sectional resonances in wind tunnels.

  1. 8. VIEW SOUTHWEST, INTERIOR VIEW, WIND TUNNEL 139 Naval ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. VIEW SOUTHWEST, INTERIOR VIEW, WIND TUNNEL 139 - Naval Surface Warfare Center, Subsonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  2. 2. VIEW SOUTH OF WIND TUNNEL 138 AND COOLING SYSTEM ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    2. VIEW SOUTH OF WIND TUNNEL 138 AND COOLING SYSTEM 140, NORTH ELEVATION - Naval Surface Warfare Center, Subsonic Wind Tunnel Building, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

  3. Effect of Winglets on a First-Generation Jet Transport Wing. 2: Pressure and Spanwise Load Distributions for a Semispan Model at High Subsonic Speeds. [in the Langley 8 ft transonic tunnel

    NASA Technical Reports Server (NTRS)

    Montoya, L. C.; Flechner, S. G.; Jacobs, P. F.

    1977-01-01

    Pressure and spanwise load distributions on a first-generation jet transport semispan model at high subsonic speeds are presented for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. Selected data are discussed to show the general trends and effects of the various configurations.

  4. Pressure Distributions About Finite Wedges in Bounded and Unbounded Subsonic Streams

    NASA Technical Reports Server (NTRS)

    Donoughe, Patrick L; Prasse, Ernst I

    1953-01-01

    An analytical investigation of incompressible flow about wedges was made to determine effects of tunnel-wedge ratio and wedge angle on the wedge pressure distributions. The region of applicability of infinite wedge-type velocity distribution was examined for finite wedges. Theoretical and experimental pressure coefficients for various tunnel-wedge ratios, wedge angles, and subsonic Mach numbers were compared.

  5. Numerical Study of the High-Speed Leg of a Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Nayani, Sudheer; Sellers, William L., III; Brynildsen, Scott E.; Everhart, Joel L.

    2015-01-01

    The paper describes the numerical study of the high-speed leg of the NASA Langley 14 by 22-foot Low Speed Wind Tunnel. The high-speed leg consists of the Settling Chamber, Contraction, Test Section, and First Diffuser. Results are shown comparing two different exit boundary conditions and two different methods of determining the surface geometry.

  6. Advanced Subsonic Combustion Rig

    NASA Technical Reports Server (NTRS)

    Lee, Chi-Ming

    1998-01-01

    Researchers from the NASA Lewis Research Center have obtained the first combustion/emissions data under extreme future engine operating conditions. In Lewis' new world-class 60-atm combustor research facility--the Advanced Subsonic Combustion Rig (ASCR)--a flametube was used to conduct combustion experiments in environments as extreme as 900 psia and 3400 F. The greatest challenge for combustion researchers is the uncertainty of the effects of pressure on the formation of nitrogen oxides (NOx). Consequently, U.S. engine manufacturers are using these data to guide their future combustor designs. The flametube's metal housing has an inside diameter of 12 in. and a length of 10.5 in. The flametube can be used with a variety of different flow paths. Each flow path is lined with a high-temperature, castable refractory material (alumina) to minimize heat loss. Upstream of the flametube is the injector section, which has an inside diameter of 13 in. and a length of 0.5-in. It was designed to provide for quick changeovers. This flametube is being used to provide all U.S. engine manufacturers early assessments of advanced combustion concepts at full power conditions prior to engine production. To date, seven concepts from engine manufacturers have been evaluated and improved. This collaborated development can potentially give U.S. engine manufacturers the competitive advantage of being first in the market with advanced low-emission technologies.

  7. Boundary interference assessment and correction for open jet wind tunnels using panel methods

    NASA Astrophysics Data System (ADS)

    Mokhtar, Wael Ahmed

    . The method is also used during a test of full scale NASCAR inside the NASA Langley Research Center 14- by 22- Foot Subsonic Wind Tunnel. This part includes the effects of test section length and the inclusion of the nozzle in the solution on the predicted boundary interference. Finally, a test is conducted at the 1/15th scale Langley Full Scale Tunnel using a generic automotive model ("Davis" model) to validate the prediction of the boundary distortion and to investigate the effect of the collector. The developed method showed reliability when compared to the classical method of images. Through the studied wind tunnel tests, the method showed enough flexibility to be applied to solve both aeronautical and automotive models and several test section configurations with a reasonable computational efficiency.

  8. 12. SOUTHWEST VIEW OF BUILDING 25C (SUBSONIC AERODYNAMICS TEST FACILITY) ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    12. SOUTHWEST VIEW OF BUILDING 25C (SUBSONIC AERODYNAMICS TEST FACILITY) (1992). - Wright-Patterson Air Force Base, Area B, Buildings 25 & 24,10-foot & 20-foot Wind Tunnel Complex, Northeast side of block bounded by K, G, Third, & Fifth Streets, Dayton, Montgomery County, OH

  9. Shuttle model tailcone pressure distribution at low subsonic speeds of a 0.03614-scale model in the NASA/LaRC low-turbulence pressure tunnel (LA81), volume 1

    NASA Technical Reports Server (NTRS)

    Ball, J. W.; Lindahl, R. H.

    1976-01-01

    An investigation was conducted in the NASA/LaRC Low-Turbulence Pressure Tunnel on a 0.03614-scale orbiter model of a 089B configuration with a 139B configuration nose forward of F.S. 500. The tailcone was the TC sub 4 design and was instrumented with eighty-nine pressure orifices. Control surfaces were deflected and three wind tunnel mounting techniques were investigated over an angle-of-attack range from -2 deg to a maximum of 18 deg. In order to determine the sensitivity of the tailcone to changes in Reynolds number, most of the test was made at a Mach number of 0.20 over a Reynolds number range of 2.0 to 10 million per foot. A few runs were made at a Mach number of 0.30 at Reynolds numbers of 4.0, 6.0, and 8 million per foot.

  10. Robust, Optimal Subsonic Airfoil Shapes

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan

    2014-01-01

    A method has been developed to create an airfoil robust enough to operate satisfactorily in different environments. This method determines a robust, optimal, subsonic airfoil shape, beginning with an arbitrary initial airfoil shape, and imposes the necessary constraints on the design. Also, this method is flexible and extendible to a larger class of requirements and changes in constraints imposed.

  11. Effect of winglets on a first-generation jet transport wing. 1: Longitudinal aerodynamic characteristics of a semispan model at subsonic speeds. [in the Langley 8 ft transonic tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, P. F.; Flechner, S. G.; Montoya, L. C.

    1977-01-01

    The effects of winglets and a simple wing-tip extension on the aerodynamic forces and moments and the flow-field cross flow velocity vectors behind the wing tip of a first generation jet transport wing were investigated in the Langley 8-foot transonic pressure tunnel using a semi-span model. The test was conducted at Mach numbers of 0.30, 0.70, 0.75, 0.78, and 0.80. At a Mach number of 0.30, the configurations were tested with combinations of leading- and trailing-edge flaps.

  12. Effect of winglets on a first-generation jet transport wing. 6: Stability characteristics for a full-span model at subsonic speeds. [conducted in Langley 8 foot transonic pressure tunnel

    NASA Technical Reports Server (NTRS)

    Flechner, S. G.

    1979-01-01

    A wind tunnel investigation to identify changes in stability and control characteristics of a model KC-135A due to the addition of winglets is presented. Static longitudinal and lateral-directional aerodynamic characteristics were determined for the model with and without winglets. Variations in the aerodynamic characteristics at various Mach numbers, angles of attack, and angles of slidslip are discussed. The effect of the winglets on the drag and lift coefficients are evaluated and the low speed and high speed characteristics of the model are reported.

  13. Simulation of Atmospheric-Entry Capsules in the Subsonic Regime

    NASA Technical Reports Server (NTRS)

    Murman, Scott M.; Childs, Robert E.; Garcia, Joseph A.

    2015-01-01

    The accuracy of Computational Fluid Dynamics predictions of subsonic capsule aerodynamics is examined by comparison against recent NASA wind-tunnel data at high-Reynolds-number flight conditions. Several aspects of numerical and physical modeling are considered, including inviscid numerical scheme, mesh adaptation, rough-wall modeling, rotation and curvature corrections for eddy-viscosity models, and Detached-Eddy Simulations of the unsteady wake. All of these are considered in isolation against relevant data where possible. The results indicate that an improved predictive capability is developed by considering physics-based approaches and validating the results against flight-relevant experimental data.

  14. Robust, optimal subsonic airfoil shapes

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan (Inventor)

    2008-01-01

    Method system, and product from application of the method, for design of a subsonic airfoil shape, beginning with an arbitrary initial airfoil shape and incorporating one or more constraints on the airfoil geometric parameters and flow characteristics. The resulting design is robust against variations in airfoil dimensions and local airfoil shape introduced in the airfoil manufacturing process. A perturbation procedure provides a class of airfoil shapes, beginning with an initial airfoil shape.

  15. Subsonic Dynamic Stability Tests of a Sample Return Entry Vehicle

    NASA Technical Reports Server (NTRS)

    Fremaux, C. Michael; Johnson, R. Keith

    2006-01-01

    An investigation has been conducted in the NASA Langley 20-Foot Vertical Spin Tunnel (VST) to determine the subsonic dynamic stability characteristics of a proposed atmospheric entry vehicle for sample return missions. In particular, the effects of changes in aft-body geometry on stability were examined. Freeflying tests of a dynamically scaled model with various geometric features were conducted, including cases in which the model was perturbed to measure dynamic response. Both perturbed and non-perturbed runs were recorded as motion time histories using the VST optical data acquisition system and reduced for post-test analysis. In addition, preliminary results from a static force and moment test of a similar model in the Langley 12-Foot Low Speed Tunnel are presented. Results indicate that the configuration is dynamically stable for the baseline geometry, but exhibits degraded dynamic behavior for the geometry modifications tested.

  16. Advanced Subsonic Combustion Rig Developed

    NASA Technical Reports Server (NTRS)

    1995-01-01

    The Advanced Subsonic Combustion Rig (ASCR), a unique, state-of-the-art facility for conducting combustion research, is located at the NASA Lewis Research Center in Cleveland, Ohio. The ASCR, which was nearing completion at the close of 1995, will be capable of simulating the very high pressure and high temperature conditions that are expected to exist in future, advanced subsonic gas turbine (jet) engines. Future environmental regulations will require much cleaner burning (more environmentally friendly) aircraft engines. The ASCR is critical to the development of these cleaner engines. It will allow NASA and U.S. aircraft engine industry researchers to identify and test promising clean-burning gas turbine engine combustion concepts under the pressure and temperature conditions that are expected for those future engines. Combustion processes will be investigated for a variety of next-generation aircraft engine sizes, including engines for large, long-range aircraft (with typical trip lengths of about 3000 mi) and for regional aircraft (with typical trip lengths of about 400 mi). The ASCR design was conceived and initiated in 1993, and fabrication and construction of the rig, including the buildup of an advanced control room, took place throughout 1994 and 1995. In early 1996, the ASCR will be operational for obtaining research data. The ASCR is an intricate part of the NASA Advanced Subsonic Technology Propulsion Program, which is aimed at developing technologies critical to the next generation of gas turbine engines. This effort is in collaboration with the U.S. aircraft gas turbine engine industry. A goal of the Advanced Subsonic Technology Propulsion Program is to develop combustion concepts and technologies that will result in gas turbine engines that produce 50 percent less nitrous oxide (NO_x) pollutants than current engines do. This facility is unique in its capability to simulate advanced subsonic engine pressure, temperature, and air flow rate conditions

  17. Wind-Tunnel Investigation at Subsonic and Supersonic Speeds of the Static and Dynamic Stability Derivatives of an Airplane Model with an Unswept Wing and a High Horizontal Tail

    NASA Technical Reports Server (NTRS)

    Lessing, Henry C.; Butler, James K.

    1959-01-01

    Results are presented of a wind-tunnel investigation to evaluate the static and dynamic stability derivatives of a model with a low-aspect-ratio unswept wing and a high horizontal tail. In addition to results for the complete model, results were also obtained of the body alone, body and wing, and body and tail. Data were obtained in the Mach number range from 0.65 to 2.2, at a Reynolds number of 2 million based on the wing mean aerodynamic chord. The angle-of-attack range for most of the data was -11.5 deg to 18 deg. A limited amount of data was obtained with fixed transition. A correspondence between the damping in pitch and the static stability, previously noted in other investigations, was also observed in the present results. The effect observed was that a decrease (or increase) in the static stability was accompanied by an increase (or decrease) in the damping in pitch. A similar correspondence was observed between the damping in yaw and the static-directional stability. Results from similar tests of the same model configuration in two other facilities over different speed ranges are presented for comparison. It was found that most of the results from the three investigations correlated reasonably well. Estimates of the rotary derivatives were made using available procedures. Comparison with the experimental results indicates the need for development of more precise estimation procedures.

  18. Subsonic Aircraft Safety Icing Study

    NASA Technical Reports Server (NTRS)

    Jones, Sharon Monica; Reveley, Mary S.; Evans, Joni K.; Barrientos, Francesca A.

    2008-01-01

    NASA's Integrated Resilient Aircraft Control (IRAC) Project is one of four projects within the agency s Aviation Safety Program (AvSafe) in the Aeronautics Research Mission Directorate (ARMD). The IRAC Project, which was redesigned in the first half of 2007, conducts research to advance the state of the art in aircraft control design tools and techniques. A "Key Decision Point" was established for fiscal year 2007 with the following expected outcomes: document the most currently available statistical/prognostic data associated with icing for subsonic transport, summarize reports by subject matter experts in icing research on current knowledge of icing effects on control parameters and establish future requirements for icing research for subsonic transports including the appropriate alignment. This study contains: (1) statistical analyses of accident and incident data conducted by NASA researchers for this "Key Decision Point", (2) an examination of icing in other recent statistically based studies, (3) a summary of aviation safety priority lists that have been developed by various subject-matter experts, including the significance of aircraft icing research in these lists and (4) suggested future requirements for NASA icing research. The review of several studies by subject-matter experts was summarized into four high-priority icing research areas. Based on the Integrated Resilient Aircraft Control (IRAC) Project goals and objectives, the IRAC project was encouraged to conduct work in all of the high-priority icing research areas that were identified, with the exception of the developing of methods to sense and document actual icing conditions.

  19. LOW SUBSONIC PRESSURE DISTRIBUTIONS ON THREE RIGID WINGS SIMULATING PARAGLIDERS WITH VARIED CANOPY CURVATURE AND LEADING-EDGE SWEEP

    DTIC Science & Technology

    An investigation was made in the Langley 7- by 10-foot transonic tunnel to determine the subsonic pressure distribution of three paraglider models...through an angle-of-attack range from 0 to 74 degrees. Three rigid meta models simulated a 45 degrees basic flat planform paraglider with leading-edge

  20. An NACA Vane-Type Angle-of-Attack Indicator for use at Subsonic and Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Mitchell, Jesse L.; Peck, Robert F.

    1949-01-01

    A vane-type angle-of-attack indicator suitable for measurements at both subsonic and supersonic speeds has been developed by the National Advisory Committee for Aeronautics. A brief history is given of the development, and a wind-tunnel calibration of the indicator is presented, together with a discussion of the corrections to be applied to the indicated readings.

  1. Aeroelasticity Benchmark Assessment: Subsonic Fixed Wing Program

    NASA Technical Reports Server (NTRS)

    Florance, Jennifer P.; Chwalowski, Pawel; Wieseman, Carol D.

    2010-01-01

    The fundamental technical challenge in computational aeroelasticity is the accurate prediction of unsteady aerodynamic phenomena and the effect on the aeroelastic response of a vehicle. Currently, a benchmarking standard for use in validating the accuracy of computational aeroelasticity codes does not exist. Many aeroelastic data sets have been obtained in wind-tunnel and flight testing throughout the world; however, none have been globally presented or accepted as an ideal data set. There are numerous reasons for this. One reason is that often, such aeroelastic data sets focus on the aeroelastic phenomena alone (flutter, for example) and do not contain associated information such as unsteady pressures and time-correlated structural dynamic deflections. Other available data sets focus solely on the unsteady pressures and do not address the aeroelastic phenomena. Other discrepancies can include omission of relevant data, such as flutter frequency and / or the acquisition of only qualitative deflection data. In addition to these content deficiencies, all of the available data sets present both experimental and computational technical challenges. Experimental issues include facility influences, nonlinearities beyond those being modeled, and data processing. From the computational perspective, technical challenges include modeling geometric complexities, coupling between the flow and the structure, grid issues, and boundary conditions. The Aeroelasticity Benchmark Assessment task seeks to examine the existing potential experimental data sets and ultimately choose the one that is viewed as the most suitable for computational benchmarking. An initial computational evaluation of that configuration will then be performed using the Langley-developed computational fluid dynamics (CFD) software FUN3D1 as part of its code validation process. In addition to the benchmarking activity, this task also includes an examination of future research directions. Researchers within the

  2. Wind-Tunnel Capability at Ames Research Center

    NASA Technical Reports Server (NTRS)

    Snyder, C. T.; Presley, L. L.

    1987-01-01

    Report describes $700 million wind-tunnel complex at Ames Research Center, including auxiliary support systems, test instrumentation, and special test rigs. Planned near-term facility improvement aimed at providing new test capabilities and increased productivity, as well as some potential longer-term improvements, also discussed. Aerodynamic test facilities range from subsonic wind tunnels to highenthalpy arc jets.

  3. Subsonic Static and Dynamic Aerodynamics of Blunt Entry Vehicles

    NASA Technical Reports Server (NTRS)

    Mitcheltree, Robert A.; Fremaux, Charles M.; Yates, Leslie A.

    1999-01-01

    The incompressible subsonic aerodynamics of four entry-vehicle shapes with variable c.g. locations are examined in the Langley 20-Foot Vertical Spin Tunnel. The shapes examined are spherically-blunted cones with half-cone angles of 30, 45, and 60 deg. The nose bluntness varies between 0.25 and 0.5 times the base diameter. The Reynolds number based on model diameter for these tests is near 500,000. Quantitative data on attitude and location are collected using a video-based data acquisition system and reduced with a six deg-of-freedom inverse method. All of the shapes examined suffered from strong dynamic instabilities which could produced limit cycles with sufficient amplitudes to overcome static stability of the configuration. Increasing cone half-angle or nose bluntness increases drag but decreases static and dynamic stability.

  4. Estimation of Rotary Stability Derivatives at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Tobak, Murray; Lessing, Henry C.

    1961-01-01

    The first part of this paper pertains to the estimation of subsonic rotary stability derivatives of wings. The unsteady potential flow problem is solved by a superposition of steady flow solutions. Numerical results for the damping coefficients of triangular wings are presented as functions of aspect ratio and Mach number, and are compared with experimental results over the Mach number range 0 to 1. In the second part, experimental results are used. to point out a close correlation between the nonlinear variations with angle of attack of the static pitching-moment curve slope and the damping-in-pitch coefficient. The underlying basis for the correlation is found as a result of an analysis in which the indicial function concept and. the principle of super-position are adapted to apply to the nonlinear problem. The form of the result suggests a method of estimating nonlinear damping coefficients from results of static wind-tunnel measurements.

  5. A study to determine methods of improving the subsonic performance of a proposed Personnel Launch System (PLS) concept

    NASA Technical Reports Server (NTRS)

    Spencer, Bernard, Jr.; Fox, Charles H.; Huffman, Jarrett K.

    1995-01-01

    An investigation has been conducted in the Langley 7- by 10-Foot High Speed Wind Tunnel to determine the longitudinal and lateral directional aerodynamic characteristics of a series of personnel launch system concepts. This series of configurations evolved during an effort to improve the subsonic characteristics of a proposed lifting entry vehicle (designated the HL-20). The primary purpose of the overall investigation was to provide a vehicle concept which was inherently stable and trimable from entry to landing while examining methods of improving subsonic aerodynamic performance.

  6. The 13-inch magnetic suspension and balance system wind tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, William G., Jr.; Dress, David A.

    1989-01-01

    NASA Langley has a small, subsonic wind tunnel in use with the 13-inch Magnetic Suspension and Balance System (MSBS). The tunnel is capable of speeds up to Mach 0.5. This report presents tunnel design and construction details. It includes flow uniformity, angularity, and velocity fluctuation data. It also compares experimental Mach number distribution data with computed results for the General Electric Streamtube Curvature Program.

  7. A Vision in Aeronautics: The K-12 Wind Tunnel Project

    NASA Technical Reports Server (NTRS)

    1997-01-01

    A Vision in Aeronautics, a project within the NASA Lewis Research Center's Information Infrastructure Technologies and Applications (IITA) K-12 Program, employs small-scale, subsonic wind tunnels to inspire students to explore the world of aeronautics and computers. Recently, two educational K-12 wind tunnels were built in the Cleveland area. During the 1995-1996 school year, preliminary testing occurred in both tunnels.

  8. Subsonic and transonic propeller noise

    NASA Astrophysics Data System (ADS)

    Lewy, S.; Gounet, H.

    Models for the noise levels from propellers are discussed, with results compared to in-flight measurements. Methods originally applied to noise from light aircraft are modified and extended to high speed passenger aircraft. Noise emitted from propellers has three components: a monopolar emission due to the air displaced by a blade; a bipolar form from average and fluctuating forces exerted by the blades; and a quadripolar component produced by deformation of the streamlines around the blade profile and defined by the Lighthill tensor. The latter is not a factor in the subsonic regime and can be neglected. Attention is given to a formalism which accounts for the sound level along each band, the frequency harmonics at each blade passage, the number of blades, and the rotation rate. The measured directivities of the two components are described. It is found that the radiated noise levels can be reduced in slow aircraft by lowering the peripheral velocity while keeping the same power with more blades. Calculations including the quadripolar term are necessary for modeling noise levels in transonic propellers.

  9. Subsonic Ultra Green Aircraft Research

    NASA Technical Reports Server (NTRS)

    Bradley, Marty K.; Droney, Christopher K.

    2011-01-01

    This Final Report summarizes the work accomplished by the Boeing Subsonic Ultra Green Aircraft Research (SUGAR) team in Phase 1, which includes the time period of October 2008 through March 2010. The team consisted of Boeing Research and Technology, Boeing Commercial Airplanes, General Electric, and Georgia Tech. The team completed the development of a comprehensive future scenario for world-wide commercial aviation, selected baseline and advanced configurations for detailed study, generated technology suites for each configuration, conducted detailed performance analysis, calculated noise and emissions, assessed technology risks, and developed technology roadmaps. Five concepts were evaluated in detail: 2008 baseline, N+3 reference, N+3 high span strut braced wing, N+3 gas turbine battery electric concept, and N+3 hybrid wing body. A wide portfolio of technologies was identified to address the NASA N+3 goals. Significant improvements in air traffic management, aerodynamics, materials and structures, aircraft systems, propulsion, and acoustics are needed. Recommendations for Phase 2 concept and technology projects have been identified.

  10. CFD and experimental data of closed-loop wind tunnel flow.

    PubMed

    Calautit, John Kaiser; Hughes, Ben Richard

    2016-06-01

    The data presented in this article were the basis for the study reported in the research articles entitled 'A validated design methodology for a closed loop subsonic wind tunnel' (Calautit et al., 2014) [1], which presented a systematic investigation into the design, simulation and analysis of flow parameters in a wind tunnel using Computational Fluid Dynamics (CFD). The authors evaluated the accuracy of replicating the flow characteristics for which the wind tunnel was designed using numerical simulation. Here, we detail the numerical and experimental set-up for the analysis of the closed-loop subsonic wind tunnel with an empty test section.

  11. PIV Measurements in the 14 x 22 Low Speed Tunnel: Recommendations for Future Testing

    NASA Technical Reports Server (NTRS)

    Watson, Ralph D.; Jenkins, Luther N.; Yao, Chung-Sheng; McGinley, Catherine B.; Paschal, Keith B.; Neuhart, Dan H.

    2003-01-01

    During the period from February 4 to March 21, 2003 stereo digital particle imaging velocimetry measurements were made on a generic high lift model, the Trap Wing, as part of the High Lift Flow Physics Experiment. These measurements were the first PIV measurements made in the NASA, Langley Research Center 14 x 22 Foot Low Speed Tunnel, and several problems were encountered and solved in the acquisition of the data. It is the purpose of this paper to document the solutions to these problems and to make recommendations for further improvements to the tunnel/setup in order to facilitate future measurements of this type.

  12. Integral equations for flows in wind tunnels

    NASA Technical Reports Server (NTRS)

    Fromme, J. A.; Golberg, M. A.

    1979-01-01

    This paper surveys recent work on the use of integral equations for the calculation of wind tunnel interference. Due to the large number of possible physical situations, the discussion is limited to two-dimensional subsonic and transonic flows. In the subsonic case, the governing boundary value problems are shown to reduce to a class of Cauchy singular equations generalizing the classical airfoil equation. The theory and numerical solution are developed in some detail. For transonic flows nonlinear singular equations result, and a brief discussion of the work of Kraft and Kraft and Lo on their numerical solution is given. Some typical numerical results are presented and directions for future research are indicated.

  13. Unsteady Aerodynamics - Subsonic Compressible Inviscid Case

    NASA Technical Reports Server (NTRS)

    Balakrishnan, A. V.

    1999-01-01

    This paper presents a new analytical treatment of Unsteady Aerodynamics - the linear theory covering the subsonic compressible (inviscid) case - drawing on some recent work in Operator Theory and Functional Analysis. The specific new results are: (a) An existence and uniqueness proof for the Laplace transform version of the Possio integral equation as well as a new closed form solution approximation thereof. (b) A new representation for the time-domain solution of the subsonic compressible aerodynamic equations emphasizing in particular the role of the initial conditions.

  14. 14 CFR 91.805 - Final compliance: Subsonic airplanes.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 2 2010-01-01 2010-01-01 false Final compliance: Subsonic airplanes. 91... § 91.805 Final compliance: Subsonic airplanes. Except as provided in §§ 91.809 and 91.811, on and after January 1, 1985, no person may operate to or from an airport in the United States any subsonic...

  15. 14 CFR 91.805 - Final compliance: Subsonic airplanes.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 2 2011-01-01 2011-01-01 false Final compliance: Subsonic airplanes. 91... § 91.805 Final compliance: Subsonic airplanes. Except as provided in §§ 91.809 and 91.811, on and after January 1, 1985, no person may operate to or from an airport in the United States any subsonic...

  16. A study on high subsonic airfoil flows in relatively high Reynolds number by using OpenFOAM

    NASA Astrophysics Data System (ADS)

    Nakao, Shinichiro; Kashitani, Masashi; Miyaguni, Takeshi; Yamaguchi, Yutaka

    2014-04-01

    In the present study, numerical calculations of the flow-field around the airfoil model are performed by using the OpenFOAM in high subsonic flows. The airfoil model is NACA 64A010. The maximum thickness is 10 % of the chord length. The SonicFOAM and the RhoCentralFOAM are selected as the solver in high subsonic flows. The grid point is 158,000 and the Mach numbers are 0.277 and 0.569 respectively. The CFD data are compared with the experimental data performed by the cryogenic wind tunnel in the past. The results are as follows. The numerical results of the pressure coefficient distribution on the model surface calculated by the SonicFOAM solver showed good agreement with the experimental data measured by the cryogenic wind tunnel. And the data calculated by the SonicFOAM have the capability for the quantitative comparison of the experimental data at low angle of attack.

  17. Hot-wire calibration in subsonic/transonic flow regimes

    NASA Technical Reports Server (NTRS)

    Nagabushana, K. A.; Ash, Robert L.

    1995-01-01

    A different approach for calibrating hot-wires, which simplifies the calibration procedure and reduces the tunnel run-time by an order of magnitude was sought. In general, it is accepted that the directly measurable quantities in any flow are velocity, density, and total temperature. Very few facilities have the capability of varying the total temperature over an adequate range. However, if the overheat temperature parameter, a(sub w), is used to calibrate the hot-wire then the directly measurable quantity, voltage, will be a function of the flow variables and the overheat parameter i.e., E = f(u,p,a(sub w), T(sub w)) where a(sub w) will contain the needed total temperature information. In this report, various methods of evaluating sensitivities with different dependent and independent variables to calibrate a 3-Wire hot-wire probe using a constant temperature anemometer (CTA) in subsonic/transonic flow regimes is presented. The advantage of using a(sub w) as the independent variable instead of total temperature, t(sub o), or overheat temperature parameter, tau, is that while running a calibration test it is not necessary to know the recovery factor, the coefficients in a wire resistance to temperature relationship for a given probe. It was deduced that the method employing the relationship E = f (u,p,a(sub w)) should result in the most accurate calibration of hot wire probes. Any other method would require additional measurements. Also this method will allow calibration and determination of accurate temperature fluctuation information even in atmospheric wind tunnels where there is no ability to obtain any temperature sensitivity information at present. This technique greatly simplifies the calibration process for hot-wires, provides the required calibration information needed in obtaining temperature fluctuations, and reduces both the tunnel run-time and the test matrix required to calibrate hotwires. Some of the results using the above techniques are presented

  18. The Langley Annular Transonic Tunnel

    NASA Technical Reports Server (NTRS)

    Habel, Louis W; Henderson, James H; Miller, Mason F

    1952-01-01

    Report describes the development of the Langley annular transonic tunnel, a facility in which test Mach numbers from 0.6 to slightly over 1.0 are achieved by rotating the test model in an annular passage between two concentric cylinders. Data obtained for two-dimensional airfoil models in the Langley annular transonic tunnel at subsonic and sonic speeds are shown to be in reasonable agreement with experimental data from other sources and with theory when comparisons are made for nonlifting conditions or for equal normal-force coefficients rather than for equal angles of attack. The trends of pressure distributions obtained from measurements in the Langley annular transonic tunnel are consistent with distributions calculated for Prandtl-Meyer flow.

  19. Airfoil shape for flight at subsonic speeds

    DOEpatents

    Whitcomb, Richard T.

    1976-01-01

    An airfoil having an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency.

  20. Flight-determined aerodynamic stability and control derivatives of the M2-F2 lifting body vehicle at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Kempel, R. W.; Thompson, R. C.

    1971-01-01

    Aerodynamic derivatives were obtained for the M2-F2 lifting body flight vehicle in the subsonic flight region between Mach numbers of 0.41 and 0.64 and altitudes of 7000 feet to 45,000 feet. The derivatives were determined by a flight time history curve-fitting process utilizing a hybrid computer. The flight-determined derivatives are compared with wind-tunnel and predicted values. Modal-response characteristics, calculated from the flight derivatives, are presented.

  1. Calibration and test capabilities of the Langley 7- by 10- foot high speed tunnel

    NASA Technical Reports Server (NTRS)

    Fox, C. H., Jr.; Huffman, J. K.

    1977-01-01

    The results of a new subsonic calibration of the Langley 7 by 10 foot high speed tunnel with the test section in a solid wall configuration are presented. A description of the test capabilities of the 7 by 10 foot high speed tunnel is also given.

  2. CFD and experimental data of closed-loop wind tunnel flow

    PubMed Central

    Calautit, John Kaiser; Hughes, Ben Richard

    2016-01-01

    The data presented in this article were the basis for the study reported in the research articles entitled ‘A validated design methodology for a closed loop subsonic wind tunnel’ (Calautit et al., 2014) [1], which presented a systematic investigation into the design, simulation and analysis of flow parameters in a wind tunnel using Computational Fluid Dynamics (CFD). The authors evaluated the accuracy of replicating the flow characteristics for which the wind tunnel was designed using numerical simulation. Here, we detail the numerical and experimental set-up for the analysis of the closed-loop subsonic wind tunnel with an empty test section. PMID:26958641

  3. Pressure Distribution at Subsonic Speeds over the Forepart of Two Blunt Circular Cylinders

    NASA Technical Reports Server (NTRS)

    Lockwood, V. E.

    1975-01-01

    A wind tunnel investigation was made at subsonic speeds to determine the pressure distribution over the forward part of a circular cylinder. The cylinder was equipped with interchangeable faces, one having a flat face and one having a dome shaped face. The investigation was made over angle of attack range from -1 deg to 26 deg and a Mach number range from 0.30 to 0.89. Pressure coefficients are presented in tabular form and plotted data are presented for some selected angles of attack about the surface of the cylinder.

  4. Exploratory subsonic investigation of vortex-flap concept on arrow wing configuration

    NASA Technical Reports Server (NTRS)

    Rao, D. M.

    1981-01-01

    The drag reduction potential of a vortex flap concept, utilizing the thrust contribution of separation vortices maintained over leading edge flap surfaces, was explored in subsonic wind tunnel tests on a highly swept arrow wing configuration. Several flap geometries were tested in comparison with a previous study on the same model with leading edges drooped for attached flow. The most promising vortex flap arrangements produced drag reductions comparable with leading edge droop over a range of lift coefficients from 0.3 to 0.6 (untrimmed), and also indicated beneficial effects in the longitudinal and lateral static stability characteristics.

  5. Evaluation of laminar flow control systems for subsonic commercial transport aircraft: Executive summary

    NASA Technical Reports Server (NTRS)

    Pearce, W. E.

    1982-01-01

    An evaluation was made of laminar flow control (LFC) system concepts for subsonic commercial transport aircraft. Configuration design studies, performance analyses, fabrication development, structural testing, wind tunnel testing, and contamination-avoidance techniques were included. As a result of trade studies, a configuration with LFC on the upper wing surface only, utilizing an electron beam-perforated suction surface, and employing a retractable high-lift shield for contamination avoidance, was selected as the most practical LFC system. The LFC aircraft was then compared with an advanced turbulent aircraft designed for the same mission. This comparison indicated significant fuel savings.

  6. Evaluation of laminar flow control systems concepts for subsonic commercial transport aircraft

    NASA Technical Reports Server (NTRS)

    Pearce, W. E.

    1983-01-01

    An evaluation was made of laminar flow control (LFC) system concepts for subsonic commercial transport aircraft. Configuration design studies, performance analyses, fabrication development, structural testing, wind tunnel testing, and contamination-avoidance techniques were included. As a result of trade studies, a configuration with LFC on the upper wing surface only, utilizing an electron beam-perforated suction surface, and employing a retractable high-lift shield for contamination avoidance, was selected as the most practical LFC system. The LFC aircraft was then compared with an advanced turbulent aircraft designed for the same mission. This comparison indicated significant fuel savings and reduced direct operating cost benefits would result from using LFC.

  7. Preliminary measurements of velocity, density and total temperature fluctuations in compressible subsonic flow

    NASA Technical Reports Server (NTRS)

    Stainback, P. C.; Johnson, C. B.; Basnett, C. B.

    1983-01-01

    The heat transfer characteristics of a three-wire hot-wire probe operated with a constant temperature anemometer were investigated in the subsonic compressible flow regime. The sensitivity coefficients, with respect to velocity, density and total temperature, were measured and the results were used to calculate the velocity, density, and total temperature fluctuations in the test section of the Langley 0.3-m Transonic Cryogenic Tunnel (TCT). These results were extended to give estimates for fluctuations due to vorticity, sound, and entropy. In addition, attempts were made to determine the major source of disturbances in the 0.3-m TCT.

  8. Guidelines for Computing Longitudinal Dynamic Stability Characteristics of a Subsonic Transport

    NASA Technical Reports Server (NTRS)

    Thompson, Joseph R.; Frank, Neal T.; Murphy, Patrick C.

    2010-01-01

    A systematic study is presented to guide the selection of a numerical solution strategy for URANS computation of a subsonic transport configuration undergoing simulated forced oscillation about its pitch axis. Forced oscillation is central to the prevalent wind tunnel methodology for quantifying aircraft dynamic stability derivatives from force and moment coefficients, which is the ultimate goal for the computational simulations. Extensive computations are performed that lead in key insights of the critical numerical parameters affecting solution convergence. A preliminary linear harmonic analysis is included to demonstrate the potential of extracting dynamic stability derivatives from computational solutions.

  9. Automatic control study of the icing research tunnel refrigeration system

    NASA Technical Reports Server (NTRS)

    Kieffer, Arthur W.; Soeder, Ronald H.

    1991-01-01

    The Icing Research Tunnel (IRT) at the NASA Lewis Research Center is a subsonic, closed-return atmospheric tunnel. The tunnel includes a heat exchanger and a refrigeration plant to achieve the desired air temperature and a spray system to generate the type of icing conditions that would be encountered by aircraft. At the present time, the tunnel air temperature is controlled by manual adjustment of freon refrigerant flow control valves. An upgrade of this facility calls for these control valves to be adjusted by an automatic controller. The digital computer simulation of the IRT refrigeration plant and the automatic controller that was used in the simulation are discussed.

  10. Investigations and Experiments in the Guidonia Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1939-01-01

    This paper is a presentation of the experiments and equipment used in investigations at the Guidonia wind tunnel. The equipment consisted of: a number of subsonic and supersonic cones, an aerodynamic balance, and optical instruments operating on the Schlieren and interferometer principle.

  11. Aeropropulsion 1987. Session 5: Subsonic Propulsion Technology

    NASA Technical Reports Server (NTRS)

    1987-01-01

    NASA is conducting aeropropulsion research over a broad range of Mach numbers. In addition to the high-speed propulsion research described, major progress was recorded in research aimed at the subsonic flight regimes of interest to many commercial and military users. Recent progress and future directions in such areas as small engine technology, rotorcraft transmissions, icing, Hot Section Technology (HOST) and the Advanced Turboprop Program (ATP) are covered.

  12. Subsonic/supersonic aerodynamic characteristics for a tactical supercruiser

    NASA Technical Reports Server (NTRS)

    Capone, F. J.; Bare, E. A.; Hollenback, D.; Hutchison, R.

    1984-01-01

    A series of cooperative NASA-Langley/Boeing experimental investigations have been conducted to determine the aeropropulsive characteristics of an advanced tactical fighter designed for supersonic cruise. These investigations were conducted in the Langley 16-Foot Transonic and Lewis 10 x 10-Foot Supersonic Wind Tunnels at Mach numbers from 0.60 to 2.47. This fighter is a Mach 2.0, 49,000 pound class vehicle that features a close-coupled canard and underwing propulsion units that utilize multifunction two-dimensional exhaust nozzles. Tests were conducted to determine the basic aerodynamic characteristics of the configuration with flow-through nacelles in which the spillage effects of representative inlets were measured. The effects of thrust-induced forces on overall aerodynamic performance were evaluated with a series of multifunction nozzles installed on air-powered nacelles. An axisymmetric nozzle configuration was also tested to obtain comparative aeropropulsive performance. Trim aerodynamic characteristics for the flow-through and powered configurations and the effect of thrust vectoring at subsonic speeds are presented.

  13. Cavity Unsteady-Pressure Measurements at Subsonic and Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Tracy, Maureen B.; Plentovich, E. B.

    1997-01-01

    An experimental investigation was conducted in the Langley 8-Foot Transonic Pressure Tunnel to determine the flow characteristics of rectangular cavities with varying relative dimensions at subsonic and transonic speeds. Cavities were tested with width-to-depth ratios of 1, 4, 8, and 16 for length-to-depth ratios l/h of 1 through 17.5. The maximum cavity depth was 2.4 in., and the turbulent boundary layer approaching the cavity was approximately 0.5 in. thick. Unsteady- and mean static-pressure measurements were made at free-stream Mach numbers from 0.20 to 0.95 at a unit Reynolds number per foot of approximately 3 x 10(exp 6); however, only unsteady-pressure results are presented in this paper. Results indicate that as l/h increases, cavity flows changed from resonant to nonresonant with resonant amplitudes decreasing gradually. Resonant spectra are obtained largely in cavities with mean static-pressure distributions characteristic of open and transitional flows. Resonance sometimes occurred for closed flow. Increasing cavity width or decreasing cavity depth while holding l/h fixed had the effect of increasing resonant amplitudes and sometimes induced resonance. The effects due to changes in width are more pronounced. Decreasing Mach number has the effect of broadening the resonances.

  14. Supersonic Jet Exhaust Noise at High Subsonic Flight Speed

    NASA Technical Reports Server (NTRS)

    Norum, Thomas D.; Garber, Donald P.; Golub, Robert A.; Santa Maria, Odilyn L.; Orme, John S.

    2004-01-01

    An empirical model to predict the effects of flight on the noise from a supersonic transport is developed. This model is based on an analysis of the exhaust jet noise from high subsonic flights of the F-15 ACTIVE Aircraft. Acoustic comparisons previously attainable only in a wind tunnel were accomplished through the control of both flight operations and exhaust nozzle exit diameter. Independent parametric variations of both flight and exhaust jet Mach numbers at given supersonic nozzle pressure ratios enabled excellent correlations to be made for both jet broadband shock noise and jet mixing noise at flight speeds up to Mach 0.8. Shock noise correlated with flight speed and emission angle through a Doppler factor exponent of about 2.6. Mixing noise at all downstream angles was found to correlate well with a jet relative velocity exponent of about 7.3, with deviations from this behavior only at supersonic eddy convection speeds and at very high flight Mach numbers. The acoustic database from the flight test is also provided.

  15. Modeling the Launch Abort Vehicle's Subsonic Aerodynamics from Free Flight Testing

    NASA Technical Reports Server (NTRS)

    Hartman, Christopher L.

    2010-01-01

    An investigation into the aerodynamics of the Launch Abort Vehicle for NASA's Constellation Crew Launch Vehicle in the subsonic, incompressible flow regime was conducted in the NASA Langley 20-ft Vertical Spin Tunnel. Time histories of center of mass position and Euler Angles are captured using photogrammetry. Time histories of the wind tunnel's airspeed and dynamic pressure are recorded as well. The primary objective of the investigation is to determine models for the aerodynamic yaw and pitch moments that provide insight into the static and dynamic stability of the vehicle. System IDentification Programs for AirCraft (SIDPAC) is used to determine the aerodynamic model structure and estimate model parameters. Aerodynamic models for the aerodynamic body Y and Z force coefficients, and the pitching and yawing moment coefficients were identified.

  16. Shape memory alloy actuation effect on subsonic static aeroelastic deformation of composite cantilever plate

    NASA Astrophysics Data System (ADS)

    Hussein, A. M. H.; Majid, D. L. Abdul; Abdullah, E. J.

    2016-10-01

    Shape memory alloy (SMA) is one of the smart materials that have unique properties and used recently in several aerospace applications. SMAs are metallic alloys that can recover permanent strains when they are heated above a certain temperature. In this study, the effects of SMA actuation on the composite plate under subsonic aeroelastic conditions are examined. The wind tunnel test is carried out for two configurations of a cantilever shape memory alloy composite plate with a single SMA wire fixed eccentrically. Strain gage data for both bending and torsional strain are recorded and demonstrated during the aeroelastic test for active and non-active SMA wire in two locations. The cyclic actuation of the SMA wire embedded inside the composite plate is also investigated during the aeroelastic test. The results show reduction in both bending and torsional strain of the composite plate after activation of the SMA wire during the wind tunnel test.

  17. Subsonic Analysis of 0.04-Scale F-16XL Models Using an Unstructured Euler Code

    NASA Technical Reports Server (NTRS)

    Lessard, Wendy B.

    1996-01-01

    The subsonic flow field about an F-16XL airplane model configuration was investigated with an inviscid unstructured grid technique. The computed surface pressures were compared to wind-tunnel test results at Mach 0.148 for a range of angles of attack from 0 deg to 20 deg. To evaluate the effect of grid dependency on the solution, a grid study was performed in which fine, medium, and coarse grid meshes were generated. The off-surface vortical flow field was locally adapted and showed improved correlation to the wind-tunnel data when compared to the nonadapted flow field. Computational results are also compared to experimental five-hole pressure probe data. A detailed analysis of the off-body computed pressure contours, velocity vectors, and particle traces are presented and discussed.

  18. CFD Prediction of Magnus Effect in Subsonic to Supersonic Flight

    DTIC Science & Technology

    2009-09-01

    CFD Prediction of Magnus Effect in Subsonic to Supersonic Flight by James DeSpirito ARL-TR-4929 September 2009...of Magnus Effect in Subsonic to Supersonic Flight James DeSpirito Weapons and Materials Research Directorate, ARL...TITLE AND SUBTITLE CFD Prediction of Magnus Effect in Subsonic to Supersonic Flight 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT

  19. Implementation of a Particle Image Velocimetry System for Wind Tunnel Flowfield Measurements

    DTIC Science & Technology

    2014-12-01

    conducted to implement Particle Image Velocimetry (PIV) as a flow measurement technique in the 8’ x 10’ Subsonic Wind Tunnel at Naval Surface Warfare...discussed and summarized. PIV and SPIV were successfully demonstrated in the wind tunnel , and are now available as powerful flowfield measurement tools...for future test programs. 15. SUBJECT TERMS Particle Image Velocimetry, PIV, SPIV, wind tunnel , airwake, SFS2, flow seeding, flow survey 16

  20. Atmospheric Effects of Subsonic Aircraft: Interim Assessment Report of the Advanced Subsonic Technology Program

    NASA Technical Reports Server (NTRS)

    Friedl, Randall R. (Editor)

    1997-01-01

    This first interim assessment of the subsonic assessment (SASS) project attempts to summarize concisely the status of our knowledge concerning the impacts of present and future subsonic aircraft fleets. It also highlights the major areas of scientific uncertainty, through review of existing data bases and model-based sensitivity studies. In view of the need for substantial improvements in both model formulations and experimental databases, this interim assessment cannot provide confident numerical predictions of aviation impacts. However, a number of quantitative estimates are presented, which provide some guidance to policy makers.

  1. Subsonic Aerodynamics of Spinning and Non-Spinning Type 200 Lightcraft: Progress Report

    NASA Astrophysics Data System (ADS)

    Kenoyer, David A.; Myrabo, Leik N.

    2010-05-01

    A combined experimental and numerical investigation of subsonic aerodynamics for Type 200 laser lightcraft is underway for both spinning and non-spinning cases. A 12.2 cm diameter aluminum model with a "closed" annular airbreathing inlet was fitted to a sting balance in RPI's 61 cm by 61 cm subsonic wind tunnel. Aerodynamic forces and moments were measured first for the non-spinning case vs. angle of attack, at several freestream flow velocities (e.g., 30, 45, and 60 m/s) to assess Reynolds number effects. The CFD analysis was performed for 0-180° angles of attack for a fixed coordinate system (i.e., non-spinning Type 200 model), and predictions compared favorably with the experimental data. In the near future, for the spinning case, a brushless electric motor has been installed to rotate the wind tunnel model at 3000 to 13,000 RPM; Magnus force effects upon the coefficients (Cd, Cl, and Cm) are expected to reveal interesting departures from the non-spinning database in forthcoming experiments.

  2. Subsonic tests of an all-flush-pressure-orifice air data system

    NASA Technical Reports Server (NTRS)

    Larson, T. J.; Siemers, P. M., III

    1981-01-01

    The use of an all-flush-pressure-orifice array as a subsonic air data system was evaluated in flight and wind tunnel tests. Two orifice configurations were investigated. Both used orifices arranged in a cruciform pattern on the airplane nose. One configuration also used orifices on the sides of the fuselage for a source of static pressure. The all-nose-orifice configuration was similar to the shuttle entry air data system (SEADS). The flight data were obtained with a KC-135A airplane. The wind tunnel data were acquired with a 0.035-scale model of the KC-135A airplane. With proper calibration, several orifices on the vertical centerline of the vehicle's nose were found to be satisfactory for the determination of total pressure and angle of attack. Angle of sideslip could be accurately determined from pressure measurements made on the horizontal centerline of the aircraft. Orifice pairs were also found that provided pressure ratio relationships suitable for the determination of Mach number. The accuracy that can be expected for the air data determined with SEADS during subsonic orbiter flight is indicated.

  3. Analysis of supersonic combustion flow fields with embedded subsonic regions

    NASA Technical Reports Server (NTRS)

    Dash, S.; Delguidice, P.

    1972-01-01

    The viscous characteristic analysis for supersonic chemically reacting flows was extended to include provisions for analyzing embedded subsonic regions. The numerical method developed to analyze this mixed subsonic-supersonic flow fields is described. The boundary conditions are discussed related to the supersonic-subsonic and subsonic-supersonic transition, as well as a heuristic description of several other numerical schemes for analyzing this problem. An analysis of shock waves generated either by pressure mismatch between the injected fluid and surrounding flow or by chemical heat release is also described.

  4. V/STOL wind-tunnel testing

    NASA Technical Reports Server (NTRS)

    Koenig, D. G.

    1984-01-01

    Factors influencing effective program planning for V/STOL wind-tunnel testing are discussed. The planning sequence itself, which includes a short checklist of considerations that could enhance the value of the tests, is also described. Each of the considerations, choice of wind tunnel, type of model installation, model development and test operations, is discussed, and examples of appropriate past and current V/STOL test programs are provided. A short survey of the moderate to large subsonic wind tunnels is followed by a review of several model installations, from two-dimensional to large-scale models of complete aircraft configurations. Model sizing, power simulation, and planning are treated, including three areas is test operations: data-acquisition systems, acoustic measurements in wind tunnels, and flow surveying.

  5. Control of Wind Tunnel Operations Using Neural Net Interpretation of Flow Visualization Records

    NASA Technical Reports Server (NTRS)

    Buggele, Alvin E.; Decker, Arthur J.

    1994-01-01

    Neural net control of operations in a small subsonic/transonic/supersonic wind tunnel at Lewis Research Center is discussed. The tunnel and the layout for neural net control or control by other parallel processing techniques are described. The tunnel is an affordable, multiuser platform for testing instrumentation and components, as well as parallel processing and control strategies. Neural nets have already been tested on archival schlieren and holographic visualizations from this tunnel as well as recent supersonic and transonic shadowgraph. This paper discusses the performance of neural nets for interpreting shadowgraph images in connection with a recent exercise for tuning the tunnel in a subsonic/transonic cascade mode of operation. That mode was operated for performing wake surveys in connection with NASA's Advanced Subsonic Technology (AST) noise reduction program. The shadowgraph was presented to the neural nets as 60 by 60 pixel arrays. The outputs were tunnel parameters such as valve settings or tunnel state identifiers for selected tunnel operating points, conditions, or states. The neural nets were very sensitive, perhaps too sensitive, to shadowgraph pattern detail. However, the nets exhibited good immunity to variations in brightness, to noise, and to changes in contrast. The nets are fast enough so that ten or more can be combined per control operation to interpret flow visualization data, point sensor data, and model calculations. The pattern sensitivity of the nets will be utilized and tested to control wind tunnel operations at Mach 2.0 based on shock wave patterns.

  6. Analysis of an advanced ducted propeller subsonic inlet

    NASA Technical Reports Server (NTRS)

    Iek, Chanthy; Boldman, Donald R.; Ibrahim, Mounir

    1992-01-01

    A time marching Navier-Stokes code called PARC (PARC2D for 2-D/axisymmetric and PARC3D for 3-D flow simulations) was validated for an advanced ducted propeller (ADP) subsonic inlet. The code validation for an advanced ducted propeller (ADP) subsonic inlet. The code validation was implemented for a non-separated flow condition associated with the inlet operating at angles-of-attack of 0 and 25 degrees. The inlet test data were obtained in the 9 x 15 ft Low Speed Wind Tunnel at NASA Lewis Research Center as part of a cooperative study with Pratt and Whitney. The experimental study focused on the ADP inlet performance for take-off and approach conditions. The inlet was tested at a free stream Mach number of 0.2, at angles-of-attack between O and 35 degrees, and at a maximum propeller speed of 12,000 RPM which induced a corrected air flow rate of about 46 lb/sec based on standard day conditions. The computational grid and flow boundary conditions (BC) were based on the actual inlet geometry and the funnel flow conditions. At the propeller face, two types of BC's were applied: a mass flow BC and a fixed flow properties BC. The fixed flow properties BC was based on a combination of data obtained from the experiment and calculations using a potential flow code. Comparison of the computational results with the test data indicates that the PARC code with the propeller face fixed flow properties BC provided a better prediction of the inlet surface static pressures than the predictions when the mass flow BC was used. For an angle-of-attack of 0 degrees, the PARC2D code with the propeller face mass flow BC provided a good prediction of inlet static pressures except in the region of high pressure gradient. With the propeller face fixed flow properties BC, the PARC2D code provided a good prediction of the inlet static pressures. For an angle-of-attack of 25 degrees with the mass flow BC, the PARC3D code predicted statis pressures which deviated significantly from the test data

  7. User's guide to PANCOR: A panel method program for interference assessment in slotted-wall wind tunnels

    NASA Technical Reports Server (NTRS)

    Kemp, William B., Jr.

    1990-01-01

    Guidelines are presented for use of the computer program PANCOR to assess the interference due to tunnel walls and model support in a slotted wind tunnel test section at subsonic speeds. Input data requirements are described in detail and program output and general program usage are described. The program is written for effective automatic vectorization on a CDC CYBER 200 class vector processing system.

  8. Propulsion technology for an advanced subsonic transport

    NASA Technical Reports Server (NTRS)

    Beheim, M. A.; Antl, R. J.; Povolny, J. H.

    1972-01-01

    Engine design studies for future subsonic commercial transport aircraft were conducted in parallel with airframe studies. These studies surveyed a broad distribution of design variables, including aircraft configuration, payload, range, and speed, with particular emphasis on reducing noise and exhaust emissions without severe economic and performance penalties. The results indicated that an engine for an advanced transport would be similar to the currently emerging turbofan engines. Application of current technology in the areas of noise suppression and combustors imposed severe performance and economic penalties.

  9. Rectangular subsonic jet flow field measurements

    NASA Technical Reports Server (NTRS)

    Morrison, Gerald L.; Swan, David H.

    1990-01-01

    Flow field measurements of three subsonic rectangular cold air jets are presented. The three cases had aspect ratios of 1x2, 1x4 at a Mach number of 0.09 and an aspect ratio of 1x2 at a Mach number of 0.9. All measurements were made using a 3-D laser Doppler anemometer system. The data includes the mean velocity vector, all Reynolds stress tensor components, turbulent kinetic energy and velocity correlation coefficients. The data are presented in tabular and graphical form. No analysis of the measured data or comparison to other published data is made.

  10. 14 CFR 91.853 - Final compliance: Civil subsonic airplanes.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 2 2011-01-01 2011-01-01 false Final compliance: Civil subsonic airplanes... Noise Limits § 91.853 Final compliance: Civil subsonic airplanes. Except as provided in § 91.873, after... airplane subject to § 91.801(c) of this subpart, unless that airplane has been shown to comply with Stage...

  11. 14 CFR 91.853 - Final compliance: Civil subsonic airplanes.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 2 2010-01-01 2010-01-01 false Final compliance: Civil subsonic airplanes... Noise Limits § 91.853 Final compliance: Civil subsonic airplanes. Except as provided in § 91.873, after... airplane subject to § 91.801(c) of this subpart, unless that airplane has been shown to comply with Stage...

  12. FPGA development for high altitude subsonic parachute testing

    NASA Technical Reports Server (NTRS)

    Kowalski, James E.; Gromov, Konstantin G.; Konefat, Edward H.

    2005-01-01

    This paper describes a rapid, top down requirements-driven design of a Field Programmable Gate Array (FPGA) used in an Earth qualification test program for a new Mars subsonic parachute. The FPGA is used to process and control storage of telemetry data from multiple sensors throughout launch, ascent, deployment and descent phases of the subsonic parachute test.

  13. FPGA development for high altitude subsonic parachute testing

    NASA Technical Reports Server (NTRS)

    Kowalski, James E.; Konefat, Edward H.; Gromovt, Konstantin

    2005-01-01

    This paper describes a rapid, top down requirements-driven design of an FPGA used in an Earth qualification test program for a new Mars subsonic parachute. The FPGA is used to process and store data from multiple sensors at multiple rates during launch, ascent, deployment and descent phases of the subsonic parachute test.

  14. High altitude subsonic parachute field programmable gate array

    NASA Technical Reports Server (NTRS)

    Kowalski, James E.; Gromov, Konstantin; Konefat, Edward H.

    2005-01-01

    This paper describes a rapid, top down requirements-driven design of an FPGA used in an Earth qualification test program for a new Mars subsonic parachute. The FPGA is used to process and control storage of telemetry data from multiple sensors throughout; launch, ascent, deployment and descent phases of the subsonic parachute test.

  15. Wind tunnel testing of low-drag airfoils

    NASA Technical Reports Server (NTRS)

    Harvey, W. Donald; Mcghee, R. J.; Harris, C. D.

    1986-01-01

    Results are presented for the measured performance recently obtained on several airfoil concepts designed to achieve low drag by maintaining extensive regions of laminar flow without compromising high-lift performance. The wind tunnel results extend from subsonic to transonic speeds and include boundary-layer control through shaping and suction. The research was conducted in the NASA Langley 8-Ft Transonic Pressure Tunnel (TPT) and Low Turbulence Pressure Tunnel (LTPT) which have been developed for testing such low-drag airfoils. Emphasis is placed on identifying some of the major factors influencing the anticipated performance of low-drag airfoils.

  16. Study and evaluation of ferro-cement for use in wind tunnel construction

    NASA Technical Reports Server (NTRS)

    Larsen, H. J., Jr. (Compiler)

    1972-01-01

    The structural suitability and cost effectiveness of ferro-cement for large subsonic wind tunnel structures is investigated. This investigation was carried out in the following four main categories: (1) a state-of-the-art survey into the uses, properties, and costs of ferro-cement; (2) an evaluation of those ferro-cement properties critical to construction of large, subsonic wind tunnels, which have not been adequately established to date; (3) a laboratory testing program to determine preliminary values for those properties; and (4) a study to establish cost factors for ferro-cement as related to a preliminary construction scheme for a nacelle and shroud unit.

  17. Quantification of the Uncertainties for the Space Launch System Liftoff/Transition and Ascent Databases

    NASA Technical Reports Server (NTRS)

    Favaregh, Amber L.; Houlden, Heather P.; Pinier, Jeremy T.

    2016-01-01

    A detailed description of the uncertainty quantification process for the Space Launch System Block 1 vehicle configuration liftoff/transition and ascent 6-Degree-of-Freedom (DOF) aerodynamic databases is presented. These databases were constructed from wind tunnel test data acquired in the NASA Langley Research Center 14- by 22-Foot Subsonic Wind Tunnel and the Boeing Polysonic Wind Tunnel in St. Louis, MO, respectively. The major sources of error for these databases were experimental error and database modeling errors.

  18. Acoustic Prediction Methodology and Test Validation for an Efficient Low-Noise Hybrid Wing Body Subsonic Transport

    NASA Technical Reports Server (NTRS)

    Kawai, Ronald T. (Compiler)

    2011-01-01

    This investigation was conducted to: (1) Develop a hybrid wing body subsonic transport configuration with noise prediction methods to meet the circa 2007 NASA Subsonic Fixed Wing (SFW) N+2 noise goal of -52 dB cum relative to FAR 36 Stage 3 (-42 dB cum re: Stage 4) while achieving a -25% fuel burned compared to current transports (re :B737/B767); (2) Develop improved noise prediction methods for ANOPP2 for use in predicting FAR 36 noise; (3) Design and fabricate a wind tunnel model for testing in the LaRC 14 x 22 ft low speed wind tunnel to validate noise predictions and determine low speed aero characteristics for an efficient low noise Hybrid Wing Body configuration. A medium wide body cargo freighter was selected to represent a logical need for an initial operational capability in the 2020 time frame. The Efficient Low Noise Hybrid Wing Body (ELNHWB) configuration N2A-EXTE was evolved meeting the circa 2007 NRA N+2 fuel burn and noise goals. The noise estimates were made using improvements in jet noise shielding and noise shielding prediction methods developed by UC Irvine and MIT. From this the Quiet Ultra Integrated Efficient Test Research Aircraft #1 (QUIET-R1) 5.8% wind tunnel model was designed and fabricated.

  19. Subsonic Glideback Rocket Demonstrator Flight Testing

    NASA Technical Reports Server (NTRS)

    DeTurris, Dianne J.; Foster, Trevor J.; Barthel, Paul E.; Macy, Daniel J.; Droney, Christopher K.; Talay, Theodore A. (Technical Monitor)

    2001-01-01

    For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.

  20. Rectangular subsonic jet flow field measurements

    NASA Technical Reports Server (NTRS)

    Morrison, Gerald L.; Swan, David H.

    1989-01-01

    Flow field measurements are presented of 3 subsonic rectangular cold air jets. The 3 cases presented had aspect ratios of 1 x 2, 1 x 4 at a Mach number of 0.09 and an aspect ratio of 1 x 2 at a Mach number of 0.9. All measurements were made using a 3-D laser Doppler anemoneter system. The presented data includes the mean velocity vector, all Reynolds stress tensor components, turbulent kinetic energy and velocity correlation coefficients. The data is presented in tabular and graphical form. No analysis of the measured data or comparison to other published data is made. All tabular data are available in ASCII format on MS-DOS compatible disks.

  1. a Numerical Model for Subsonic Acoustic Choking.

    NASA Astrophysics Data System (ADS)

    Walkington, Noel John

    In aircraft turbofan inlets, fan generated noise is observed experimentally to be significantly attenuated at high subsonic inlet Mach numbers. This phenomenon cannot be predicted by linear acoustic theory. In order to study the physical process by which this may occur, a numerical algorithm has been developed to solve a related nonlinear problem in one dimensional gas dynamics. The nonlinear solution admits the possibility of wave steepening and shock waves. Approximate solutions are obtained using several finite difference schemes. The boundary conditions required to model an acoustic source and an anechoic termination are developed. The numerical solutions agree closely with those obtained using the method of matched asymptotic expansions. Solutions involving shock waves exhibit a large reduction in the ratio of transmitted to incident power. This offers an explanation for acoustic choking. The results indicate that more power is dissipated as the Mach number, sound amplitude and frequency are increased. These observations are in agreement with those observed experimentally.

  2. On the stability of subsonic thermal fronts

    SciTech Connect

    Ibanez S, Miguel H.; Shchekinov, Yuri; Bessega L, Maria C.

    2005-08-15

    The stability of subsonic thermal fronts against corrugation is analyzed and an exact dispersion relation is obtained taking into account the compressibility of the gas. For heat fronts, this dispersion equation has an unstable root ({omega}{sub ex}) corresponding to the Landau-Darrieus unstable mode ({omega}{sub 0}) modified by the compressional effects. In particular, the exact solution shows a conspicuous maximum very close to the value of the intake Mach number M{sub 1} at which a Chapman-Jouguet deflagration wave behind the heat front is formed. Cooling fronts are stable for corrugation-like disturbances. A maximum damping as well as a maximum in the frequency occur at a value of M{sub 1} depending on the value of the normalized cooling q.

  3. Wind Tunnel Measured Effects on a Twin-Engine Short-Haul Transport Caused by Simulated Ice Accretions: Data Report

    NASA Technical Reports Server (NTRS)

    Reehorst, Andrew; Potapczuk, Mark; Ratvasky, Thomas; Laflin, Brenda Gile

    1997-01-01

    The purpose of this report is to release the data from the NASA Langley/Lewis 14 by 22 foot wind tunnel test that examined icing effects on a 1/8 scale twin-engine short-haul jet transport model. Presented in this document are summary data from the major configurations tested. The entire test database in addition to ice shape and model measurements is available as a data supplement in CD-ROM form. Data measured and presented are: wing pressure distributions, model force and moment, and wing surface flow visualization.

  4. Wind tunnel investigation of the aerodynamic characteristics of symmetrically deflected ailerons of the F-8C airplane. [conducted in the Langley 8-foot transonic pressure tunnel

    NASA Technical Reports Server (NTRS)

    Gera, J.

    1977-01-01

    A .042-scale model of the F-8C airplane was investigated in a transonic wind tunnel at high subsonic Mach numbers and a range of angles of attack between-3 and 20 degrees. The effect of symmetrically deflected ailerons on the longitudinal aerodynamic characteristics was measured. Some data were also obtained on the lateral control effectiveness of asymmetrically deflected horizontal tail surfaces.

  5. Low Subsonic Pressure Distributions on Three Rigid Wings Simulating Paragliders with Varied Canopy Curvature and Leading-Edge Sweep

    NASA Technical Reports Server (NTRS)

    Fournier, Paul G.; Bell, B. Ann

    1961-01-01

    An investigation has been made in the Langley 7- by 10-foot transonic tunnel to determine the subsonic pressure distribution of three paraglider models through an angle-of-attack range from 0 deg to 74 deg. Three rigid metal models simulated a 45 deg basic flat planform paraglider with leading-edge sweep angles of 61.6 deg, 52.5 deg, and 48.6 deg. These configurations resulted in one-half-circle, one-third-circle, and one-quarter-circle semispan trailing-edge curvature when viewed from downstream. The results of the investigation are presented as curves of chordwise pressure distributions at four spanwise locations.

  6. Test-to-Test Repeatability of Results From a Subsonic Wing-Body Configuration in the National Transonic Facility

    NASA Technical Reports Server (NTRS)

    Mineck, Raymond E.; Pendergraft, Odis C., Jr.

    2000-01-01

    Results from three wind tunnel tests in the National Transonic Facility of a model of an advanced-technology, subsonic-transport wing-body configuration have been analyzed to assess the test-to-test repeatability of several aerodynamic parameters. The scatter, as measured by the prediction interval, in the longitudinal force and moment coefficients increases as the Mach number increases. Residual errors with and without the ESP tubes installed suggest a bias leading to lower drag with the tubes installed. Residual errors as well as average values of the longitudinal force and moment coefficients show that there are small bias errors between the different tests.

  7. Subsonic stability and control flight test results of the Space Shuttle /tail cone off/

    NASA Technical Reports Server (NTRS)

    Cooke, D. R.

    1980-01-01

    The subsonic stability and control testing of the Space Shuttle Orbiter in its two test flights in the tailcone-off configuration is discussed, and test results are presented. Flight test maneuvers were designed to maximize the quality and quantity of stability and control data in the minimal time allotted using the Space Shuttle Functional Simulator and the Modified Maximum Likelihood Estimator (MMLE) programs, and coefficients were determined from standard sensor data sets using the MMLE, despite problems encountered in timing due to the different measurement systems used. Results are included for lateral directional and longitudinal maneuvers as well as the Space Shuttle aerodynamic data base obtained using the results of wind tunnel tests. The flight test data are found to permit greater confidence in the data base since the differences found are well within control system capability. It is suggested that the areas of major differences, including lateral directional data with open speedbrake, roll due to rudder and normal force due to elevon, be investigated in any further subsonic flight testing. Improvements in sensor data and data handling techniques for future orbital test flights are indicated.

  8. Method for the prediction of the installation aerodynamics of a propfan at subsonic speeds: User manual

    NASA Technical Reports Server (NTRS)

    Chandrasekaran, B.

    1986-01-01

    This document is the user's guide for the method developed earlier for predicting the slipstream wing interaction at subsonic speeds. The analysis involves a subsonic panel code (HESS code) modified to handle the propeller onset flow. The propfan slipstream effects are superimposed on the normal flow boundary condition and are applied over the surface washed by the slipstream. The effects of the propeller slipstream are to increase the axial induced velocity, tangential velocity, and a total pressure rise in the wake of the propeller. Principles based on blade performance theory, momentum theory, and vortex theory were used to evaluate the slipstream effects. The code can be applied to any arbitrary three dimensional geometry, expressed in the form of HESS input format. The code can handle a propeller alone configuration or a propeller/nacelle/airframe configuration, operating up to high subcritical Mach numbers over a range of angles of attack. Inclusion of a viscous modelling is briefly outlined. Wind tunnel results/theory comparisons are included as examples for the application of the code to a generic supercritical wing/overwing Nacelle with a powered propfan. A sample input/output listing is provided.

  9. Multi-Mission Earth Vehicle Subsonic Dynamic Stability Testing and Analyses

    NASA Technical Reports Server (NTRS)

    Glaab, Louis J.; Fremaux, C. Michael

    2013-01-01

    Multi-Mission Earth Entry Vehicles (MMEEVs) are blunt-body vehicles designed with the purpose of transporting payloads from outer space to the surface of the Earth. To achieve high-reliability and minimum weight, MMEEVs avoid use of limited-reliability systems, such as parachutes, retro-rockets, and reaction control systems and rely on the natural aerodynamic stability of the vehicle throughout the Entry, Descent, and Landing (EDL) phase of flight. The Multi-Mission Systems Analysis for Planetary Entry (M-SAPE) parametric design tool is used to facilitate the design of MMEEVs for an array of missions and develop and visualize the trade space. Testing in NASA Langley?s Vertical Spin Tunnel (VST) was conducted to significantly improve M-SAPE?s subsonic aerodynamic models. Vehicle size and shape can be driven by entry flight path angle and speed, thermal protection system performance, terminal velocity limitations, payload mass and density, among other design parameters. The objectives of the VST testing were to define usable subsonic center of gravity limits, and aerodynamic parameters for 6-degree-of-freedom (6-DOF) simulations, for a range of MMEEV designs. The range of MMEEVs tested was from 1.8m down to 1.2m diameter. A backshell extender provided the ability to test a design with a much larger payload for the 1.2m MMEEV.

  10. Measurement of the static and dynamic coefficients of a cross-type parachute in subsonic flow

    NASA Technical Reports Server (NTRS)

    Shpund, Zalman; Levin, Daniel

    1991-01-01

    An experimental parametric investigation of the aerodynamic qualities of cross-type parachutes was performed in a subsonic wind tunnel, using a new experimental technique. This investigation included the measurement of the static and dynamic aerodynamic coefficients, utilizing the measuring apparatus modified specifically for this type of testing. It is shown that the static aerodynamic coefficients of several configurations are in good agreement with available data, and assisted in validating the experimental technique employed. Two configuration parameters were varied in the static tests, the cord length and the canopy aspect ratio, with both parameters having a similar effect on the drag measurement, i.e., any increase in either of them increased the effective blocking area, and therefore the axial force.

  11. Aerodynamic Characteristics of an Aerospace Vehicle During a Subsonic Pitch-Over Maneuver

    NASA Technical Reports Server (NTRS)

    Kleb, William L.

    1996-01-01

    Time-dependent CFD has been used to predict aerospace vehicle aerodynamics during a subsonic rotation maneuver. The inviscid 3D3U code is employed to solve the 3-D unsteady flow field using an unstructured grid of tetrahedra. As this application represents a challenge to time-dependent CFD, observations concerning spatial and temporal resolution are included. It is shown that even for a benign rotation rate, unsteady aerodynamic effects are significant during the maneuver. Possibly more significant, however, the rotation maneuver creates ow asymmetries leading to yawing moment, rolling moment, and side force which are not present in the quasi-steady case. A series of steady solutions at discrete points in the maneuver are also computed for comparison with wind tunnel measurements and as a means of quantifying unsteady effects.

  12. Influence of configuration details on the subsonic characteristics of a space shuttle orbiter design

    NASA Technical Reports Server (NTRS)

    Decker, J. P.; Phillips, W. P.

    1974-01-01

    An investigation was conducted in the Langley low-turbulence pressure tunnel of a model of a space shuttle orbiter design in order to determine the influence of minor configuration geometric details on the aerodynamic characteristics at subsonic speeds. A plane wing was tested with a small planform fillet; a twisted wing was tested with both a small and a large planform fillet. Tailored attitude-control propulsion-system wing-tip and body pods, trisegmented elevons, and canopy effects were also investigated. The tests were conducted at angles of attack from -3 deg to 24 deg for sideslip angles of 0 deg and 6 deg and at a Mach number of 0.25.

  13. Design and Development of Low-Cost Wind Tunnel for Educational Purpose

    NASA Astrophysics Data System (ADS)

    Yong, T. H.; Dol, S. S.

    2015-04-01

    The presence of wind tunnel is undoubtedly bringing infinite possibilities to studying and understanding complex fluid flows. However, commercial wind tunnel is expensive and only limited to highly-focus researchers or exclusive institutions. This paper discusses the design and development of a low-cost, educational-purposed, open-typed subsonic wind tunnel. In this work, an open-typed subsonic wind tunnel is designed with the aim of achieving turbulent intensity (in the working section) below or equal to 5%, within the budget of RM 1500 and a working speed of 6 m/s - 8 m/s to meet the Reynolds number in the order of 105. The conceptual design was studied using Ansys Fluent 14.5 and the optimal design was then developed and experimentally verified.

  14. The future of very large subsonic transports

    NASA Technical Reports Server (NTRS)

    Justice, R. Steven; Hays, Anthony P.; Parrott, Ed L.

    1996-01-01

    The Very Large Subsonic Transport (VLST) is a multi-use commercial passenger, commercial cargo, and military airlifter roughly 50% larger than the current Lockheed C-5 and Boeing 747. Due to the large size and cost of the VLST, it is unlikely that the commercial market can support more than one aircraft production line, while declining defense budgets will not support a dedicated military VLST. A successful VLST must therefore meet airline requirements for more passenger and cargo capacity on congested routes into slot-limited airports and also provide a cost effective heavy airlift capacity to support the overseas deployment of US military forces. A successful VLST must satisfy three key missions: commercial passenger service with nominal seating capacity at a minimum of 650 passengers with a range capability of 7,000 to 10,000 miles; commercial air cargo service for containerized cargo to support global manufacturing of high value added products, 'just-in-time' parts delivery, and the general globalization of trade; and military airlift with adequate capacity to load current weapon systems, with minimal break-down, over global ranges (7,000 to 10,000 miles) required to reach the operational theater without need of overseas bases and midair refueling. The development of the VLST poses some technical issues specific to large aircraft, but also key technologies applicable to a wide range of subsonic transport aircraft. Key issues and technologies unique to the VLST include: large composite structures; dynamic control of a large, flexible structure; aircraft noise requirements for aircraft over 850,000 pounds; and increased aircraft separation due to increased wake vortex generation. Other issues, while not unique to the VLST, will critically impact the ability to build an efficient and affordable aircraft include: active control systems: Fly-By-Light/Power-By-Wire (FBL/PBW); high lift systems; flight deck associate systems; laminar flow; emergency egress; and

  15. User's guide to STIPPAN: A panel method program for slotted tunnel interference prediction

    NASA Technical Reports Server (NTRS)

    Kemp, W. B., Jr.

    1985-01-01

    Guidelines are presented for use of the computer program STIPPAN to simulate the subsonic flow in a slotted wind tunnel test section with a known model disturbance. Input data requirements are defined in detail and other aspects of the program usage are discussed in more general terms. The program is written for use in a CDC CYBER 200 class vector processing system.

  16. Study of several factors affecting crew escape trajectories from the Space Shuttle Orbiter at low-subsonic speeds

    NASA Technical Reports Server (NTRS)

    Scallion, W. I.; Spencer, B., Jr.; Ware, G. M.; Phillips, W. P.; Powell, R. W.; Young, J. C.; Gentry, G. L., Jr.; Applin, Z. T.

    1985-01-01

    Factors affecting the bailout characteristics from the space shuttle orbiter at low-subsonic speeds were investigated. In the 12-foot low-speed tunnel and the 4 by 7-meter tunnel with 0.03-scale models. The effect of crew-model exit velocity, body position, and body weight were studied with egress from the main side hatch with the orbiter upright and from the upper cabin hatch with the orbiter inverted. Crew model drag and flow field measurements around the orbiter were estimated. The high-angle-of-attack trim characteristics of the orbiter was determined by force tests in an attempt to improve bailout conditions. A computer simulation was made to evaluate the maneuver necessary to attain the high-angle-of-attack trim.

  17. An experimental investigation of nacelle-pylon installation on an unswept wing at subsonic and transonic speeds

    NASA Technical Reports Server (NTRS)

    Carlson, J. R.; Compton, W. B., III

    1984-01-01

    A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.

  18. SHARP: Subsonic High Altitude Research Platform

    NASA Technical Reports Server (NTRS)

    Beals, Todd; Burton, Craig; Cabatan, Aileen; Hermano, Christine; Jones, Tom; Lee, Susan; Radloff, Brian

    1991-01-01

    The Universities Space Research Association is sponsoring an undergraduate program which is geared to designing an aircraft that can study the ozone layer at the equator. This aircraft must be able to satisfy four mission profiles. Mission one is a polar mission that ranges from Chile to the South Pole and back to Chile, a total range of 6000 n.mi. at 100,000 ft with a 2500 lb payload. The second mission is also a polar mission, with an altitude of 70,000 ft and an increased payload of 4000 lbs. For the third mission, the aircraft will takeoff at NASA Ames, cruise at 100,000 ft carrying a 2500 lb payload, and land at Puerto Montt, Chile. The final mission requires the aircraft to take off at NASA Ames, cruise at 100,000 ft with a 1000 lb payload, make an excursion to 120,000 ft, and land at Howard AFB, Panama. Three missions require that a subsonic Mach number be maintained due to constraints imposed by the air sampling equipment. The aircraft need not be manned for all four missions. Three aircraft configurations have been determined to be the most suitable for meeting the above requirements. In the event that a requirement cannot be obtained within the given constraints, recommendations for proposal modifications are given.

  19. The numerical simulation of subsonic flutter

    NASA Technical Reports Server (NTRS)

    Strganac, Thomas W.; Mitchum, Maria V.; Mook, Dean T.

    1987-01-01

    The present paper describes a numerical simulation of unsteady, subsonic aeroelastic responses. The technique accounts for aerodynamic nonlinearities associated with angles of attack, vortex-dominated flow, static deformations, and unsteady behavior. The fluid and the wing together are treated as a single dynamic system, and the equations of motion for the structure and flowfield are integrated simultaneously and interactively in the time domain. The method employs an iterative scheme based on a predictor-corrector technique. The aerodynamic loads are computed by the general unsteady vortex-lattice method and are determined simultaneously with the motion of the wing. Two models are used to demonstrate the technique: a rigid wing on an elastic support experiencing plunge and pitch about the elastic axis, and a continuous wing rigidly supported at the root chord experiencing spanwise bending and twisting. The time domain solution coupled with the unsteady vortex-lattice method provides the capability of graphically depicting wing and wake motion. Several graphs that illustrate the time domain behavior of the wing and wake are presented.

  20. Subsonic-transonic stall flutter study

    NASA Technical Reports Server (NTRS)

    Stardter, H.

    1979-01-01

    The objective of the Subsonic/Transonic Stall Flutter Program was to obtain detailed measurements of both the steady and unsteady flow field surrounding a rotor and the mechanical state of the rotor while it was operating in both steady and flutter modes to provide a basis for future analysis and for development of theories describing the flutter phenomenon. The program revealed that while all blades flutter at the same frequency, they do not flutter at the same amplitude, and their interblade phase angles are not equal. Such a pattern represents the superposition of a number of rotating nodal diameter patterns, each characterized by a different amplitude and different phase indexing, but each rotating at a speed that results in the same flutter frequency as seen in the rotor system. Review of the steady pressure contours indicated that flutter may alter the blade passage pressure distribution. The unsteady pressure amplitude contour maps reveal regions of high unsteady pressure amplitudes near the leading edge, lower amplitudes near the trailing.

  1. Advanced Subsonic Airplane Design and Economic Studies

    NASA Technical Reports Server (NTRS)

    Liebeck, Robert H.; Andrastek, Donald A.; Chau, Johnny; Girvin, Raquel; Lyon, Roger; Rawdon, Blaine K.; Scott, Paul W.; Wright, Robert A.

    1995-01-01

    A study was made to examine the effect of advanced technology engines on the performance of subsonic airplanes and provide a vision of the potential which these advanced engines offered. The year 2005 was selected as the entry-into-service (EIS) date for engine/airframe combination. A set of four airplane classes (passenger and design range combinations) that were envisioned to span the needs for the 2005 EIS period were defined. The airframes for all classes were designed and sized using 2005 EIS advanced technology. Two airplanes were designed and sized for each class: one using current technology (1995) engines to provide a baseline, and one using advanced technology (2005) engines. The resulting engine/airframe combinations were compared and evaluated on the basis on sensitivity to basic engine performance parameters (e.g. SFC and engine weight) as well as DOC+I. The advanced technology engines provided significant reductions in fuel burn, weight, and wing area. Average values were as follows: reduction in fuel burn = 18%, reduction in wing area = 7%, and reduction in TOGW = 9%. Average DOC+I reduction was 3.5% using the pricing model based on payload-range index and 5% using the pricing model based on airframe weight. Noise and emissions were not considered.

  2. NASA's Subsonic Jet Transport Noise Reduction Research

    NASA Technical Reports Server (NTRS)

    Powell, Clemans A.; Preisser, John S.

    2000-01-01

    Although new jet transport airplanes in today s fleet are considerably quieter than the first jet transports introduced about 40 years ago, airport community noise continues to be an important environmental issue. NASA s Advanced Subsonic Transport (AST) Noise Reduction program was begun in 1994 as a seven-year effort to develop technology to reduce jet transport noise 10 dB relative to 1992 technology. This program provides for reductions in engine source noise, improvements in nacelle acoustic treatments, reductions in the noise generated by the airframe, and improvements in the way airplanes are operated in the airport environs. These noise reduction efforts will terminate at the end of 2001 and it appears that the objective will be met. However, because of an anticipated 3-8% growth in passenger and cargo operations well into the 21st Century and the slow introduction of new the noise reduction technology into the fleet, world aircraft noise impact will remain essentially constant until about 2020 to 2030 and thereafter begin to rise. Therefore NASA has begun planning with the Federal Aviation Administration, industry, universities and environmental interest groups in the USA for a new noise reduction initiative to provide technology for significant further reductions.

  3. The Cylinder and Semicylinder in Subsonic Flow

    NASA Technical Reports Server (NTRS)

    Bingham, Harry J.; Weimer, David K..; Griffith, Wayland

    1952-01-01

    In studying the diffraction of shock waves around various two-dimensional obstacles we have observed that flow separation and the formation of vortices contributes in an important way to transient loading of the obstacle. The cases of a cylinder and semicylinder are especially interesting because the breakaway point is not clearly defined as it is for objects having sharp corners. Accordingly a number of experiments have been made in the shock tube to observe the influence of Reynolds number and Mach number on the transient flow patterns about a cylinder and about a semicylinder mounted on a smooth plane. Some differences might be anticipated since the plane would impose a symmetry on the flow and produce a viscous boundary layer for which there is no counterpart with the cylinder. In the course of these experiments it was noted that a condition of steady subsonic flow about both the cylinder and semicylinder was approached. Thus a comparison with von Karrnan's theoretical calculation of the drag on a cylinder, from certain characteristics of its wake or "vortex street", was undertaken.

  4. Aeroservoelastic Wind-Tunnel Test of the SUGAR Truss Braced Wing Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Scott, Robert C.; Allen, Timothy J.; Funk, Christie J.; Castelluccio, Mark A.; Sexton, Bradley W.; Claggett, Scott; Dykman, John; Coulson, David A.; Bartels, Robert E.

    2015-01-01

    The Subsonic Ultra Green Aircraft Research (SUGAR) Truss-Braced Wing (TBW) aeroservoelastic (ASE) wind-tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) and was completed in April, 2014. The primary goals of the test were to identify the open-loop flutter boundary and then demonstrate flutter suppression. A secondary goal was to demonstrate gust load alleviation (GLA). Open-loop flutter and limit cycle oscillation onset boundaries were identified for a range of Mach numbers and various angles of attack. Two sets of control laws were designed for the model and both sets of control laws were successful in suppressing flutter. Control laws optimized for GLA were not designed; however, the flutter suppression control laws were assessed using the TDT Airstream Oscillation System. This paper describes the experimental apparatus, procedures, and results of the TBW wind-tunnel test. Acquired system ID data used to generate ASE models is also discussed.2 study.

  5. Recognition Tunneling

    PubMed Central

    Lindsay, Stuart; He, Jin; Sankey, Otto; Hapala, Prokop; Jelinek, Pavel; Zhang, Peiming; Chang, Shuai; Huang, Shuo

    2010-01-01

    Single molecules in a tunnel junction can now be interrogated reliably using chemically-functionalized electrodes. Monitoring stochastic bonding fluctuations between a ligand bound to one electrode and its target bound to a second electrode (“tethered molecule-pair” configuration) gives insight into the nature of the intermolecular bonding at a single molecule-pair level, and defines the requirements for reproducible tunneling data. Simulations show that there is an instability in the tunnel gap at large currents, and this results in a multiplicity of contacts with a corresponding spread in the measured currents. At small currents (i.e. large gaps) the gap is stable, and functionalizing a pair of electrodes with recognition reagents (the “free analyte” configuration) can generate a distinct tunneling signal when an analyte molecule is trapped in the gap. This opens up a new interface between chemistry and electronics with immediate implications for rapid sequencing of single DNA molecules. PMID:20522930

  6. Subsonic potential aerodynamics for complex configurations - A general theory

    NASA Technical Reports Server (NTRS)

    Morino, L.; Kuo, C.-C.

    1974-01-01

    A general theory of subsonic potential aerodynamic flow around a lifting body having arbitrary shape and motion is presented. By using the Green function method, an integral representation for the velocity potential is obtained for both supersonic and subsonic flow. Under the small perturbation assumption, the potential at any point in the field depends only upon the values of the potential and its normal derivative on the surface of the body. On the surface of the body, this representation reduces to an integro-differential equation relating the potential and its normal derivative (which is known from the boundary conditions) on the surface. The theory is applied to finite-thickness wings in subsonic steady and oscillatory flows.

  7. Subsonic loads on wings having sharp leading edges and tips

    NASA Technical Reports Server (NTRS)

    Kandil, O. A.; Mook, D. T.; Nayfeh, A. H.

    1976-01-01

    A vortex-lattice method for predicting the aerodynamics of wings having separation at the sharp edges in incompressible flows is extended to compressible subsonic flows using a modified Prandtl-Glauert transformation. Numerical results showing the effect of freestream Mach number on the aerodynamic coefficients are compared with available experimental data for several planforms. It is shown that the proposed method is suitable for predicting the aerodynamic loads on low-aspect wings at moderate angles of attack for high subsonic freestream Mach number. The method is limited to angles of attack up to 12 deg for high subsonic freestream Mach number and to angles of attack up to 20 deg for Mach number not exceeding 0.5.

  8. Flight-determined derivatives and dynamic characteristics for the HL-10 lifting body vehicle at subsonic and transonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Strutz, L. W.

    1972-01-01

    The HL-10 lifting body stability and control derivatives were determined by using an analog-matching technique and compared with derivatives obtained from wind-tunnel results. The flight derivatives were determined as a function of angle of attack for a subsonic configuration at Mach 0.7 and for a transonic configuration at Mach 0.7, 0.9, and 1.2. At an angle of attack of 14 deg, data were obtained for a Mach number range from 0.6 to 1.4. The flight and wind-tunnel derivatives were in general agreement, with the possible exception of the longitudinal and lateral damping derivatives. Some differences were noted between the vehicle dynamic response characteristics calculated from flight-determined derivatives and those predicted by the wind-tunnel results. However, the only difference the pilots noted between the response of the vehicle in flight and the response of a simulator programed with wind-tunnel-predicted data was that the damping generally was higher in the flight vehicle.

  9. Overview of the Langley subsonic research effort on SCR configuration

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Thomas, J. D.; Huffman, J. K.; Weston, R. P.; Schoonover, W. E., Jr.; Gentry, C. L., Jr.

    1980-01-01

    Recent advances achieved in the subsonic aerodynamics of low aspect ratio, highly swept wing designs are summarized. The most significant of these advances was the development of leading edge deflection concepts which effectively reduce leading edge flow separation. The improved flow attachment results in substantial improvements in low speed performance, significant delay of longitudinal pitch up, increased trailing edge flap effectiveness, and increased lateral control capability. Various additional theoretical and/or experimental studies are considered which, in conjunction with the leading edge deflection studies, form the basis for future subsonic research effort.

  10. Second-order subsonic airfoil theory including edge effects

    NASA Technical Reports Server (NTRS)

    Van Dyke, Milton D

    1956-01-01

    Several recent advances in plane subsonic flow theory are combined into a unified second-order theory for airfoil sections of arbitrary shape. The solution is reached in three steps: the incompressible result is found by integration, it is converted into the corresponding subsonic compressible result by means of the second-order compressibility rule, and it is rendered uniformly valid near stagnation points by further rules. Solutions for a number of airfoils are given and are compared with the results of other theories and of experiment. A straight-forward computing scheme is outlined for calculating the surface velocities and pressures on any airfoil at any angle of attack

  11. Subsonic flow over thin oblique airfoils at zero lift

    NASA Technical Reports Server (NTRS)

    Jones, Robert T

    1948-01-01

    A previous report gave calculations for the pressure distribution over thin oblique airfoils at supersonic speed. The present report extends the calculations to subsonic speeds. It is found that the flows again can be obtained by the superposition of elementary conical flow fields. In the case of the swept-back wing the pressure distributions remain qualitatively similar at subsonic and supersonic speeds. Thus a distribution similar to the Ackeret type of distribution appears on the root sections of the swept-back wing at Mach=0. The resulting positive pressure drag on the root section is balanced by negative drags on outboard sections.

  12. The VIBRA-8 Subsonic Aerodynamic Nuclear Gust Vulnerability Code.

    DTIC Science & Technology

    1979-05-01

    entries along with AA(a=O, M) as given by Eq. (12) to obtain f NP(s-*). Data for the symmetric 64A006 and 64A010 airfoils (6 and 10 percent thick...N.A.C.A. Report 1977, 1952. 5. Stivers, L.S., Jr., Effects of Subsonic Mach Number on the Forces and Pressure Distributions on Four NACA 64A-Series...Airfoil Sections at Angles of Attack as High as 280, NACA TN3162, March 1954. 6. Axelson, J.A., and Haacker, J.F., Subsonic Wing Loadings on a 450 Sweptback

  13. Determination of forced convective heat transfer coefficients for subsonic flows over heated asymmetric NANA 4412 airfoil

    NASA Astrophysics Data System (ADS)

    Dag, Yusuf

    Forced convection over traditional surfaces such as flat plate, cylinder and sphere have been well researched and documented. Data on forced convection over airfoil surfaces, however, remain very scanty in literature. High altitude vehicles that employ airfoils as lifting surfaces often suffer leading edge ice accretions which have tremendous negative consequences on the lifting capabilities and stability of the vehicle. One of the ways of mitigating the effect of ice accretion involves judicious leading edge convective cooling technique which in turn depends on the accuracy of convective heat transfer coefficient used in the analysis. In this study empirical investigation of convective heat transfer measurements on asymmetric airfoil is presented at different angle of attacks ranging from 0° to 20° under subsonic flow regime. The top and bottom surface temperatures are measured at given points using Senflex hot film sensors (Tao System Inc.) and used to determine heat transfer characteristics of the airfoils. The model surfaces are subjected to constant heat fluxes using KP Kapton flexible heating pads. The monitored temperature data are then utilized to determine the heat convection coefficients modelled empirically as the Nusselt Number on the surface of the airfoil. The experimental work is conducted in an open circuit-Eiffel type wind tunnel, powered by a 37 kW electrical motor that is able to generate subsonic air velocities up to around 41 m/s in the 24 square-inch test section. The heat transfer experiments have been carried out under constant heat flux supply to the asymmetric airfoil. The convective heat transfer coefficients are determined from measured surface temperature and free stream temperature and investigated in the form of Nusselt number. The variation of Nusselt number is shown with Reynolds number at various angles of attacks. It is concluded that Nusselt number increases with increasing Reynolds number and increase in angle of attack from 0

  14. Investigation of aerodynamic characteristics of subsonic wings

    NASA Technical Reports Server (NTRS)

    Dejarnette, F. R.; Frink, N. T.

    1979-01-01

    An analytical strake design procedure is investigated. A numerical solution to the governing strake design equation is used to generate a series of strakes which are tested in a water tunnel to study their vortex breakdown characteristics. The strakes are scaled for use on a half-scale model of the NASA-LaRC general research fuselage with a 44 degrees trapezoidal wing. An analytical solution to the governing design equation is obtained. The strake design procedure relates the potential-flow leading-edge suction and pressure distributions to vortex stability. Several suction distributions are studied and those which are more triangular and peak near the tip generate strakes that reach higher angles of attack before vortex breakdown occurs at the wing trailing edge. For the same suction distribution, a conical rather than three dimensional pressure specification results in a better strake shape as judged from its vortex breakdown characteristics.

  15. Comparison of Space Shuttle Orbiter low-speed static stability and control derivatives obtained from wind-tunnel and approach and landing flight tests

    NASA Technical Reports Server (NTRS)

    Freeman, D. C., Jr.; Spencer, B., Jr.

    1980-01-01

    Tests were conducted in the 8 foot transonic pressure tunnel to obtain wind tunnel data for comparison with static stability and control parameters measured on the space shuttle orbiter approach and landing flight tests. The longitudinal stability, elevon effectiveness, lateral directional stability, and aileron effectiveness derivatives were determined from the wind tunnel data and compared with the flight test results. The comparison covers a range of angles of attack from approximately 2 deg to 10 deg at subsonic Mach numbers of 0.41 to 0.56. In general the wind tunnel results agreed well with the flight test results, indicating the wind tunnel data is applicable to the design of entry vehicles for subsonic speeds over the angle of attack range studied.

  16. FLEXWAL: A computer program for predicting the wall modifications for two-dimensional, solid, adaptive-wall tunnels

    NASA Technical Reports Server (NTRS)

    Everhart, J. L.

    1983-01-01

    A program called FLEXWAL for calculating wall modifications for solid, adaptive-wall wind tunnels is presented. The method used is the iterative technique of NASA TP-2081 and is applicable to subsonic and transonic test conditions. The program usage, program listing, and a sample case are given.

  17. Near-Field Noise Computation for a Subsonic Coannular Jet

    NASA Technical Reports Server (NTRS)

    Loh, Ching Y.; Hultgren, Lennart S.; Jorgenson, Philip C. E.

    2008-01-01

    A high-Reynolds-number, subsonic coannular jet is simulated, using a three-dimensional finite-volume LES method, with emphasis on the near field noise. The nozzle geometry used is the NASA Glenn 3BB baseline model. The numerical results are generally in good agreement with existing experimental findings.

  18. Design procedure for low-drag subsonic airfoils

    NASA Technical Reports Server (NTRS)

    Peterson, J. B.; Chen, A. B.

    1975-01-01

    Airfoil has least amount of drag under given restrictions of boundary layer transition position, lift coefficient, thickness ratio, and Reynolds number based on airfoil chord. It is suitable for use as wing and propeller aircraft sections operating at subsonic speeds and for hydrofoil sections and blades for fans, compressors, turbines, and windmills.

  19. Subsonic annular wing theory with application to flow about nacelles

    NASA Technical Reports Server (NTRS)

    Mann, M. J.

    1974-01-01

    A method has recently been developed for calculating the flow over a subsonic nacelle at zero angle of attack. The method makes use of annular wing theory and boundary-layer theory and has shown good agreement with both experimental data and more complex theoretical solutions. The method permits variation of the mass flow by changing the size of a center body.

  20. Wing-Design Program for Subsonic or Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.; Walkley, K. B.

    1986-01-01

    Surface of mildest possible camber generated. WINGDES provides analysis, design capability and is applicable to both subsonic and supersonic flows. Optimization carried out for entire wing or for designated leading- and trailing-edge areas, for design of missionadaptive surfaces. WINGDES written in FORTRAN IV.

  1. New Model Exhaust System Supports Testing in NASA Lewis' 10- by 10-Foot Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Roeder, James W., Jr.

    1998-01-01

    In early 1996, the ability to run NASA Lewis Research Center's Abe Silverstein 10- by 10- Foot Supersonic Wind Tunnel (10x10) at subsonic test section speeds was reestablished. Taking advantage of this new speed range, a subsonic research test program was scheduled for the 10x10 in the fall of 1996. However, many subsonic aircraft test models require an exhaust source to simulate main engine flow, engine bleed flows, and other phenomena. This was also true of the proposed test model, but at the time the 10x10 did not have a model exhaust capability. So, through an in-house effort over a period of only 5 months, a new model exhaust system was designed, installed, checked out, and made ready in time to support the scheduled test program.

  2. Microspheres for laser velocimetry in high temperature wind tunnel

    NASA Technical Reports Server (NTRS)

    Ghorieshi, Anthony

    1993-01-01

    The introduction of non-intrusive measurement techniques in wind tunnel experimentation has been a turning point in error free data acquisition. Laser velocimetry has been progressively implemented and utilized in various wind tunnels; e.g. subsonic, transonic, and supersonic. The success of the laser velocimeter technique is based on an accurate measurement of scattered light by seeding particles introduced into the flow stream in the wind tunnel. Therefore, application of appropriate seeding particles will affect, to a large extent the acquired data. The seeding material used depends on the type of experiment being run. Among the seeding material for subsonic tunnel are kerosene, Kaolin, and polystyrene. Polystyrene is known to be the best because of being solid particles, having high index of refraction, capable of being made both spherical and monodisperse. However for high temperature wind tunnel testing seeding material must have an additional characteristic that is high melting point. Typically metal oxide powders such as Al2O3 with melting point 3660 F are used. The metal oxides are, however polydispersed, have a high density, and a tendency to form large agglomerate that does not closely follow the flow velocity. The addition of flame phase silica to metal oxide helps to break up the agglomerates, yet still results in a narrow band of polydispersed seeding. The less desirable utility of metal oxide in high temperature wind tunnels necessitates the search for a better alternative particle seeding which this paper addresses. The Laser Velocimetry (LV) characteristic of polystyrene makes it a prime candidate as a base material in achieving the high temperature particle seeding inexpensively. While polystyrene monodisperse seeding particle reported has been successful in a subsonic wind tunnel, it lacks the high melting point and thus is not practically usable in a high temperature wind tunnel. It is well known that rise in melting point of polystyrene can be

  3. Experimental investigations of an 0.0405 scale space shuttle configuration 3 orbiter to determine subsonic stability characteristics (OA21A/OA21B), volume 2

    NASA Technical Reports Server (NTRS)

    Cameron, B. W.; Ritschel, A. J.

    1974-01-01

    Aerodynamic investigations were conducted in a low speed wind tunnel from June 18 through June 25, 1973 on a 0.0405 scale -139B model Space Shuttle Vehicle orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional subsonic aerodynamic characteristics of the proposed PRR Space Shuttle Orbiter. Emphasis was placed on component buildup effects, elevon, rudder, body flaps, rudder flare effectiveness, and canard and speed brake development. Angles of attack from -4 to 24 and angles of sideslip of -10 to 10 were tested. Static pressures were recorded on the base. The aerodynamic force balance results are presented in plotted and tabular form.

  4. Low Reynolds Number Aerodynamic Characteristics of Several Airplane Configurations Designed to Fly in the Mars Atmosphere at Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Re, Richard J.; Pendergraft, Odis C., Jr.; Campbell, Richard L.

    2006-01-01

    A 1/4-scale wind tunnel model of an airplane configuration developed for short duration flight at subsonic speeds in the Martian atmosphere has been tested in the Langley Research Center Transonic Dynamics Tunnel. The tunnel was pumped down to extremely low pressures to represent Martian Mach/Reynolds number conditions. Aerodynamic data were obtained and upper and lower surface wind pressures were measured at one spanwise station on some configurations. Three unswept wings of the same planform but different airfoil sections were tested. Horizontal tail incidence was varied as was the deflection of plain and split trailing-edge flaps. One unswept wing configuration was tested with the lower part of the fuselage removed and the vertical/horizontal tail assembly inverted and mounted from beneath the fuselage. A sweptback wing was also tested. Tests were conducted at Mach numbers from 0.50 to 0.90. Wing chord Reynolds number was varied from 40,000 to 100,000 and angles of attack and sideslip were varied from -10deg to 20deg and -10deg to 10deg, respectively.

  5. Analytical and experimental study of the effects of wing-body aerodynamic interaction on space shuttle subsonic flutter

    NASA Technical Reports Server (NTRS)

    Chipman, R. R.; Rauch, F. J.

    1975-01-01

    The effects on flutter of the aerodynamic interaction between the space shuttle bodies and wing, 1/80th-scale semispan models of the orbiter wing, the complete shuttle and intermediate component combinations were tested in the NASA Langley Research Center 26-inch Transonic Blowdown Wind Tunnel. Using the double lattice method combined with slender body theory to calculate unsteady aerodynamic forces, subsonic flutter speeds were computed for comparison. Using calculated complete vehicle modes, flutter speed trends were computed for the full scale vehicle at an altitude of 15,200 meters and a Mach number of 0.6. Consistent with findings of the model studies, analysis shows the shuttle to have the same flutter speed as an isolated cantilevered wing.

  6. Effect of aileron deflections on the aerodynamic characteristics of a semispan model of a subsonic energy-efficient transport

    NASA Technical Reports Server (NTRS)

    Jacobs, P. F.

    1985-01-01

    An investigation was conducted in the Langley 8 Foot Transonic Pressure Tunnel to determine the effect of aileron deflections on the aerodynamic characteristics of a subsonic energy efficient transport (EET) model. The semispan model had an aspect ratio 10 supercritical wing and was configured with a conventionally located set of ailerons (i.e., a high speed aileron located inboard and a low speed aileron located outboard). Data for the model were taken over a Mach number range from 0.30 to 0.90 and an angle of attack range from approximately -2 deg to 10 deg. The Reynolds number was 2.5 million per foot for Mach number = 0.30 and 4 million per foot for the other Mach numbers. Model force and moment data, aileron effectiveness parameters, aileron hinge moment data, otherwise pressure distributions, and spanwise load data are presented.

  7. Preliminary flight-determined subsonic lift and drag characteristics of the X-29A forward-swept-wing airplane

    NASA Technical Reports Server (NTRS)

    Hicks, John W.; Huckabine, Thomas

    1989-01-01

    The X-29A subsonic lift and drag characteristics determined, met, or exceeded predictions, particularly with respect to the drag polar shapes. Induced drag levels were as great as 20 percent less than wind tunnel estimates, particularly at coefficients of lift above 0.8. Drag polar shape comparisons with other modern fighter aircraft showed the X-29A to have a better overall aircraft aerodynamic Oswald efficiency factor for the same aspect ratio. Two significant problems arose in the data reduction and analysis process. These included uncertainties in angle of attack upwash calibration and effects of maneuver dynamics on drag levels. The latter problem resulted from significantly improper control surface automatic camber control scheduling. Supersonic drag polar results were not obtained during this phase because of a lack of engine instrumentation to measure afterburner fuel flow.

  8. Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D

    SciTech Connect

    Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.

    2016-01-04

    This paper presents results from an explanatory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected in the resulting steady-state analyses using NASA's FUN3D CFD software.

  9. Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D

    NASA Technical Reports Server (NTRS)

    Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.

    2016-01-01

    This paper presents results from an exploratory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected and the resulting steady-state analyses using NASA's FUN3D CFD software.

  10. A study of sound generation in subsonic rotors, volume 2

    NASA Technical Reports Server (NTRS)

    Chalupnik, J. D.; Clark, L. T.

    1975-01-01

    Computer programs were developed for use in the analysis of sound generation by subsonic rotors. Program AIRFOIL computes the spectrum of radiated sound from a single airfoil immersed in a laminar flow field. Program ROTOR extends this to a rotating frame, and provides a model for sound generation in subsonic rotors. The program also computes tone sound generation due to steady state forces on the blades. Program TONE uses a moving source analysis to generate a time series for an array of forces moving in a circular path. The resultant time series are than Fourier transformed to render the results in spectral form. Program SDATA is a standard time series analysis package. It reads in two discrete time series and forms auto and cross covariances and normalizes these to form correlations. The program then transforms the covariances to yield auto and cross power spectra by means of a Fourier transformation.

  11. User's manual: Subsonic/supersonic advanced panel pilot code

    NASA Technical Reports Server (NTRS)

    Moran, J.; Tinoco, E. N.; Johnson, F. T.

    1978-01-01

    Sufficient instructions for running the subsonic/supersonic advanced panel pilot code were developed. This software was developed as a vehicle for numerical experimentation and it should not be construed to represent a finished production program. The pilot code is based on a higher order panel method using linearly varying source and quadratically varying doublet distributions for computing both linearized supersonic and subsonic flow over arbitrary wings and bodies. This user's manual contains complete input and output descriptions. A brief description of the method is given as well as practical instructions for proper configurations modeling. Computed results are also included to demonstrate some of the capabilities of the pilot code. The computer program is written in FORTRAN IV for the SCOPE 3.4.4 operations system of the Ames CDC 7600 computer. The program uses overlay structure and thirteen disk files, and it requires approximately 132000 (Octal) central memory words.

  12. Advanced surface paneling method for subsonic and supersonic flow

    NASA Technical Reports Server (NTRS)

    Erickson, L. L.; Johnson, F. T.; Ehlers, F. E.

    1976-01-01

    Numerical results illustrating the capabilities of an advanced aerodynamic surface paneling method are presented. The method is applicable to both subsonic and supersonic flow, as represented by linearized potential flow theory. The method is based on linearly varying sources and quadratically varying doublets which are distributed over flat or curved panels. These panels are applied to the true surface geometry of arbitrarily shaped three dimensional aerodynamic configurations.

  13. Effect of Riblets upon Flow Separation in a Subsonic Diffuser

    DTIC Science & Technology

    1988-12-01

    Technology Air University In Partial Fulfillment of the Requirements for the Degree of Master of Science in Aeronautical Engineering Nathan W. Martens...afterburner where the flow leaving the turbine must be 4 slowed from a high subsonic Mach number to a Mach number of about 0.2" (8:305). Physicall ., a...Second Edition). New York: McGraw-Hill Book Company, 1975. 3. Cebeci, Tuncer and A. M. 0. Smith. Analysis of Turbulent Boundary Layers. Orlando: Academic

  14. Development of panel methods for subsonic analysis and design

    NASA Technical Reports Server (NTRS)

    Bristow, D. R.

    1980-01-01

    Two computer programs, developed for subsonic inviscid analysis and design are described. The first solves arbitrary mixed analysis design problems for multielement airfoils in two dimensional flow. The second calculates the pressure distribution for arbitrary lifting or nonlifting three dimensional configurations. In each program, inviscid flow is modelled by using distributed source doublet singularities on configuration surface panels. Numerical formulations and representative solutions are presented for the programs.

  15. Design and Development of a Deep Acoustic Lining for the 40-by 80-Foot Wind Tunnel Test Section

    NASA Technical Reports Server (NTRS)

    Soderman, Paul T.; Schmitz, Fredric H.; Allen, Christopher S.; Jaeger, Stephen M.; Sacco, Joe N.; Mosher, Marianne; Hayes, Julie A.

    2002-01-01

    The work described in this report has made effective use of design teams to build a state-of-the-art anechoic wind-tunnel facility. Many potential design solutions were evaluated using engineering analysis, and computational tools. Design alternatives were then evaluated using specially developed testing techniques, Large-scale coupon testing was then performed to develop confidence that the preferred design would meet the acoustic, aerodynamic, and structural objectives of the project. Finally, designs were frozen and the final product was installed in the wind tunnel. The result of this technically ambitious project has been the creation of a unique acoustic wind tunnel. Its large test section (39 ft x 79 ft x SO ft), potentially near-anechoic environment, and medium subsonic speed capability (M = 0.45) will support a full range of aeroacoustic testing-from rotorcraft and other vertical takeoff and landing aircraft to the take-off/landing configurations of both subsonic and supersonic transports.

  16. Subsonic Flow for the Multidimensional Euler-Poisson System

    NASA Astrophysics Data System (ADS)

    Bae, Myoungjean; Duan, Ben; Xie, Chunjing

    2016-04-01

    We establish the existence and stability of subsonic potential flow for the steady Euler-Poisson system in a multidimensional nozzle of a finite length when prescribing the electric potential difference on a non-insulated boundary from a fixed point at the exit, and prescribing the pressure at the exit of the nozzle. The Euler-Poisson system for subsonic potential flow can be reduced to a nonlinear elliptic system of second order. In this paper, we develop a technique to achieve a priori {C^{1,α}} estimates of solutions to a quasi-linear second order elliptic system with mixed boundary conditions in a multidimensional domain enclosed by a Lipschitz continuous boundary. In particular, we discovered a special structure of the Euler-Poisson system which enables us to obtain {C^{1,α}} estimates of the velocity potential and the electric potential functions, and this leads us to establish structural stability of subsonic flows for the Euler-Poisson system under perturbations of various data.

  17. Design Methodology for Multi-Element High-Lift Systems on Subsonic Civil Transport Aircraft

    NASA Technical Reports Server (NTRS)

    Pepper, R. S.; vanDam, C. P.

    1996-01-01

    The choice of a high-lift system is crucial in the preliminary design process of a subsonic civil transport aircraft. Its purpose is to increase the allowable aircraft weight or decrease the aircraft's wing area for a given takeoff and landing performance. However, the implementation of a high-lift system into a design must be done carefully, for it can improve the aerodynamic performance of an aircraft but may also drastically increase the aircraft empty weight. If designed properly, a high-lift system can improve the cost effectiveness of an aircraft by increasing the payload weight for a given takeoff and landing performance. This is why the design methodology for a high-lift system should incorporate aerodynamic performance, weight, and cost. The airframe industry has experienced rapid technological growth in recent years which has led to significant advances in high-lift systems. For this reason many existing design methodologies have become obsolete since they are based on outdated low Reynolds number wind-tunnel data and can no longer accurately predict the aerodynamic characteristics or weight of current multi-element wings. Therefore, a new design methodology has been created that reflects current aerodynamic, weight, and cost data and provides enough flexibility to allow incorporation of new data when it becomes available.

  18. Experimental cavity pressure measurements at subsonic and transonic speeds. Static-pressure results

    NASA Technical Reports Server (NTRS)

    Plentovich, E. B.; Stallings, Robert L., Jr.; Tracy, M. B.

    1993-01-01

    An experimental investigation was conducted to determine cavity flow-characteristics at subsonic and transonic speeds. A rectangular box cavity was tested in the Langley 8-Foot Transonic Pressure Tunnel at Mach numbers from 0.20 to 0.95 at a unit Reynolds number of approximately 3 x 10(exp 6) per foot. The boundary layer approaching the cavity was turbulent. Cavities were tested over a range of length-to-depth ratios (l/h) of 1 to 17.5 for cavity width-to-depth ratios of 1, 4, 8, and 16. Fluctuating- and static-pressure data in the cavity were obtained; however, only static-pressure data is analyzed. The boundaries between the flow regimes based on cavity length-to-depth ratio were determined. The change to transitional flow from open flow occurs at l/h at approximately 6-8 however, the change from transitional- to closed-cavity flow occurred over a wide range of l/h and was dependent on Mach number and cavity configuration. The change from closed to open flow as found to occur gradually. The effect of changing cavity dimensions showed that if the vlaue of l/h was kept fixed but the cavity width was decreased or cavity height was increased, the cavity pressure distribution tended more toward a more closed flow distribution.

  19. Measurements in Vortex Wakes Shed by Conventional and Modified Subsonic Aircraft

    NASA Technical Reports Server (NTRS)

    Rossow, Vernon J.

    1996-01-01

    A theoretical and experimental program is underway at NASA Ames Research Center to first obtain a better understanding of the hazard posed by the vortex wakes of subsonic transports, and then to develop methods on how to modify the wake-generating aircraft in order to make the vortices less hazardous. This paper summarizes results obtained in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center on the characteristics of the vortex wakes that trail from 0.03 scale models of a B-747 and of a DC-10. Measurements are first described that were taken in the wakes with a hot-film anemometer probe, and with wings that range in size from 0.2 to 1.0 times the span of the wake generating models at downstream distances of 81 ft and 162 ft. behind the wake-generating model; i.e., at scale distances of 0.5 and 1.0 mile. The data are then used to evaluate the accuracy of a vortex-lattice method for prediction of the loads induced on following wings by vortex wakes.

  20. Subsonic, transonic, and supersonic stability and control characteristics of the -147B space shuttle orbiter

    NASA Technical Reports Server (NTRS)

    Mennell, R. C.

    1973-01-01

    Experimental aerodynamic investigations were conducted on 0.015 scale representations of two Space Shuttle Orbiter configurations in a trisonic wind tunnel from June 20, 1973 to June 30, 1973. The primary test objective was to define subsonic, transonic, and supersonic stability and control characteristics of the -147B Orbiter. Six-component aerodynamic force and moment data for the -147B Orbiter were recorded over an angle of attack range of -2 deg to 30 deg at Mach numbers of 0.6, 0.9, 1.2, 2.0, and 3.0. Reynolds numbers of 5.0, 7.0, 8.0, and 9.0 x 100000 6/ft were tested at Mach numbers less than 2.0 while testing at Mach 2.0 and 3.0 was conducted at a Reynolds number of 11.0 x 100000/ft. Eleven deflections of 0 deg, +15 deg, -20, deg and -40 deg; body flap deflections of 0 deg, +13.75 deg and -14.25 deg; and rudder flare angles of 24.92 deg and 54.92 deg were tested on the -147B Orbiter over the entire Mach number range. Testing of the -139B Orbiter was for data verification and configuration comparison purposes only.

  1. Aerodynamic Characteristics of a Canard and an Outboard-Tail Airplane Model at High Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Fournier, Paul G.

    1961-01-01

    An investigation has been made in the Langley high-speed 7- by 10-foot tunnel through a range of Mach numbers from 0.60 to 0.95 of the static longitudinal and lateral stability and control characteristics of a canard airplane configuration and an outboard-tail configuration. The canard model had a twisted wing with approximately 67 deg of sweepback and an aspect ratio of 2.91 and was tested with three trapezoidal canard surfaces having ratios of exposed area to wing area of 0.032, 0.076, and 0.121. The canard model had a single body-mounted vertical tail. The outboard-tail model had its horizontal- and vertical-tail surfaces mounted on slender bodies attached to the wing tips and located to the rear and outboard of the 67 deg sweptback wing of aspect ratio 1.00. The data, which are presented with limited analysis, provide information at high subsonic speeds on these two types of high-speed airplanes which have previously been tested at supersonic speeds and reported in NACA RM L58BO7 and NACA RM L58E20.

  2. Dynamic ground effects

    NASA Technical Reports Server (NTRS)

    Paulson, John W., Jr.; Kemmerly, Guy T.; Gilbert, William P.

    1990-01-01

    A research program is underway at the NASA Langley Research Center to study the effect of rate of descent on ground effects. A series of powered models were tested in the Vortex Research Facility under conditions with rate of descent and in the 14 x 22 Foot Subsonic Tunnel under identical conditions but without rate of descent. These results indicate that the rate of descent can have a significant impact on ground effects particularly if vectored or reversed thrust is used.

  3. 8-Foot High Speed Tunnel (HST

    NASA Technical Reports Server (NTRS)

    1957-01-01

    Interior view of the slotted throat test section installed in the 8-Foot High Speed Tunnel (HST) in 1950. The slotted region is about 160 inches in length. In this photograph, the sting-type model support is seen straight on. In a NASA report, the test section is described as follows: 'The test section of the Langley 8-foot transonic tunnel is dodecagonal in cross section and has a cross-sectional area of about 43 square feet. Longitudinal slots are located between each of the 12 wall panels to allow continuous operation through the transonic speed range. The slots contain about 11 percent of the total periphery of the test section. Six of the twelve panels have windows in them to allow for schlieren observations. The entire test section is enclosed in a hemispherical shaped chamber.' John Becker noted that the tunnel's 'final achievement was the development and use in routine operations of the first transonic slotted throat. The investigations of wing-body shapes in this tunnel led to Whitcomb's discovery of the transonic area rule.' James Hansen described the origins of the the slotted throat as follows: 'In 1946 Langley physicist Ray H. Wright conceived a way to do transonic research effectively in a wind tunnel by placing slots in the throat of the test section. The concept for what became known as the slotted-throat or slotted-wall tunnel came to Wright not as a solution to the chronic transonic problem, but as a way to get rid of wall interference (i.e., the mutual effect of two or more meeting waves or vibrations of any kind caused by solid boundaries) at subsonic speeds. For most of the year before Wright came up with this idea, he had been trying to develop a theoretical understanding of wall interference in the 8-Foot HST, which was then being repowered for Mach 1 capability.' When Wright presented these ideas to John Stack, the response was enthusiastic but neither Wright nor Stack thought of slotted-throats as a solution to the transonic problem, only

  4. Tunnel-to-tunnel correlation

    NASA Technical Reports Server (NTRS)

    Steinle, F. W., Jr.

    1982-01-01

    Flow quality is discussed. Incremental comparisons of: (1) the angle of attack, (2) the axial force coefficient, and (3) the base cavity axial force coefficient against the normal force coefficient are presented. Relative blockage determination, relative buoyancy corrections, and boundary layer transition length are discussed. Blockage buoyancy caused by tunnel model wall dynamic interaction is discussed in terms of adaptive walls. The effect of 'transonic turbulence factor' is considered.

  5. Carpal Tunnel Syndrome

    MedlinePlus

    ... Just a passing cramp? It could be carpal tunnel syndrome. The carpal tunnel is a narrow passageway of ligament and bones ... from irritated tendons or other swelling narrows the tunnel and causes the nerve to be compressed. Symptoms ...

  6. Looking into Tunnel Books.

    ERIC Educational Resources Information Center

    Hinshaw, Craig

    1999-01-01

    Describes how to make tunnel books, which are viewed by looking into a "tunnel" created by accordion-folded expanding sides. Suggests possible themes. Describes how to create a walk-through tunnel book for first grade students. (CMK)

  7. A compilation of the pressures measured on a wing and aileron with various amounts of sweep in the Langley 8-foot high-speed tunnel

    NASA Technical Reports Server (NTRS)

    Whitcomb, Richard T

    1948-01-01

    A compilation is made in tabular form of all the pressures measured on a thin high-aspect-ratio wing and aileron with no sweep and with 30 degree and 45 degree of sweepback and sweepforward at high subsonic Mach numbers in the Langley 8-foot high-speed tunnel.

  8. The F2 wind tunnel at Fauga-Mauzac

    NASA Technical Reports Server (NTRS)

    Afchain, D.; Broussaud, P.; Frugier, M.; Rancarani, G.

    1984-01-01

    Details on the French subsonic wind-tunnel F2 that becomes operational on July 1983 are presented. Some of the requirements were: (1) installation of models on any wall of the facility, (2) good observation points due to transparent walls, (3) smooth flow, (4) a laser velocimeter, and (5) easy access and handling. The characteristics include a nonpressurized return circuit, dimensions of 5 x 1.4 x 1.8 m, maximum velocity of 100 m/s and a variable speed fan of 683 kW.

  9. RITD – Wind tunnel testing

    NASA Astrophysics Data System (ADS)

    Haukka, Harri; Harri, Ari-Matti; Aleksashkin, Sergei; Koryanov, Valeri; Schmidt, Walter; Heilimo, Jyri; Finchenko, Valeri; Martynov, Maxim; Ponomarenko, Andrey; Kazakovtsev, Victor; Arruego, Ignazio

    2015-04-01

    An atmospheric re-entry and descent and landing system (EDLS) concept based on inflatable hypersonic decelerator techniques is highly promising for the Earth re-entry missions. We developed such EDLS for the Earth re-entry utilizing a concept that was originally developed for Mars. This EU-funded project is called RITD - Re-entry: Inflatable Technology Development - and it was to assess the bene¬fits of this technology when deploying small payloads from low Earth orbits to the surface of the Earth with modest costs. The principal goal was to assess and develope a preliminary EDLS design for the entire relevant range of aerodynamic regimes expected to be encountered in Earth's atmosphere during entry, descent and landing. The RITD entry and descent system utilizes an inflatable hypersonic decelerator. Development of such system requires a combination of wind tunnel tests and numerical simulations. This included wind tunnel tests both in transsonic and subsonic regimes. The principal aim of the wind tunnel tests was the determination of the RITD damping factors in the Earth atmosphere and recalculation of the results for the case of the vehicle descent in the Mars atmosphere. The RITD mock-up model used in the tests was in scale of 1:15 of the real-size vehicle as the dimensions were (midsection) diameter of 74.2 mm and length of 42 mm. For wind tunnel testing purposes the frontal part of the mock-up model body was manufactured by using a PolyJet 3D printing technology based on the light curing of liquid resin. The tail part of the mock-up model body was manufactured of M1 grade copper. The structure of the mock-up model placed th center of gravity in the same position as that of the real-size RITD. The wind tunnel test program included the defining of the damping factor at seven values of Mach numbers 0.85; 0.95; 1.10; 1.20; 1.25; 1.30 and 1.55 with the angle of attack ranging from 0 degree to 40 degrees with the step of 5 degrees. The damping characteristics of

  10. Characteristics of the Langley 8-foot Transonic Tunnel with Slotted Test Section

    NASA Technical Reports Server (NTRS)

    Wright, Ray H; Ritchie, Virgil S; Pearson, Albin O

    1958-01-01

    A large wind tunnel, approximately 8 feet in diameter, has been converted to transonic operation by means of slots in the boundary extending in the direction of flow. The usefulness of such a slotted wind tunnel, already known with respect to the reduction of the subsonic blockage interference and the production of continuously variable supersonic flows, has been augmented by devising a slot shape with which a supersonic test region with excellent flow quality could be produced. Experimental locations of detached shock waves ahead of axially symmetric bodies at low supersonic speeds in the slotted test section agreed satisfactorily with predictions obtained by use of existing approximate methods.

  11. NASA Subsonic Rotary Wing Project - Structures and Materials Discipline

    NASA Technical Reports Server (NTRS)

    Halbig, Michael C.; Johnson, Susan M.

    2008-01-01

    The Structures & Materials Discipline within the NASA Subsonic Rotary Wing Project is focused on developing rotorcraft technologies. The technologies being developed are within the task areas of: 5.1.1 Life Prediction Methods for Engine Structures & Components 5.1.2 Erosion Resistant Coatings for Improved Turbine Blade Life 5.2.1 Crashworthiness 5.2.2 Methods for Prediction of Fatigue Damage & Self Healing 5.3.1 Propulsion High Temperature Materials 5.3.2 Lightweight Structures and Noise Integration The presentation will discuss rotorcraft specific technical challenges and needs as well as details of the work being conducted in the six task areas.

  12. Subsonic/transonic stall flutter investigation of a rotating rig

    NASA Technical Reports Server (NTRS)

    Jutras, R. R.; Fost, R. B.; Chi, R. M.; Beacher, B. F.

    1981-01-01

    Stall flutter is investigated by obtaining detailed quantitative steady and aerodynamic and aeromechanical measurements in a typical fan rotor. The experimental investigation is made with a 31.3 percent scale model of the Quiet Engine Program Fan C rotor system. Both subsonic/transonic (torsional mode) flutter and supersonic (flexural) flutter are investigated. Extensive steady and unsteady data on the blade deformations and aerodynamic properties surrounding the rotor are acquired while operating in both the steady and flutter modes. Analysis of this data shows that while there may be more than one traveling wave present during flutter, they are all forward traveling waves.

  13. Computation of subsonic base flow on a vector processor

    NASA Technical Reports Server (NTRS)

    Rudy, D. H.

    1987-01-01

    Two-dimensional subsonic laminar compressible base flow has been studied using numerical solutions of the time-dependent Navier-Stokes equations. These solutions were obtained using an explicit finite-difference scheme which is highly efficient on a vector processor. The organization of the code for a CDC CYBER-205 computer is described. Solutions were obtained for Mach 0.4 and 0.6 flows past a slender blunt-based model at moderately high Reynolds numbers. The flow in the wake is unsteady with periodic shedding of vortices from the trailing edge. The computed shedding frequency was found to increase with increasing Reynolds number.

  14. Sound radiation from a subsonic rotor subjected to turbulence

    NASA Technical Reports Server (NTRS)

    Sevik, M.

    1974-01-01

    The broadband sound radiated by a subsonic rotor subjected to turbulence in the approach stream has been analyzed. The power spectral density of the sound intensity has been found to depend on a characteristic time scale-namely, the integral scale of the turbulence divided by the axial flow velocity-as well as several length-scale ratios. These consist of the ratio of the integral scale to the acoustic wavelength, rotor radius, and blade chord. Due to the simplified model chosen, only a limited number of cascade parameters appear. Limited comparisons with experimental data indicate good agreement with predicted values.

  15. Solution of the unsteady subsonic thin airfoil problem

    NASA Technical Reports Server (NTRS)

    Williams, M. H.

    1982-01-01

    The problem of a thin airfoil subject to simple harmonic disturbances in a uniform subsonic free stream is solved by extension of a technique developed earlier for a stationary strip vibrating in a uniform fluid. Explicit expressions are given for the lift and moment, acoustic directivity pattern, and total acoustic power for arbitrary upwash and, in particular, for the 'elementary disturbances': plunge, pitch and a stationary transverse gust. Numerical results for a simple skewed gust are presented and compared to the high-frequency asymptotic theory of Martinez and Widnall.

  16. Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Slater, J. W.; Saunders, J. D.

    2015-01-01

    Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

  17. Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Slater, John; Saunders, John

    2014-01-01

    Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

  18. Evaluation of flow quality in two large NASA wind tunnels at transonic speeds

    NASA Technical Reports Server (NTRS)

    Harvey, W. D.; Stainback, P. C.; Owen, F. K.

    1980-01-01

    Wind tunnel testing of low drag airfoils and basic transition studies at transonic speeds are designed to provide high quality aerodynamic data at high Reynolds numbers. This requires that the flow quality in facilities used for such research be excellent. To obtain a better understanding of the characteristics of facility disturbances and identification of their sources for possible facility modification, detailed flow quality measurements were made in two prospective NASA wind tunnels. Experimental results are presented of an extensive and systematic flow quality study of the settling chamber, test section, and diffuser in the Langley 8 foot transonic pressure tunnel and the Ames 12 foot pressure wind tunnel. Results indicate that the free stream velocity and pressure fluctuation levels in both facilities are low at subsonic speeds and are so high as to make it difficult to conduct meaningful boundary layer control and transition studies at transonic speeds.

  19. Wind tunnel wall effects in a linear oscillating cascade

    NASA Technical Reports Server (NTRS)

    Buffum, Daniel H.; Fleeter, Sanford

    1991-01-01

    Experiments in a linear oscillating cascade reveal that the wind tunnel walls enclosing the airfoils have, in some cases, a detrimental effect on the oscillating cascade aerodynamics. In a subsonic flow field, biconvex airfoils are driven simultaneously in harmonic, torsion-mode oscillations for a range of interblade phase angle values. It is found that the cascade dynamic periodicity - the airfoil to airfoil variation in unsteady surface pressure - is good for some values of interblade phase angle but poor for others. Correlation of the unsteady pressure data with oscillating flat plate cascade predictions is generally good for conditions where the periodicity is good and poor where the periodicity is poor. Calculations based upon linearized unsteady aerodynamic theory indicate that pressure waves reflected from the wind tunnel walls are responsible for the cases where there is poor periodicity and poor correlation with the predictions.

  20. NASA Noise Reduction Program for Advanced Subsonic Transports

    NASA Technical Reports Server (NTRS)

    Stephens, David G.; Cazier, F. W., Jr.

    1995-01-01

    Aircraft noise is an important byproduct of the world's air transportation system. Because of growing public interest and sensitivity to noise, noise reduction technology is becoming increasingly important to the unconstrained growth and utilization of the air transportation system. Unless noise technology keeps pace with public demands, noise restrictions at the international, national and/or local levels may unduly constrain the growth and capacity of the system to serve the public. In recognition of the importance of noise technology to the future of air transportation as well as the viability and competitiveness of the aircraft that operate within the system, NASA, the FAA and the industry have developed noise reduction technology programs having application to virtually all classes of subsonic and supersonic aircraft envisioned to operate far into the 21st century. The purpose of this paper is to describe the scope and focus of the Advanced Subsonic Technology Noise Reduction program with emphasis on the advanced technologies that form the foundation of the program.

  1. Control of Subsonic and Sonic Jets with Limiting Tabs

    NASA Astrophysics Data System (ADS)

    Berrueta, T.; Rathakrishnan, E.

    2017-04-01

    Aerodynamic mixing of subsonic and sonic jets with limiting tabs, with and without corrugations, has been studied experimentally. Limiting tab located at the nozzle exit and at a downstream distance of 0.5D has been considered in this study. Mixing caused by the tab at nozzle exit is found to be better that of tab at 0.5D, for both plain and corrugated geometries. Also, both tabs caused better mixing for underexpanded sonic jets than the correctly expanded sonic jet and subsonic jets. At nozzle pressure ratio 3 the plain tab at the nozzle exit reduced the core by about 56 % and the corrugated tab by about 51 %. But when the plain tab is placed at 0.5D the jet mixing is retarded. However, the corrugated tab at 0.5D enhances the mixing, though not up to the level of the same tab at 0D, at all Mach numbers except 0.6. The maximum reduction of core caused by shifted corrugated tab is 14 % for Mach 0.8 jet.

  2. Subsonic and Supersonic Effects in Bose-Einstein Condensate

    NASA Technical Reports Server (NTRS)

    Zak, Michail

    2003-01-01

    A paper presents a theoretical investigation of subsonic and supersonic effects in a Bose-Einstein condensate (BEC). The BEC is represented by a time-dependent, nonlinear Schroedinger equation that includes terms for an external confining potential term and a weak interatomic repulsive potential proportional to the number density of atoms. From this model are derived Madelung equations, which relate the quantum phase with the number density, and which are used to represent excitations propagating through the BEC. These equations are shown to be analogous to the classical equations of flow of an inviscid, compressible fluid characterized by a speed of sound (g/Po)1/2, where g is the coefficient of the repulsive potential and Po is the unperturbed mass density of the BEC. The equations are used to study the effects of a region of perturbation moving through the BEC. The excitations created by a perturbation moving at subsonic speed are found to be described by a Laplace equation and to propagate at infinite speed. For a supersonically moving perturbation, the excitations are found to be described by a wave equation and to propagate at finite speed inside a Mach cone.

  3. Subsonic Round and Rectangular Twin Jet Flow Effects

    NASA Technical Reports Server (NTRS)

    Bozak, Rick; Wernet, Mark

    2014-01-01

    Subsonic and supersonic aircraft concepts proposed by NASAs Fundamental Aeronautics Program have integrated propulsion systems with asymmetric nozzles. The asymmetry in the exhaust of these propulsion systems creates asymmetric flow and acoustic fields. The flow asymmetries investigated in the current study are from two parallel round, 2:1, and 8:1 aspect ratio rectangular jets at the same nozzle conditions. The flow field was measured with streamwise and cross-stream particle image velocimetry (PIV). A large dataset of single and twin jet flow field measurements was acquired at subsonic jet conditions. The effects of twin jet spacing and forward flight were investigated. For round, 2:1, and 8:1 rectangular twin jets at their closest spacings, turbulence levels between the two jets decreased due to enhanced jet mixing at near static conditions. When the flight Mach number was increased to 0.25, the flow around the twin jet model created a velocity deficit between the two nozzles. This velocity deficit diminished the effect of forward flight causing an increase in turbulent kinetic energy relative to a single jet. Both of these twin jet flow field effects decreased with increasing twin jet spacing relative to a single jet. These variations in turbulent kinetic energy correlate with changes in far-field sound pressure level.

  4. Advanced Configurations for Very Large Subsonic Transport Airplanes

    NASA Technical Reports Server (NTRS)

    McMasters, John H.; Paisley, David J.; Hubert, Richard J.; Kroo, Ilan; Bofah, Kwasi K.; Sullivan, John P.; Drela, Mark

    1996-01-01

    Recent aerospace industry interest in developing a subsonic commercial transport airplane with 50 percent greater passenger capacity than the largest existing aircraft in this category (the Boeing 747-400 with approximately 400-450 seats) has generated a range of proposals based largely on the configuration paradigm established nearly 50 years ago with the Boeing B-47 bomber. While this basic configuration paradigm has come to dominate subsonic commercial airplane development since the advent of the Boeing 707/Douglas DC-8 in the mid-1950's, its extrapolation to the size required to carry more than 600-700 passengers raises several questions. To explore these and a number of related issues, a team of Boeing, university, and NASA engineers was formed under the auspices of the NASA Advanced Concepts Program. The results of a Research Analysis focused on a large, unconventional transport airplane configuration for which Boeing has applied for a patent are the subject of this report. It should be noted here that this study has been conducted independently of the Boeing New Large Airplane (NLA) program, and with the exception of some generic analysis tools which may be common to this effort and the NLA (as will be described later), no explicit Boeing NLA data other than that published in the open literature has been used in the conduct of the study reported here.

  5. Extension of sonic anemometry to high subsonic Mach number flows

    NASA Astrophysics Data System (ADS)

    Otero, R.; Lowe, K. T.; Ng, W. F.

    2017-03-01

    In the literature, the application of sonic anemometry has been limited to low subsonic Mach number, near-incompressible flow conditions. To the best of the authors’ knowledge, this paper represents the first time a sonic anemometry approach has been used to characterize flow velocity beyond Mach 0.3. Using a high speed jet, flow velocity was measured using a modified sonic anemometry technique in flow conditions up to Mach 0.83. A numerical study was conducted to identify the effects of microphone placement on the accuracy of the measured velocity. Based on estimated error strictly due to uncertainty in time-of-acoustic flight, a random error of +/- 4 m s‑1 was identified for the configuration used in this experiment. Comparison with measurements from a Pitot probe indicated a velocity RMS error of +/- 9 m s‑1. The discrepancy in error is attributed to a systematic error which may be calibrated out in future work. Overall, the experimental results from this preliminary study support the use of acoustics for high subsonic flow characterization.

  6. Steady subsonic flow around finite-thickness wings

    NASA Technical Reports Server (NTRS)

    Kuo, C. C.; Morino, L.

    1975-01-01

    The general method for analyzing steady subsonic potential aerodynamic flow around a lifting body having arbitrary shape is presented. By using the Green function method, an integral representation for the potential is obtained. Under small perturbation assumption, the potential at any point, P, in the field depends only upon the values of the potential and its normal derivative on the surface of the body. Hence if the point P approaches the surface of the body, the representation reduces to an integral equation relating the potential and its normal derivative (which is known from the boundary conditions) on the surface. The question of uniqueness is examined and it is shown that, for thin wings, the operator becomes singular as the thickness approaches zero. This fact may yield numerical problems for very thin wings. However, numerical results obtained for a rectangular wing in subsonic flow show that these problems do not appear even for thickness ratio tau = .001. Comparison with existing results shows that the proposed method is at least as fast and accurate as the lifting surface theories.

  7. 14 CFR 91.881 - Final compliance: Civil subsonic jet airplanes weighing 75,000 pounds or less.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 2 2014-01-01 2014-01-01 false Final compliance: Civil subsonic jet... OPERATING AND FLIGHT RULES Operating Noise Limits § 91.881 Final compliance: Civil subsonic jet airplanes... operate to or from an airport in the contiguous United States a civil subsonic jet airplane subject...

  8. An experimental study of the flow field surrounding a subsonic jet in a cross flow. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Dennis, Robert Foster

    1993-01-01

    An experimental investigation of the flow interaction of a 5.08 cm (2.00 in.) diameter round subsonic jet exhausting perpendicularly to a flat plate in a subsonic cross flow was conducted in the NASA Ames 7x1O ft. Wind Tunnel Number One. Flat plate surface pressures were measured at 400 locations in a 30.48 cm (12.0 in.) concentric circular array surrounding the jet exit. Results from these measurements are provided in tabular and graphical form for jet-to-crossflow velocity ratios ranging from 4 to 12, and for jet exit Mach numbers ranging from 0.50 to 0.93. Laser doppler velocimeter (LDV) three component velocity measurements were made in selected regions in the developed jet plume and near the flat plate surface, at a jet Mach number of 0.50 and jet-to-crossflow velocity ratios of 6 and 8. The results of both pressure and LDV measurements are compared with the results of previous experiments. In addition, pictures of the jet plume shape at jet velocity ratios ranging from 4 to 12 were obtained using schleiren photography. The LDV measurements are consistent with previous work, but more extensive measurements will be necessary to provide a detailed picture of the flow field. The surface pressure results compare closely with previous work and provide a useful characterization of jet induced surface pressures. The results demonstrate the primary influence of jet velocity ratio and the secondary influence of jet Mach number in determining such surface pressures.

  9. In-flight pressure distributions and skin-friction measurements on a subsonic transport high-lift wing section

    NASA Technical Reports Server (NTRS)

    Yip, Long P.; Vijgen, Paul M. H. W.; Hardin, Jay D.; Vandam, C. P.

    1993-01-01

    Flight experiments are being conducted as part of a multiphased subsonic transport high-lift research program for correlation with wind-tunnel and computational results. The NASA Langley Transport Systems Research Vehicle (B737-100 aircraft) is used to obtain in-flight flow characteristics at full-scale Reynolds numbers to contribute to the understanding of 3-D high-lift, multi-element flows including attachment-line transition and relaminarization, confluent boundary-layer development, and flow separation characteristics. Flight test results of pressure distributions and skin friction measurements were obtained for a full-chord wing section including the slat, main-wing, and triple-slotted, Fowler flap elements. Test conditions included a range of flap deflections, chord Reynolds numbers (10 to 21 million), and Mach numbers (0.16 to 0.40). Pressure distributions were obtained at 144 chordwise locations of a wing section (53-percent wing span) using thin pressure belts over the slat, main-wing, and flap elements. Flow characteristics observed in the chordwise pressure distributions included leading-edge regions of high subsonic flows, leading-edge attachment-line locations, slat and main-wing cove-flow separation and reattachment, and trailing-edge flap separation. In addition to the pressure distributions, limited skin-friction measurements were made using Preston-tube probes. Preston-tube measurements on the slat upper surface suggested relaminarization of the turbulent flow introduced by the pressure belt on the slat leading-edge surface when the slat attachment line was laminar. Computational analysis of the in-flight pressure measurements using two-dimensional, viscous multielement methods modified with simple-sweep theory showed reasonable agreement. However, overprediction of the pressures on the flap elements suggests a need for better detailed measurements and improved modeling of confluent boundary layers as well as inclusion of three-dimensional viscous

  10. Cryogenic wind tunnels. II

    NASA Technical Reports Server (NTRS)

    Kilgore, Robert A.

    1987-01-01

    The application of the cryogenic concept to various types of tunnels including Ludwieg tube tunnel, Evans clean tunnel, blowdown, induced-flow, and continuous-flow fan-driven tunnels is discussed. Benefits related to construction and operating costs are covered, along with benefits related to new testing capabilities. It is noted that cooling the test gas to very low temperatures increases Reynolds number by more than a factor of seven. From the energy standpoint, ambient-temperature fan-driven closed-return tunnels are considered to be the most efficient type of tunnel, while a large reduction in the required tunnel stagnation pressure can be achieved through cryogenic operation. Operating envelopes for three modes of operation for a cryogenic transonic pressure tunnel with a 2.5 by 2.5 test section are outlined. A computer program for calculating flow parameters and power requirements for wind tunnels with operating temperatures from saturation to above ambient is highlighted.

  11. An Impact-Location Estimation Algorithm for Subsonic Uninhabited Aircraft

    NASA Technical Reports Server (NTRS)

    Bauer, Jeffrey E.; Teets, Edward

    1997-01-01

    An impact-location estimation algorithm is being used at the NASA Dryden Flight Research Center to support range safety for uninhabited aerial vehicle flight tests. The algorithm computes an impact location based on the descent rate, mass, and altitude of the vehicle and current wind information. The predicted impact location is continuously displayed on the range safety officer's moving map display so that the flightpath of the vehicle can be routed to avoid ground assets if the flight must be terminated. The algorithm easily adapts to different vehicle termination techniques and has been shown to be accurate to the extent required to support range safety for subsonic uninhabited aerial vehicles. This paper describes how the algorithm functions, how the algorithm is used at NASA Dryden, and how various termination techniques are handled by the algorithm. Other approaches to predicting the impact location and the reasons why they were not selected for real-time implementation are also discussed.

  12. Review of Propulsion Technologies for N+3 Subsonic Vehicle Concepts

    NASA Technical Reports Server (NTRS)

    Ashcraft, Scott W.; Padron, Andres S.; Pascioni, Kyle A.; Stout, Gary W., Jr.; Huff, Dennis L.

    2011-01-01

    NASA has set aggressive fuel burn, noise, and emission reduction goals for a new generation (N+3) of aircraft targeting concepts that could be viable in the 2035 timeframe. Several N+3 concepts have been formulated, where the term "N+3" indicate aircraft three generations later than current state-of-the-art aircraft, "N". Dramatic improvements need to be made in the airframe, propulsion systems, mission design, and the air transportation system in order to meet these N+3 goals. The propulsion system is a key element to achieving these goals due to its major role with reducing emissions, fuel burn, and noise. This report provides an in-depth description and assessment of propulsion systems and technologies considered in the N+3 subsonic vehicle concepts. Recommendations for technologies that merit further research and development are presented based upon their impact on the N+3 goals and likelihood of being operational by 2035.

  13. Unsteady Subsonic and Transonic Potential Flow over Helicopter Rotor Blades

    NASA Technical Reports Server (NTRS)

    Isom, M. P.

    1974-01-01

    Differential equations and boundary conditions for a rotor blade in forward flight, with subsonic or transonic tip Mach number, are derived. A variety of limiting flow regimes determined by different limits involving blade thickness ratio, aspect ratio, advance ratio and maximum tip Mach number is discussed. The transonic problem is discussed in some detail, and in particular the conditions that make this problem quasi-steady or essentially unsteady are determined. Asymptotic forms of equations and boundary conditions that are valid in an appropriately scaled region of the tip and an azimuthal sector on the advancing side are derived. The equations are then put in a form that is valid from the blade tip inboard through the strip theory region.

  14. Prediction of subsonic vortex shedding from forebodies with chines

    NASA Technical Reports Server (NTRS)

    Mendenhall, Michael R.; Lesieutre, Daniel J.

    1990-01-01

    An engineering prediction method and associated computer code VTXCHN to predict nose vortex shedding from circular and noncircular forebodies with sharp chine edges in subsonic flow at angles of attack and roll are presented. Axisymmetric bodies are represented by point sources and doublets, and noncircular cross sections are transformed to a circle by either analytical or numerical conformal transformations. The lee side vortex wake is modeled by discrete vortices in crossflow planes along the body; thus the three-dimensional steady flow problem is reduced to a two-dimensional, unsteady, separated flow problem for solution. Comparison of measured and predicted surface pressure distributions, flow field surveys, and aerodynamic characteristics are presented for noncircular bodies alone and forebodies with sharp chines.

  15. Analysis of an advanced technology subsonic turbofan incorporating revolutionary materials

    NASA Technical Reports Server (NTRS)

    Knip, Gerald, Jr.

    1987-01-01

    Successful implementation of revolutionary composite materials in an advanced turbofan offers the possibility of further improvements in engine performance and thrust-to-weight ratio relative to current metallic materials. The present analysis determines the approximate engine cycle and configuration for an early 21st century subsonic turbofan incorporating all composite materials. The advanced engine is evaluated relative to a current technology baseline engine in terms of its potential fuel savings for an intercontinental quadjet having a design range of 5500 nmi and a payload of 500 passengers. The resultant near optimum, uncooled, two-spool, advanced engine has an overall pressure ratio of 87, a bypass ratio of 18, a geared fan, and a turbine rotor inlet temperature of 3085 R. Improvements result in a 33-percent fuel saving for the specified misssion. Various advanced composite materials are used throughout the engine. For example, advanced polymer composite materials are used for the fan and the low pressure compressor (LPC).

  16. Persistent pressure anisotropy in the subsonic magnetosheath region

    NASA Technical Reports Server (NTRS)

    Crooker, N. U.; Siscoe, G. L.; Geller, R. B.

    1976-01-01

    Observations of proton fluxes in the subsonic flow region of the magnetosheath show double peaks as a function of angle in the equatorial plane of the spacecraft. The peaks are separated by a wide angle, usually more than 90 deg. Suggestions of double flux peaks are present in 95 per cent of the data. They are interpreted as the effect of a persistent pressure anisotropy. The clearest cases were analyzed to determine the orientation of the flux peaks relative to the magnetic field and direction of a model hydrodynamic flow. The peaks are shown to be consistent with a greater pressure perpendicular to the field. Possible sources of the pressure anisotropy in the magnetosheath are discussed.

  17. Inviscid and Viscous Interactions in Subsonic Corner Flows

    PubMed Central

    Chung, Kung-Ming; Chang, Po-Hsiung; Chang, Keh-Chin

    2013-01-01

    A flap can be used as a high-lift device, in which a downward deflection results in a gain in lift at a given geometric angle of attack. To characterize the aerodynamic performance of a deflected surface in compressible flows, the present study examines a naturally developed turbulent boundary layer past the convex and concave corners. This investigation involves the analysis of mean and fluctuating pressure distributions. The results obtained indicate strong inviscid-viscous interactions. There are upstream expansion and downstream compression for the convex-corner flows, while the opposite trend is observed for the concave-corner flows. A combined flow similarity parameter, based on the small perturbation theory, is proposed to scale the flow characteristics in both subsonic convex- and concave-corner flows. PMID:23935440

  18. Evaluation of the Advanced Subsonic Technology Program Noise Reduction Benefits

    NASA Technical Reports Server (NTRS)

    Golub, Robert A.; Rawls, John W., Jr.; Russell, James W.

    2005-01-01

    This report presents a detailed evaluation of the aircraft noise reduction technology concepts developed during the course of the NASA/FAA Advanced Subsonic Technology (AST) Noise Reduction Program. In 1992, NASA and the FAA initiated a cosponsored, multi-year program with the U.S. aircraft industry focused on achieving significant advances in aircraft noise reduction. The program achieved success through a systematic development and validation of noise reduction technology. Using the NASA Aircraft Noise Prediction Program, the noise reduction benefit of the technologies that reached a NASA technology readiness level of 5 or 6 were applied to each of four classes of aircraft which included a large four engine aircraft, a large twin engine aircraft, a small twin engine aircraft and a business jet. Total aircraft noise reductions resulting from the implementation of the appropriate technologies for each class of aircraft are presented and compared to the AST program goals.

  19. Subsonic Wing Optimization for Handling Qualities Using ACSYNT

    NASA Technical Reports Server (NTRS)

    Soban, Danielle Suzanne

    1996-01-01

    The capability to accurately and rapidly predict aircraft stability derivatives using one comprehensive analysis tool has been created. The PREDAVOR tool has the following capabilities: rapid estimation of stability derivatives using a vortex lattice method, calculation of a longitudinal handling qualities metric, and inherent methodology to optimize a given aircraft configuration for longitudinal handling qualities, including an intuitive graphical interface. The PREDAVOR tool may be applied to both subsonic and supersonic designs, as well as conventional and unconventional, symmetric and asymmetric configurations. The workstation-based tool uses as its model a three-dimensional model of the configuration generated using a computer aided design (CAD) package. The PREDAVOR tool was applied to a Lear Jet Model 23 and the North American XB-70 Valkyrie.

  20. 4. 'Ring Stones & Tunnel Sections, Tunnel #33,' Southern Pacific ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    4. 'Ring Stones & Tunnel Sections, Tunnel #33,' Southern Pacific Standard Double-Track Tunnel, ca. 1913. Compare to photos in documentation sets for Tunnel 18 (HAER No. CA-197), Tunnel 34 (HAER No. CA-206), and Tunnel 1 (HAER No. CA-207). - Central Pacific Transcontinental Railroad, Sacramento to Nevada state line, Sacramento, Sacramento County, CA

  1. Aeroelastic characteristics of a cascade of mistuned blades in subsonic and supersonic flows. [turbofan engines

    NASA Technical Reports Server (NTRS)

    Kielb, R. E.; Kaza, K. R. V.

    1981-01-01

    The effects of mistuning on flutter and forced response of a cascade in subsonic in subsonic and supersonic flow were investigated. The aerodynamic and structural coupling between the bending and torsional motions and the aerodynamic coupling between the blades were studied. It is shown that frequency mistuning always has a beneficial effect on flutter. For the cascade considered, the potential for raising flutter speed is greater in subsonic than in supersonic flow. Preliminary results for structural damping mistuning show that there are no additional benefits over adding damping mistuning may have either a beneficial or an adverse effect on forced response, depending on the engine order of the excitation and Mach number.

  2. Carpal tunnel repair - slideshow

    MedlinePlus

    ... page: //medlineplus.gov/ency/presentations/100078.htm Carpal tunnel repair - series—Normal anatomy To use the sharing ... in the wrist and the wrist bones (carpal tunnel). Review Date 5/9/2015 Updated by: C. ...

  3. Aerodynamic Assessment of Flight-Determined Subsonic Lift and Drag Characteristics of Seven Lifting-Body and Wing-Body Reentry Vehicle Configurations

    NASA Technical Reports Server (NTRS)

    Saltzman, Edwin J.; Wang, K. Charles; Iliff, Kenneth W.

    2002-01-01

    This report examines subsonic flight-measured lift and drag characteristics of seven lifting-body and wing-body reentry vehicle configurations with truncated bases. The seven vehicles are the full-scale M2-F1, M2-F2, HL-10, X-24A, X-24B, and X-15 vehicles and the Space Shuttle Enterprise. Subsonic flight lift and drag data of the various vehicles are assembled under aerodynamic performance parameters and presented in several analytical and graphical formats. These formats are intended to unify the data and allow a greater understanding than individually studying the vehicles allows. Lift-curve slope data are studied with respect to aspect ratio and related to generic wind-tunnel model data and to theory for low-aspect-ratio platforms. The definition of reference area is critical for understanding and comparing the lift data. The drag components studied include minimum drag coefficient, lift-related drag, maximum lift-to drag ratio, and, where available, base pressure coefficients. The influence of forebody drag on afterbody and base drag at low lift is shown to be related to Hoerner's compilation for body, airfoil, nacelle, and canopy drag. This feature may result in a reduced need of surface smoothness for vehicles with a large ratio of base area to wetted area. These analyses are intended to provide a useful analytical framework with which to compare and evaluate new vehicle configurations of the same generic family.

  4. HSR Model Deformation Measurements from Subsonic to Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Burner, A. W.; Erickson, G. E.; Goodman, W. L.; Fleming, G. A.

    1999-01-01

    This paper describes the video model deformation technique (VMD) used at five NASA facilities and the projection moire interferometry (PMI) technique used at two NASA facilities. Comparisons between the two techniques for model deformation measurements are provided. Facilities at NASA-Ames and NASA-Langley where deformation measurements have been made are presented. Examples of HSR model deformation measurements from the Langley Unitary Wind Tunnel, Langley 16-foot Transonic Wind Tunnel, and the Ames 12-foot Pressure Tunnel are presented. A study to improve and develop new targeting schemes at the National Transonic Facility is also described. The consideration of milled targets for future HSR models is recommended when deformation measurements are expected to be required. Finally, future development work for VMD and PMI is addressed.

  5. Variable Density Tunnel

    NASA Technical Reports Server (NTRS)

    1931-01-01

    Variable Density Tunnel in operation. Man at far right is probably Harold J. 'Cannonball' Tuner, longtime safety officer, who started with Curtiss in the teens. This view of the Variable Density Tunnel clearly shows the layout of the Tunnel's surroundings, as well as the plumbing and power needs of the this innovative research tool.

  6. Development of an intelligent videogrammetric wind tunnel measurement system

    NASA Astrophysics Data System (ADS)

    Graves, Sharon S.; Burner, Alpheus W.

    2001-11-01

    A videogrammetric technique developed at NASA Langley Research Center has been used at five NASA facilities at the Langley and Ames Research Centers for deformation measurements on a number of sting mounted and semispan models. These include high-speed research and transport models tested over a wide range of aerodynamic conditions including subsonic, transonic, and supersonic regimes. The technique, based on digital photogrammetry, has been used to measure model attitude, deformation, and sting bending. In addition, the technique has been used to study model injection rate effects and to calibrate and validate methods for predicting static aeroelastic deformations of wind tunnel models. An effort is currently underway to develop an intelligent videogrammetric measurement system that will be both useful and usable in large production wind tunnels while providing accurate data in a robust and timely manner. Designed to encode a higher degree of knowledge through computer vision, the system features advanced pattern recognition techniques to improve automated location and identification of targets placed on the wind tunnel model to be used for aerodynamic measurements such as attitude and deformation. This paper will describe the development and strategy of the new intelligent system that was used in a recent test at a large transonic wind tunnel.

  7. Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Pinier, Jeremy T.; Niskey, Charles J.; Hanke, Jeremy L.; Tomek, William G.

    2011-01-01

    Throughout three full design analysis cycles, the Ares I project within the Constellation program has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley s Unitary Plan Wind Tunnel (UPWT) at the overlapping condition ofMach=1.6. Much insight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

  8. Development of an Intelligent Videogrammetric Wind Tunnel Measurement System

    NASA Technical Reports Server (NTRS)

    Graves, Sharon S.; Burner, Alpheus W.

    2004-01-01

    A videogrammetric technique developed at NASA Langley Research Center has been used at five NASA facilities at the Langley and Ames Research Centers for deformation measurements on a number of sting mounted and semispan models. These include high-speed research and transport models tested over a wide range of aerodynamic conditions including subsonic, transonic, and supersonic regimes. The technique, based on digital photogrammetry, has been used to measure model attitude, deformation, and sting bending. In addition, the technique has been used to study model injection rate effects and to calibrate and validate methods for predicting static aeroelastic deformations of wind tunnel models. An effort is currently underway to develop an intelligent videogrammetric measurement system that will be both useful and usable in large production wind tunnels while providing accurate data in a robust and timely manner. Designed to encode a higher degree of knowledge through computer vision, the system features advanced pattern recognition techniques to improve automated location and identification of targets placed on the wind tunnel model to be used for aerodynamic measurements such as attitude and deformation. This paper will describe the development and strategy of the new intelligent system that was used in a recent test at a large transonic wind tunnel.

  9. Parametric Study of the Flutter Stability of a Semi-Rigid 3-D Wing-With-Engine Nacelle Model in Subsonic Flow

    NASA Astrophysics Data System (ADS)

    Försching, H.; Knaack, J. M.

    1993-08-01

    A parametric investigation is performed of the aeroelastic flutter stability behaviour of a semi-rigid 3-D wing-with-engine nacelle model in subsonic flow. The system under investigation is a wind tunnel model that was flutter tested some years ago. It consists of a swept-back half wing with a pylon-mounted engine nacelle, representative of modern large transport aircraft, and is elastically restrained at its wing root, so that it may execute decoupled (rigid body) rolling and pitching oscillations about two orthogonal axes. For this binary aeroelastic system, first the equations of motion and then the aeroelastic stability equations are set up in terms of generalized coordinates. In addition to the basic wind tunnel model configuration, two artificial configurations with other positions of the rotation axes and corresponding mode shapes are investigated. For the computation of the motion-induced generalized airloads, a panel technique is used for both the wing and the engine nacelle that is replaced by an annular wing. Numerical results are presented for several systematic parameter variations and Mach numbers, where special emphasis is placed on the effects of the motion-induced unsteady airloads acting on the engine nacelle, the position of the rotation axes, and the frequency ratio of the two modes in roll and pitch. Moreover, a comparison is made with some wind tunnel test results.

  10. Experimental investigations of an 0.0405 scale Space Shuttle Configuration 3 orbiter to determine subsonic stability characteristics. Volume 1: OA21A

    NASA Technical Reports Server (NTRS)

    Cameron, B. W.; Ritschel, A. J.

    1973-01-01

    Experimental aerodynamic investigations were conducted in a low speed wind tunnel from May 21 through June 4 and from June 18 through June 25, 1973 on a 0.0405 scale -139B model Space Shuttle Vehicle (SSV) orbiter. The purpose of the test was to investigate the longitudinal and lateral-directional subsonic aerodynamic characteristics of the proposed PRR Space Shuttle orbiter. Emphasis was placed on component buildup effects, elevon, rudder, body flaps, rudder flare effectiveness, and canard and speed brake development. Angles of attack from -4 deg. to 24 deg. and angles of sideslip of -10 deg. to 10 deg. were tested. Static pressures were recorded on the base. The aerodynamic force balance results are presented in plotted and tabular form.

  11. Low-Subsonic-Speed Static Longitudinal Stability and Control Characteristics of a Winged Reentry-Vehicle Configuration Having Wingtip Panels that Fold up for High-Drag Reentry

    NASA Technical Reports Server (NTRS)

    Ware, George M.

    1960-01-01

    An investigation of the low-subsonic-speed static longitudinal stability and control characteristics of a model of a manned reentry-vehicle configuration capable of high-drag reentry and glide landing has been a made in the Langley free-flight tunnel. The model had a modified 63 deg delta plan-form wing with a fuselage on the upper surface. This configuration had wingtip panels designed to fold up 90 deg for the high-drag reentry phase of the flight and to extend horizontally for the glide landing. Data for the basic configurations and modifications to determine the effects of plan form, wingtip panel incidence, dihedral, and vertical position of the wingtip panels are presented without analysis.

  12. A user's guide to the Langley 16-foot transonic tunnel complex. Revision 1

    NASA Technical Reports Server (NTRS)

    1990-01-01

    The operational characteristics and equipment associated with the Langley 16-foot transonic tunnel complex which is located in buildings 1146 and 1234 at the Langley Research Center are described in detail. This complex consists of the 16-foot transonic wind tunnel, the static test facility, and the 16- by 24-inch water tunnel research facilities. The 16-foot transonic tunnel is a single-return atmospheric wind tunnel with a 15.5 foot diameter test section and a Mach number capability from 0.20 to 1.30. The emphasis for research conducted in this research complex is on the integration of the propulsion system into advanced aircraft concepts. In the past, the primary focus has been on the integration of nozzles and empennage into the afterbody of fighter aircraft. During the last several years this experimental research has been expanded to include developing the fundamental data base necessary to verify new theoretical concepts, inlet integration into fighter aircraft, nozzle integration for supersonic and hypersonic transports, nacelle/pylon/wing integration for subsonic transport configurations, and the study of vortical flows (in the 16- by 24-inch water tunnel). The purpose here is to provide a comprehensive description of the operational characteristics of the research facilities of the 16-foot transonic tunnel complex and their associated systems and equipments.

  13. Validation of US3D for Capsule Aerodynamics using 05-CA Wind Tunnel Test Data

    NASA Technical Reports Server (NTRS)

    Schwing, Alan

    2012-01-01

    Several comparisons of computational fluid dynamics to wind tunnel test data are shown for the purpose of code validation. The wind tunnel test, 05-CA, uses a 7.66% model of NASA's Multi-Purpose Crew Vehicle in the 11-foot test section of the Ames Unitary Plan Wind tunnel. A variety of freestream conditions over four Mach numbers and three angles of attack are considered. Test data comparisons include time-averaged integrated forces and moments, time-averaged static pressure ports on the surface, and Strouhal Number. The applicability of the US3D code to subsonic and transonic flow over a bluff body is assessed on a comprehensive data set. With close comparison, this work validates US3D for highly separated flows similar to those examined here.

  14. High speed wind tunnel tests of the PTA aircraft. [Propfan Test Assessment Program

    NASA Technical Reports Server (NTRS)

    Aljabri, A. S.; Little, B. H., Jr.

    1986-01-01

    Propfans, advanced highly-loaded propellers, are proposed to power transport aircraft that cruise at high subsonic speeds, giving significant fuel savings over the equivalent turbofan-powered aircraft. NASA is currently sponsoring the Propfan Test Assessment Program (PTA) to provide basic data on the structural integrity and acoustic performance of the propfan. The program involves installation design, wind-tunnel tests, and flight tests of the Hamilton Standard SR-7 propfan in a wing-mount tractor installation on the Gulfstream II aircraft. This paper reports on the high-speed wind-tunnel tests and presents the computational aerodynamic methods that were employed in the analyses, design, and evaluation of the configuration. In spite of the complexity of the configuration, these methods provide aerodynamic predictions which are in excellent agreement with wind-tunnel data.

  15. Quasispherical subsonic accretion in X-ray pulsars

    NASA Astrophysics Data System (ADS)

    Shakura, Nikolai I.; Postnov, Konstantin A.; Kochetkova, A. Yu; Hjalmarsdotter, L.

    2013-04-01

    A theoretical model is considered for quasispherical subsonic accretion onto slowly rotating magnetized neutron stars. In this regime, the accreting matter settles down subsonically onto the rotating magnetosphere, forming an extended quasistatic shell. Angular momentum transfer in the shell occurs via large-scale convective motions resulting, for observed pulsars, in an almost iso-angular-momentum \\omega \\sim 1/R^2 rotation law inside the shell. The accretion rate through the shell is determined by the ability of the plasma to enter the magnetosphere due to Rayleigh-Taylor instabilities, with allowance for cooling. A settling accretion regime is possible for moderate accretion rates \\dot M \\lesssim \\dot M_* \\simeq 4\\times 10^{16} g s ^{-1}. At higher accretion rates, a free-fall gap above the neutron star magnetosphere appears due to rapid Compton cooling, and the accretion becomes highly nonstationary. Observations of spin-up/spin-down rates of quasispherically wind accreting equilibrium X-ray pulsars with known orbital periods (e.g., GX 301-2 and Vela X-1) enable us to determine the main dimensionless parameters of the model, as well as to estimate surface magnetic field of the neutron star. For equilibrium pulsars, the independent measurements of the neutron star magnetic field allow for an estimate of the stellar wind velocity of the optical companion without using complicated spectroscopic measurements. For nonequilibrium pulsars, a maximum value is shown to exist for the spin-down rate of the accreting neutron star. From observations of the spin-down rate and the X-ray luminosity in such pulsars (e.g., GX 1+4, SXP 1062, and 4U 2206+54), a lower limit can be put on the neutron star magnetic field, which in all cases turns out to be close to the standard value and which agrees with cyclotron line measurements. Furthermore, both explains the spin-up/spin-down of the pulsar frequency on large time-scales and also accounts for the irregular short

  16. Subsonic Aircraft With Regression and Neural-Network Approximators Designed

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Hopkins, Dale A.

    2004-01-01

    At the NASA Glenn Research Center, NASA Langley Research Center's Flight Optimization System (FLOPS) and the design optimization testbed COMETBOARDS with regression and neural-network-analysis approximators have been coupled to obtain a preliminary aircraft design methodology. For a subsonic aircraft, the optimal design, that is the airframe-engine combination, is obtained by the simulation. The aircraft is powered by two high-bypass-ratio engines with a nominal thrust of about 35,000 lbf. It is to carry 150 passengers at a cruise speed of Mach 0.8 over a range of 3000 n mi and to operate on a 6000-ft runway. The aircraft design utilized a neural network and a regression-approximations-based analysis tool, along with a multioptimizer cascade algorithm that uses sequential linear programming, sequential quadratic programming, the method of feasible directions, and then sequential quadratic programming again. Optimal aircraft weight versus the number of design iterations is shown. The central processing unit (CPU) time to solution is given. It is shown that the regression-method-based analyzer exhibited a smoother convergence pattern than the FLOPS code. The optimum weight obtained by the approximation technique and the FLOPS code differed by 1.3 percent. Prediction by the approximation technique exhibited no error for the aircraft wing area and turbine entry temperature, whereas it was within 2 percent for most other parameters. Cascade strategy was required by FLOPS as well as the approximators. The regression method had a tendency to hug the data points, whereas the neural network exhibited a propensity to follow a mean path. The performance of the neural network and regression methods was considered adequate. It was at about the same level for small, standard, and large models with redundancy ratios (defined as the number of input-output pairs to the number of unknown coefficients) of 14, 28, and 57, respectively. In an SGI octane workstation (Silicon Graphics

  17. Airfoil modification effects on subsonic and transonic pressure distributions and performance for the EA-6B airplane

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Sewall, William G.

    1995-01-01

    Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

  18. Transonic single-mode flutter and buffet of a low aspect ratio wing having a subsonic airfoil shape

    NASA Technical Reports Server (NTRS)

    Erickson, L. L.

    1974-01-01

    Transonic flutter and buffet results obtained from wind-tunnel tests of a low aspect ratio semispan wing model are presented. The tests were conducted to investigate potential transonic aeroelastic problems of vehicles having subsonic airfoil sections. The model employed NACA 00XX-64 airfoil sections in the streamwise direction and had a 14 deg leading edge sweep angle. Aspect ratio, and average thickness were 4.0, 0.35, and 8 percent, respectively. The model was tested at Mach numbers from 0.6 to 0.95 at angles of attack from 0 deg to 15 deg. Two zero lift flutter conditions were found that involved essentially single normal mode vibrations. With boundary layer trips on the model, flutter occurred in a narrow Mach number range centered at about Mach 0.90. The frequency and motion of this flutter were like that of the first normal mode vibration. With the trips removed flutter occurred at a slightly high Mach number but in a mode strongly resembling that of the second normal mode.

  19. Analysis of a high speed civil transport configuration at subsonic flow conditions using a Navier-Stokes solver

    NASA Technical Reports Server (NTRS)

    Lessard, Victor R.

    1993-01-01

    Computations of three dimensional vortical flows over a generic High Speed Civil Transport (HSCT) configuration with an aspect ratio of 3.04 are performed using a thin-layer Navier-Stokes solver. The HSCT cruise configuration is modeled without leading or trailing edge flap deflections and without engine nacelles. The flow conditions, which correspond to tests done in the NASA Langley 8-Foot Transonic Pressure Tunnel (TPT), are a subsonic Mach number of 0.3 and Reynolds number of 4.4 million for a range-of-attack (-.23 deg to 17.78 deg). The effects of the farfield boundary location with respect to the body are investigated. The boundary layer is assumed turbulent and simulated using an algebraic turbulence model. The key features of the vortices and their interactions are captured. Grid distribution in the vortex regions is critical for predicting the correct induced lift. Computed forces and surface pressures compare reasonably well with the experimental TPT data.

  20. Effects of Boundary-Layer Thickness on Unsteady Flow Characteristics Inside Open Cavities at Subsonic and Transonic Speeds

    NASA Astrophysics Data System (ADS)

    Yang, Dang-Guo; Li, Jian-Qiang; Fan, Zhao-Lin; Luo, Xin-Fu

    An experimental study was conducted in a 0.6m by 0.6m wind-tunnel to analyze effects of boundary-layer thickness on unsteady flow characteristics inside a rectangular open cavity at subsonic and transonic speeds. The sound pressure level (SPL) distributions at the centerline of the cavity floor and Sound pressure frequency spectrum (SPFS) characteristics on some measurement positions presented herein was obtained with cavity length-to-depth ratio (L/D) of 8 over Mach numbers (Ma) of 0.6 and 1.2 at a Reynolds numbers (Re) of 1.23 × 107 and 2.02 × 107 per meter under different boundary-layer thickness to cavity-depth ratios (δ/D). The experimental angle of attack, yawing and rolling angles were 0°. The results indicate that decrease in δ/D leads to severe flow separation and unsteady pressure fluctuation, which induces increase in SPL at same measurement points inside the cavity at Ma of 0.6. At Ma of 1.2, decrease in δ/D results in enhancing compressible waves. Generally, decrease in δ/D induces more flow self-sustained oscillation frequencies. It also makes severer aerodynamic noise inside the open cavity.

  1. Propulsion System for Very High Altitude Subsonic Unmanned Aircraft

    NASA Technical Reports Server (NTRS)

    Bents, David J.; Mockler, Ted; Maldonado, Jaime; Harp, James L., Jr.; King, Joseph F.; Schmitz, Paul C.

    1998-01-01

    This paper explains why a spark ignited gasoline engine, intake pressurized with three cascaded stages of turbocharging, was selected to power NASA's contemplated next generation of high altitude atmospheric science aircraft. Beginning with the most urgent science needs (the atmospheric sampling mission) and tracing through the mission requirements which dictate the unique flight regime in which this aircraft has to operate (subsonic flight at greater then 80 kft) we briefly explore the physical problems and constraints, the available technology options and the cost drivers associated with developing a viable propulsion system for this highly specialized aircraft. The paper presents the two available options (the turbojet and the turbocharged spark ignited engine) which are discussed and compared in the context of the flight regime. We then show how the unique nature of the sampling mission, coupled with the economic considerations pursuant to aero engine development, point to the spark ignited engine as the only cost effective solution available. Surprisingly, this solution compares favorably with the turbojet in the flight regime of interest. Finally, some remarks are made about NASA's present state of development, and future plans to flight demonstrate the three stage turbocharged powerplant.

  2. CFD validation of subsonic turbulent planar shear layers

    NASA Technical Reports Server (NTRS)

    Lai, H. T.; Raju, M. S.

    1993-01-01

    The primary objective of the present study is to assess the limitations and capabilities of RPLUS in predicting the entrainment, mixing, and burning characteristics of a high subsonic planar shear layer; these computations are performed in conjunction with an experiment being conducted at NASA Lewis Research Center. Turbulence is modeled by a two-equation k-epsilon closure modified for compressibility effects. Combustion is modeled by two different models: one in which a finite-rate laminar chemistry model is used for H2-O2 oxidation, and the other based on a composition joint pdf approach to account for the turbulence-chemistry interaction effects. Between upwind and centered differencing schemes, minimal differences are observed for the flow properties of a nonreacting shear layer case studied. In comparison with experimental data, computed growth rates are underpredicted in both reacting and nonreacting cases of a shear layer, but more severely for the reacting flows. However, the qualitative agreement between computation and experiment is reasonably good for the other observable characteristics, and indicative of potential for further improvements.

  3. The 1996 Subsonic Aircraft: Contrail and Cloud Effects Special Study

    NASA Technical Reports Server (NTRS)

    Toon, Owen B.; Condon, Estelle P. (Technical Monitor)

    1996-01-01

    During April 1996 NASA, in conjunction with the DOE, will sponsor a multi-aircraft field campaign to better understand the microphysical and radiative properties of cirrus clouds, the origins of ice nuclei and cloud condensation nuclei in the upper troposphere, and the possible role that the commercial subsonic aircraft fleet might play in altering cloud or aerosol properties. The NASA ER-2 aircraft will be used as a remote sensing platform, while the NASA DC-8 aircraft will be used as an in situ measurement platform. In situ observations will include a full set of size distribution measurements from nano-meter to millimeter sizes, ice water content measurements, gas phase and condensed phase chemical measurements, ice crystal optical phase function measurements, lidar observations of cloud top and cloud base, and atmospheric state measurement. The ER-2 will have lidar, microwave ice water path measurements, as well as visible and infrared spectral measurement. In this presentation the highlights of the mission will be presented. The goal will be to address fundamental questions such as the mode of nucleation of cirrus clouds, the composition of the nuclei on which cirrus form, the degree to which aircraft impact cirrus cloud properties.

  4. The Impact of Subsonic Twin Jets on Airport Noise

    NASA Technical Reports Server (NTRS)

    Bozak, Richard F.

    2012-01-01

    Subsonic and supersonic aircraft concepts proposed through NASA s Fundamental Aeronautics Program have multiple engines mounted near one another. Engine configurations with multiple jets introduce an asymmetry to the azimuthal directivity of the jet noise. Current system noise predictions add the jet noise from each jet incoherently, therefore, twin jets are estimated by adding 3 EPNdB to the far-field noise radiated from a single jet. Twin jet effects have the ability to increase or decrease the radiated noise to different azimuthal observation locations. Experiments have shown that twin jet effects are reduced with forward flight and increasing spacings. The current experiment investigates the impact of spacing, and flight effects on airport noise for twin jets. Estimating the jet noise radiated from twin jets as that of a single jet plus 3 EPNdB may be sufficient for horizontal twin jets with an s/d of 4.4 and 5.5, where s is the center-to-center spacing and d is the jet diameter. However, up to a 3 EPNdB error could be present for jet spacings with an s/d of 2.6 and 3.2.

  5. CFD validation of subsonic turbulent planar shear layers

    NASA Astrophysics Data System (ADS)

    Lai, H. T.; Raju, M. S.

    1993-06-01

    The primary objective of the present study is to assess the limitations and capabilities of RPLUS in predicting the entrainment, mixing, and burning characteristics of a high subsonic planar shear layer; these computations are performed in conjunction with an experiment being conducted at NASA Lewis Research Center. Turbulence is modeled by a two-equation k-epsilon closure modified for compressibility effects. Combustion is modeled by two different models: one in which a finite-rate laminar chemistry model is used for H2-O2 oxidation, and the other based on a composition joint pdf approach to account for the turbulence-chemistry interaction effects. Between upwind and centered differencing schemes, minimal differences are observed for the flow properties of a nonreacting shear layer case studied. In comparison with experimental data, computed growth rates are underpredicted in both reacting and nonreacting cases of a shear layer, but more severely for the reacting flows. However, the qualitative agreement between computation and experiment is reasonably good for the other observable characteristics, and indicative of potential for further improvements.

  6. Transverse Injection into Subsonic Crossflow with Various Injector Orifice Geometries

    NASA Technical Reports Server (NTRS)

    Foster, Lancert E.; Zaman, Khairul B.

    2010-01-01

    Computational and experimental results are presented for a case study of single injectors employed in 90 deg transverse injection into a non-reacting subsonic flow. Different injector orifice shapes are used (including circular, square, diamond-shaped, and wide rectangular slot), all with constant cross-sectional area, to observe the effects of this variation on injector penetration and mixing. Whereas the circle, square, and diamond injector produce similar jet plumes, the wide rectangular slot produces a plume with less vertical penetration than the others. There is also some evidence that the diamond injector produces slightly faster penetration with less mixing of the injected fluid. In addition, a variety of rectangular injectors were analyzed, with varying length/width ratios. Both experimental and computational data show improved plume penetration with increased streamwise orifice length. 3-D Reynolds-Averaged Navier-Stokes (RANS) results are obtained for the various injector geometries using NCC (National Combustion Code) with the kappa-epsilon turbulence model in multi-species modes on an unstructured grid. Grid sensitivity results are also presented which indicate consistent qualitative trends in the injector performance comparisons with increasing grid refinement.

  7. High frequency flow-structural interaction in dense subsonic fluids

    NASA Technical Reports Server (NTRS)

    Liu, Baw-Lin; Ofarrell, J. M.

    1995-01-01

    Prediction of the detailed dynamic behavior in rocket propellant feed systems and engines and other such high-energy fluid systems requires precise analysis to assure structural performance. Designs sometimes require placement of bluff bodies in a flow passage. Additionally, there are flexibilities in ducts, liners, and piping systems. A design handbook and interactive data base have been developed for assessing flow/structural interactions to be used as a tool in design and development, to evaluate applicable geometries before problems develop, or to eliminate or minimize problems with existing hardware. This is a compilation of analytical/empirical data and techniques to evaluate detailed dynamic characteristics of both the fluid and structures. These techniques have direct applicability to rocket engine internal flow passages, hot gas drive systems, and vehicle propellant feed systems. Organization of the handbook is by basic geometries for estimating Strouhal numbers, added mass effects, mode shapes for various end constraints, critical onset flow conditions, and possible structural response amplitudes. Emphasis is on dense fluids and high structural loading potential for fatigue at low subsonic flow speeds where high-frequency excitations are possible. Avoidance and corrective measure illustrations are presented together with analytical curve fits for predictions compiled from a comprehensive data base.

  8. Coherent structures and wavepackets in subsonic transitional turbulent jets

    NASA Astrophysics Data System (ADS)

    Yang, Haihua; Zhang, Xingchen; Ran, Lingke; Sun, Dejun; Wan, Zhenhua

    2017-02-01

    A large eddy simulation (LES) is performed for two subsonic jets with a Reynolds number of Re=10^5, which have different core temperatures, i.e., the cold and hot jet. The far-field overall sound pressure levels (OASPL) and noise spectra are well validated against previous experimental results. It is found that the OASPL is raised by heating at shallow angles. The most energetic coherent structures are extracted with specified frequencies using the filter based on the frequency domain variant of the snapshot method of proper orthogonal decomposition (POD). The m=0,1 modes have high coherence of near-field pressure for both jets, while the coherence of m=0 modes is enhanced greatly by heating. Based on the coherent structures, spatial wavepackets are educed and the characteristics of growth, saturation and decay are analyzed and compared between the two jets in detail. The results show that heating would enhance the linear growth rate for high frequency components, and nonlinear growth rates for low frequency components in general, which are responsible for higher OASPL in the hot jet. The far-field sound generated by wavepackets is computed using the Kirchhoff extrapolation, which matches well with that of LES at shallow angles. This indicates that the wavepackets associated with coherent structures are dominant sound sources in forced transitional turbulent jets. Additionally, the present POD method is proven to be a robust tool to extract the salient features of the wavepackets in turbulent flows.

  9. An Overview of NASA's Subsonic Research Aircraft Testbed (SCRAT)

    NASA Technical Reports Server (NTRS)

    Baumann, Ethan; Hernandez, Joe; Ruhf, John C.

    2013-01-01

    National Aeronautics and Space Administration Dryden Flight Research Center acquired a Gulfstream III (GIII) aircraft to serve as a testbed for aeronautics flight research experiments. The aircraft is referred to as SCRAT, which stands for SubsoniC Research Aircraft Testbed. The aircraft's mission is to perform aeronautics research; more specifically raising the Technology Readiness Level (TRL) of advanced technologies through flight demonstrations and gathering high-quality research data suitable for verifying the technologies, and validating design and analysis tools. The SCRAT has the ability to conduct a range of flight research experiments throughout a transport class aircraft's flight envelope. Experiments ranging from flight-testing of a new aircraft system or sensor to those requiring structural and aerodynamic modifications to the aircraft can be accomplished. The aircraft has been modified to include an instrumentation system and sensors necessary to conduct flight research experiments along with a telemetry capability. An instrumentation power distribution system was installed to accommodate the instrumentation system and future experiments. An engineering simulation of the SCRAT has been developed to aid in integrating research experiments. A series of baseline aircraft characterization flights has been flown that gathered flight data to aid in developing and integrating future research experiments. This paper describes the SCRAT's research systems and capabilities.

  10. Computational Study of Separating Flow in a Planar Subsonic Diffuser

    NASA Technical Reports Server (NTRS)

    DalBello, Teryn; Dippold, Vance, III; Georgiadis, Nicholas J.

    2005-01-01

    A computational study of the separated flow through a 2-D asymmetric subsonic diffuser has been performed. The Wind Computational Fluid Dynamics code is used to predict the separation and reattachment behavior for an incompressible diffuser flow. The diffuser inlet flow is a two-dimensional, turbulent, and fully-developed channel flow with a Reynolds number of 20,000 based on the centerline velocity and the channel height. Wind solutions computed with the Menter SST, Chien k-epsilon, Spalart-Allmaras and Explicit Algebraic Reynolds Stress turbulence models are compared with experimentally measured velocity profiles and skin friction along the upper and lower walls. In addition to the turbulence model study, the effects of grid resolution and use of wall functions were investigated. The grid studies varied the number of grid points across the diffuser and varied the initial wall spacing from y(sup +) = 0.2 to 60. The wall function study assessed the applicability of wall functions for analysis of separated flow. The SST and Explicit Algebraic Stress models provide the best agreement with experimental data, and it is recommended wall functions should only be used with a high level of caution.

  11. Chaos control for the plates subjected to subsonic flow

    NASA Astrophysics Data System (ADS)

    Norouzi, Hamed; Younesian, Davood

    2016-07-01

    The suppression of chaotic motion in viscoelastic plates driven by external subsonic air flow is studied. Nonlinear oscillation of the plate is modeled by the von-Kármán plate theory. The fluid-solid interaction is taken into account. Galerkin's approach is employed to transform the partial differential equations of the system into the time domain. The corresponding homoclinic orbits of the unperturbed Hamiltonian system are obtained. In order to study the chaotic behavior of the plate, Melnikov's integral is analytically applied and the threshold of the excitation amplitude and frequency for the occurrence of chaos is presented. It is found that adding a parametric perturbation to the system in terms of an excitation with the same frequency of the external force can lead to eliminate chaos. Variations of the Lyapunov exponent and bifurcation diagrams are provided to analyze the chaotic and periodic responses. Two perturbation-based control strategies are proposed. In the first scenario, the amplitude of control forces reads a constant value that should be precisely determined. In the second strategy, this amplitude can be proportional to the deflection of the plate. The performance of each controller is investigated and it is found that the second scenario would be more efficient.

  12. Aerodynamics in ground effect and predicted landing ground roll of a fighter configuration with a secondary-nozzle thrust reverser

    NASA Technical Reports Server (NTRS)

    Banks, Daniel W.

    1988-01-01

    An experimental investigation of the in-ground effect aerodynamic characteristics and predicted landing-ground-roll performance of wing-canard fighter configuration with a secondary nozzle thrust reverser was completed. These tests were conducted in the Langley 14 by 22 foot Subsonic Wind Tunnel using a model equipped with a pneumatic jet for thrust simulation of nozzle pressure ratios up to 4.0. The model was tested in the landing rollout configuration at approx. wheel touchdown height for a range of decreasing dynamic pressure from 50 psf down to 10 psf. Landing-ground-roll predictions of the configuration were calculated using the wind tunnel results.

  13. The cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Kilgore, R. A.

    1976-01-01

    Based on theoretical studies and experience with a low speed cryogenic tunnel and with a 1/3-meter transonic cryogenic tunnel, the cryogenic wind tunnel concept was shown to offer many advantages with respect to the attainment of full scale Reynolds number at reasonable levels of dynamic pressure in a ground based facility. The unique modes of operation available in a pressurized cryogenic tunnel make possible for the first time the separation of Mach number, Reynolds number, and aeroelastic effects. By reducing the drive-power requirements to a level where a conventional fan drive system may be used, the cryogenic concept makes possible a tunnel with high productivity and run times sufficiently long to allow for all types of tests at reduced capital costs and, for equal amounts of testing, reduced total energy consumption in comparison with other tunnel concepts.

  14. Simulator of Road Tunnel

    NASA Astrophysics Data System (ADS)

    Danišovič, Peter; Schlosser, František; Šrámek, Juraj; Rázga, Martin

    2015-05-01

    A Tunnel Traffic & Operation Simulator is a device of the Centre of Transport Research at the University of Žilina. The Simulator allows managing technological equipment of virtual two-tube highway tunnel, which is interconnected with simulation of vehicle traffic in tunnel. Changes of the traffic-operation states and other equipment are reflecting at the simulated traffic, as well as simulations of various emergency events in traffic initiate changes in tunnel detecting and measuring devices. It is thus possible to simulate emergency states, which can be affected by various faults of technology as well as by climatic conditions. The solutions can be found in irreplaceable experiences of Slovak road tunnel operators, changes of trafficoperation states, visualizations of operator technological display screens, technological devices labelling in order to increase operational safety of road tunnels.

  15. Pipetron Tunnel Construction Issues

    SciTech Connect

    Friant, James E.; Bauer, Robert A.; Gross, David L.; May, Michael; Lach, Joseph

    1997-01-01

    This report examines issues involved in the civil construction aspects of the tunneling that could be done in the region of Fermilab to support the Pipetron along, moderately deep, tunnel loop. Cost, technical and political aspects of tunneling are addressed in this preliminary guide for further study. At Snowmass 96, in a series of informal, but comprehensive discussions, several guidelines were developed to frame this report.

  16. Charge Islands Through Tunneling

    NASA Technical Reports Server (NTRS)

    Robinson, Daryl C.

    2002-01-01

    It has been recently reported that the electrical charge in a semiconductive carbon nanotube is not evenly distributed, but rather it is divided into charge "islands." This paper links the aforementioned phenomenon to tunneling and provides further insight into the higher rate of tunneling processes, which makes tunneling devices attractive. This paper also provides a basis for calculating the charge profile over the length of the tube so that nanoscale devices' conductive properties may be fully exploited.

  17. Tunnel closure calculations

    SciTech Connect

    Moran, B.; Attia, A.

    1995-07-01

    When a deeply penetrating munition explodes above the roof of a tunnel, the amount of rubble that falls inside the tunnel is primarily a function of three parameters: first the cube-root scaled distance from the center of the explosive to the roof of the tunnel. Second the material properties of the rock around the tunnel, and in particular the shear strength of that rock, its RQD (Rock Quality Designator), and the extent and orientation of joints. And third the ratio of the tunnel diameter to the standoff distance (distance between the center of explosive and the tunnel roof). The authors have used CALE, a well-established 2-D hydrodynamic computer code, to calculate the amount of rubble that falls inside a tunnel as a function of standoff distance for two different tunnel diameters. In particular they calculated three of the tunnel collapse experiments conducted in an iron ore mine near Kirkeness, Norway in the summer of 1994. The failure model that they used in their calculations combines an equivalent plastic strain criterion with a maximum tensile strength criterion and can be calibrated for different rocks using cratering data as well as laboratory experiments. These calculations are intended to test and improve the understanding of both the Norway Experiments and the ACE (Array of conventional Explosive) phenomenology.

  18. Cryogenic wind tunnel technology. A way to measurement at higher Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Beck, J. W.

    1984-01-01

    The goals, design, problems, and value of cryogenic transonic wind tunnels being developed in Europe are discussed. The disadvantages inherent in low-Reynolds-number (Re) wind tunnel simulations of aircraft flight at high Re are reviewed, and the cryogenic tunnel is shown to be the most practical method to achieve high Re. The design proposed for the European Transonic Wind tunnel (ETW) is presented: parameters include cross section. DISPLAY 83A46484/2 = 4 sq m, operating pressure = 5 bar, temperature = 110-120 K, maximum Re = 40 x 10 to the 6th, liquid N2 consumption = 40,000 metric tons/year, and power = 39,5 MW. The smaller Cologne subsonic tunnel being adapted to cryogenic use for preliminary studies is described. Problems of configuration, materials, and liquid N2 evaporation and handling and the research underway to solve them are outlined. The benefits to be gained by the construction of these costly installations are seen more in applied aerodynamics than in basic research in fluid physics. The need for parallel development of both high Re tunnels and computers capable of performing high-Re numerical analysis is stressed.

  19. Modernization and Activation of the NASA Ames 11- by 11-Foot Transonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Kmak, Frank J.

    2000-01-01

    The Unitary Plan Wind Tunnel (UPWT) was modernized to improve performance, capability, productivity, and reliability. Automation systems were installed in all three UPWT tunnel legs and the Auxiliaries facility. Major improvements were made to the four control rooms, model support systems, main drive motors, and main drive speed control. Pressure vessel repairs and refurbishment to the electrical distribution system were also completed. Significant changes were made to improve test section flow quality in the 11-by 11-Foot Transonic leg. After the completion of the construction phase of the project, acceptance and checkout testing was performed to demonstrate the capabilities of the modernized facility. A pneumatic test of the tunnel circuit was performed to verify the structural integrity of the pressure vessel before wind-on operations. Test section turbulence, flow angularity, and acoustic parameters were measured throughout the tunnel envelope to determine the effects of the tunnel flow quality improvements. The new control system processes were thoroughly checked during wind-off and wind-on operations. Manual subsystem modes and automated supervisory modes of tunnel operation were validated. The aerodynamic and structural performance of both the new composite compressor rotor blades and the old aluminum rotor blades was measured. The entire subsonic and supersonic envelope of the 11-by 11-Foot Transonic leg was defined up to the maximum total pressure.

  20. 16-foot transonic tunnel test section flowfield survey

    NASA Technical Reports Server (NTRS)

    Yetter, J. A.; Abeyounis, W. K.

    1994-01-01

    A flow survey has been made of the test section of the NASA Langley Research Center 16-Foot Transonic Tunnel at subsonic and supersonic speeds. The survey was performed using five five-hole pyramid-head probes mounted at 14 inch intervals on a survey rake. Probes were calibrated at freestream Mach numbers from 0.50 to 0.95 and from 1.18 to 1.23. Flowfield surveys were made at Mach numbers from 0.50 to 0.90 and at Mach 1.20. The surveys were made at tunnel stations 130.6, 133.6, and 136.0. By rotating the survey rake through 180 degrees, a cylindrical volume of the test section 4.7 feet in diameter and 5.4 feet long centered about the tunnel centerline was surveyed. Survey results showing the measured test section upflow and sideflow characteristics and local Mach number distributions are presented. The report documents the survey probe calibration techniques used, summarizes the procedural problems encountered during testing, and identifies the data discrepancies observed during the post-test data analysis.

  1. Variable Stiffness Spar Wind-Tunnel Model Development and Testing

    NASA Technical Reports Server (NTRS)

    Florance, James R.; Heeg, Jennifer; Spain, Charles V.; Ivanco, Thomas G.; Wieseman, Carol D.; Lively, Peter S.

    2004-01-01

    The concept of exploiting wing flexibility to improve aerodynamic performance was investigated in the wind tunnel by employing multiple control surfaces and by varying wing structural stiffness via a Variable Stiffness Spar (VSS) mechanism. High design loads compromised the VSS effectiveness because the aerodynamic wind-tunnel model was much stiffer than desired in order to meet the strength requirements. Results from tests of the model include stiffness and modal data, model deformation data, aerodynamic loads, static control surface derivatives, and fuselage standoff pressure data. Effects of the VSS on the stiffness and modal characteristics, lift curve slope, and control surface effectiveness are discussed. The VSS had the most effect on the rolling moment generated by the leading-edge outboard flap at subsonic speeds. The effects of the VSS for the other control surfaces and speed regimes were less. The difficulties encountered and the ability of the VSS to alter the aeroelastic characteristics of the wing emphasize the need for the development of improved design and construction methods for static aeroelastic models. The data collected and presented is valuable in terms of understanding static aeroelastic wind-tunnel model development.

  2. Computational Investigations of Noise Suppression in Subsonic Round Jets

    NASA Technical Reports Server (NTRS)

    Pruett, C. David

    1997-01-01

    NASA Grant NAG1-1802, originally submitted in June 1996 as a two-year proposal, was awarded one-year's funding by NASA LaRC for the period 5 Oct., 1996, through 4 Oct., 1997. Because of the inavailability (from IT at NASA ARC) of sufficient supercomputer time in fiscal 1998 to complete the computational goals of the second year of the original proposal (estimated to be at least 400 Cray C-90 CPU hours), those goals have been appropriately amended, and a new proposal has been submitted to LaRC as a follow-on to NAG1-1802. The current report documents the activities and accomplishments on NAG1-1802 during the one-year period from 5 Oct., 1996, through 4 Oct., 1997. NASA Grant NAG1-1802, and its predecessor, NAG1-1772, have been directed toward adapting the numerical tool of Large-Eddy Simulation (LES) to aeroacoustic applications, with particular focus on noise suppression in subsonic round jets. In LES, the filtered Navier-Stokes equations are solved numerically on a relatively coarse computational grid. Residual stresses, generated by scales of motion too small to be resolved on the coarse grid, are modeled. Although most LES incorporate spatial filtering, time-domain filtering affords certain conceptual and computational advantages, particularly for aeroacoustic applications. Consequently, this work has focused on the development of SubGrid-Scale (SGS) models that incorporate time- domain filters. The author is unaware of any previous attempt at purely time-filtered LES; however, Aldama and Dakhoul and Bedford have considered approaches that combine both spatial and temporal filtering. In our view, filtering in both space and time is redundant, because removal of high frequencies effects the removal of small spatial scales and vice versa.

  3. Measurements in 80- by 120-foot wind tunnel of hazard posed by lift-generated wakes

    NASA Technical Reports Server (NTRS)

    Rossow, V. J.; Sacco, J. N.; Askins, P. A.; Bisbee, L. S.; Smith, S. M.

    1993-01-01

    The large, low speed wind tunnel at NASA-Ames has been used to study the characteristics of lift-generated vortices involved in the definition of aircraft-separation criteria, in order to enhance airport capacity without compromising safety. Attention is given to the potential hazard caused by the vortex wake of several configurations of a subsonic transport. Measured downwash distributions in the wake of three different wake-generator configurations are obtained by means of a vortex-lattice method, in order to predict the lift and rolling moment on several models of wake-following aircraft.

  4. Low Speed Dynamic Model Investigation of Apollo Command Module Configuration in the Langley Spin Tunnel

    NASA Technical Reports Server (NTRS)

    Lee, Henry A.; Burk, Sanger M., Jr.

    1967-01-01

    An investigation has been conducted in the Langley spin tunnel to determine the dynamic stability of the Apollo command module at low subsonic speeds, both with and without drogue parachutes. The investigation consisted of tests to determine (1) the dynamic stability of the command module alone, (2) the motion of the command module during the deployment of a drogue parachute, (3) the effect of various drogue-parachute configurations on the stability of the command module, and (4) the effect of modifications to the command module to prevent an apex-forward trim condition.

  5. The carpal tunnel.

    PubMed

    Ellis, Harold

    2009-12-01

    The carpal bones are deeply convex anteriorly. This bony gutter is converted by the flexor retinaculum into a tube - the carpal tunnel, which conveys the median nerve, together with the long flexor tendons of the fingers and thumb, into the hand. It is of special interest to the surgeon because it is the site of a common nerve entrapment, the carpal tunnel syndrome.

  6. Shotcrete in tunnel design

    SciTech Connect

    Golser, J.; Galler, R.; Schubert, P.; Rabensteiner, K.

    1995-12-31

    Shotcrete is an important structural element for tunnel support. Green shotcrete is exposed to compression strain rates and tunnel design requires a realistic material law for shotcrete. A modified rate of flow method simulates shotcrete behavior very well and can be incorporated in Finite Element calculations.

  7. Investigation of Low-Subsonic Flight Characteristics of a Model of a Flat-Bottom Hypersonic Boost-Glide Configuration Having a 78 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Paulson, John W.; Shanks, Robert E.

    1959-01-01

    An investigation of the low-subsonic stability and control characteristics of a model of a flat-bottom hypersonic boost-glide configuration having 78 deg sweep of the leading edge has been made in the Langley full-scale tunnel. The model was flown over an angle-of-attack range from 10 to 35 deg. Static and dynamic force tests were made in the Langley free-flight tunnel. The investigation showed that the longitudinal stability and control characteristics were generally satisfactory with neutral or positive static longitudinal stability. The addition of artificial pitch damping resulted in satisfactory longitudinal characteristics being obtained with large amounts of static instability. The most rearward center-of-gravity position for which sustained flights could be made either with or without pitch damper corresponded to the calculated maneuver point. The lateral stability and control characteristics were satisfactory up to about 15 deg angle of attack. The damping of the Dutch roll oscillation decreased with increasing angle of attack; the oscillation was about neutrally stable at 20 deg angle of attack and unstable at angles of attack of about 25 deg and above. Artificial damping in roll greatly improved the lateral characteristics and resulted in flights being made up to 35 deg angle of attack.

  8. Investigation of Low-Subsonic Flight Characteristics of a Model of a Hypersonic Boost-Glide Configuration Having a 78 deg. Delta Wing

    NASA Technical Reports Server (NTRS)

    Paulson, John W.; Shanks, Robert E.

    1961-01-01

    An investigation of the low-subsonic stability and control characteristics of a model of a hypersonic boost-glide configuration having 78 deg. sweep of the leading edge has been made in the Langley full-scale tunnel. The model was flown over an angle-of-attack range from 10 to 35 deg. Static and dynamic force tests were made in the Langley free-flight tunnel. The investigation showed that the longitudinal stability and control characteristics were generally satisfactory with neutral or positive static longitudinal stability. The addition of artificial pitch damping resulted in satisfactory longitudinal characteristics being obtained with large amounts of static instability. The most rearward center-of-gravity position for which sustained flights could be made either with or without pitch damper corresponded to the calculated maneuver point. The lateral stability and control characteristics were satisfactory up to about 15 deg. angle of attack. The damping of the Dutch roll oscillation decreased with increasing angle of attack; the oscillation was about neutrally stable at 20 deg. angle of attack and unstable at angles of attack of about 25 deg. and above. Artificial damping in roll greatly improved the lateral characteristics and resulted in flights being made up to 35 deg. angle of attack.

  9. Investigation of a subsonic-arc-attachment thruster using segmented anodes

    NASA Technical Reports Server (NTRS)

    Berns, Darren H.; Sankovic, John M.; Sarmiento, Charles J.

    1993-01-01

    To investigate high frequency arc instabilities observed in subsonic-arc-attachment thrusters, a 3 kW, segmented-anode arc jet was designed and tested using hydrogen as the propellant. The thruster nozzle geometry was scaled from a 30 kW design previously tested in the 1960's. By observing the current to each segment and the arc voltage, it was determined that the 75-200 kHz instabilities were results of axial movements of the arc anode attachment point. The arc attachment point was fully contained in the subsonic portion of the nozzle for nearly all flow rates. The effects of isolating selected segments were investigated. In some cases, forcing the arc downstream caused the restrike to cease. Finally, decreasing the background pressure from 18 to 0.05 Pa affected the pressure distribution in the nozzle including the pressure in the subsonic arc chamber.

  10. Investigation of a subsonic-arc-attachment thruster using segmented anodes

    NASA Astrophysics Data System (ADS)

    Berns, Darren H.; Sankovic, John M.; Sarmiento, Charles J.

    1993-12-01

    To investigate high frequency arc instabilities observed in subsonic-arc-attachment thrusters, a 3 kW, segmented-anode arc jet was designed and tested using hydrogen as the propellant. The thruster nozzle geometry was scaled from a 30 kW design previously tested in the 1960's. By observing the current to each segment and the arc voltage, it was determined that the 75-200 kHz instabilities were results of axial movements of the arc anode attachment point. The arc attachment point was fully contained in the subsonic portion of the nozzle for nearly all flow rates. The effects of isolating selected segments were investigated. In some cases, forcing the arc downstream caused the restrike to cease. Finally, decreasing the background pressure from 18 to 0.05 Pa affected the pressure distribution in the nozzle including the pressure in the subsonic arc chamber.

  11. The Stability Tunnel

    NASA Technical Reports Server (NTRS)

    1943-01-01

    Engineers operate the controls of the Stability Tunnel: Plans for a new tunnel to study stability problems began in the late thirties. The Stability Tunnel was authorized in 1939 and began operations in June 1941. The installation was completed in December that year with the completion of a new 10,000 Horsepower Diesel-electric generating plant. It was a single return, closed jet tunnel with a 6-foot square test section. The tunnel was disassembled and shipped to Virginia Polytechnic Institute and State University in 1958. The tunnel had two separate test sections: one for curved flow, the other for rolling flow. 'The facility...simulates the motion of the aircraft in curved or rolling flight. This is done by actually curving or rolling the airstream as it passes over the model and at the same time providing the proper velocity distribution.' (From AIAA-80-0309) >From Alan Pope, Wind-Tunnel Testing: 'The only tunnel directly designed for dynamic stability work is located at the Langley Field branch of the NACA. Its most vital feature is its ability to subject the models to curving air streams that simulate those actually encountered when an airplane rolls, pitches, or yaws. the rotating airstream for simulating roll is produced by a motor-driven paddle just ahead of the test section. Curved air of properly varying velocity for simulating pitch and yaw is produced by a combination of a curved test section and velocity screens. The proper use of this apparatus makes possible the determination of the stability derivatives.' Published in F.H. Lutze, 'Experimental Determination of Pure Rotary Stability Derivatives using a Curved and Rolling Flow Wind Tunnel,' AIAA-80-0309, AIAA 18th Aerospace Sciences Meeting, Pasadena, CA, January 14-16, 1980; Alan Pope, Wind-Tunnel Testing (New York: John Wiley & Sons, 1947).

  12. The Liquid Hydrogen Option for the Subsonic Transport: A status report

    NASA Technical Reports Server (NTRS)

    Korycinski, P. F.

    1977-01-01

    Continued subsonic air transport design studies include the option for a liquid hydrogen fuel system as an aircraft fuel conservation measure. Elements of this option discussed include: (1) economical production of hydrogen; (2) efficient liquefaction of hydrogen; (3) materials for long service life LH2 fuel tanks; (4) insulation materials; (5) LH2 fuel service and installations at major air terminals; (6) assessment of LH2 hazards; and (7) the engineering definition of an LH2 fuel system for a large subsonic passenger air transport.

  13. CFD-Based Design Optimization Tool Developed for Subsonic Inlet

    NASA Technical Reports Server (NTRS)

    1995-01-01

    The traditional approach to the design of engine inlets for commercial transport aircraft is a tedious process that ends with a less-than-optimum design. With the advent of high-speed computers and the availability of more accurate and reliable computational fluid dynamics (CFD) solvers, numerical optimization processes can effectively be used to design an aerodynamic inlet lip that enhances engine performance. The designers' experience at Boeing Corporation showed that for a peak Mach number on the inlet surface beyond some upper limit, the performance of the engine degrades excessively. Thus, our objective was to optimize efficiency (minimize the peak Mach number) at maximum cruise without compromising performance at other operating conditions. Using a CFD code NPARC, the NASA Lewis Research Center, in collaboration with Boeing, developed an integrated procedure at Lewis to find the optimum shape of a subsonic inlet lip and a numerical optimization code, ADS. We used a GRAPE-based three-dimensional grid generator to help automate the optimization procedure. The inlet lip shape at the crown and the keel was described as a superellipse, and the superellipse exponents and radii ratios were considered as design variables. Three operating conditions: cruise, takeoff, and rolling takeoff, were considered in this study. Three-dimensional Euler computations were carried out to obtain the flow field. At the initial design, the peak Mach numbers for maximum cruise, takeoff, and rolling takeoff conditions were 0.88, 1.772, and 1.61, respectively. The acceptable upper limits on the takeoff and rolling takeoff Mach numbers were 1.55 and 1.45. Since the initial design provided by Boeing was found to be optimum with respect to the maximum cruise condition, the sum of the peak Mach numbers at takeoff and rolling takeoff were minimized in the current study while the maximum cruise Mach number was constrained to be close to that at the existing design. With this objective, the

  14. Ultrafast scanning tunneling microscopy

    SciTech Connect

    Botkin, D.A. |

    1995-09-01

    I have developed an ultrafast scanning tunneling microscope (USTM) based on uniting stroboscopic methods of ultrafast optics and scanned probe microscopy to obtain nanometer spatial resolution and sub-picosecond temporal resolution. USTM increases the achievable time resolution of a STM by more than 6 orders of magnitude; this should enable exploration of mesoscopic and nanometer size systems on time scales corresponding to the period or decay of fundamental excitations. USTM consists of a photoconductive switch with subpicosecond response time in series with the tip of a STM. An optical pulse from a modelocked laser activates the switch to create a gate for the tunneling current, while a second laser pulse on the sample initiates a dynamic process which affects the tunneling current. By sending a large sequence of identical pulse pairs and measuring the average tunnel current as a function of the relative time delay between the pulses in each pair, one can map the time evolution of the surface process. USTM was used to measure the broadband response of the STM`s atomic size tunnel barrier in frequencies from tens to hundreds of GHz. The USTM signal amplitude decays linearly with the tunnel junction conductance, so the spatial resolution of the time-resolved signal is comparable to that of a conventional STM. Geometrical capacitance of the junction does not appear to play an important role in the measurement, but a capacitive effect intimately related to tunneling contributes to the measured signals and may limit the ultimate resolution of the USTM.

  15. Correaltion of full-scale drag predictions with flight measurements on the C-141A aircraft. Phase 2: Wind tunnel test, analysis, and prediction techniques. Volume 1: Drag predictions, wind tunnel data analysis and correlation

    NASA Technical Reports Server (NTRS)

    Macwilkinson, D. G.; Blackerby, W. T.; Paterson, J. H.

    1974-01-01

    The degree of cruise drag correlation on the C-141A aircraft is determined between predictions based on wind tunnel test data, and flight test results. An analysis of wind tunnel tests on a 0.0275 scale model at Reynolds number up to 3.05 x 1 million/MAC is reported. Model support interference corrections are evaluated through a series of tests, and fully corrected model data are analyzed to provide details on model component interference factors. It is shown that predicted minimum profile drag for the complete configuration agrees within 0.75% of flight test data, using a wind tunnel extrapolation method based on flat plate skin friction and component shape factors. An alternative method of extrapolation, based on computed profile drag from a subsonic viscous theory, results in a prediction four percent lower than flight test data.

  16. Subsonic and supersonic indicial aerodynamics and aerodynamic transfer function for complex configurations. [aerodynamic configurations for subsonic and supersonic speeds using the finite element method

    NASA Technical Reports Server (NTRS)

    Morino, L.

    1974-01-01

    A general theory for indicial-potential-compressible aerodynamics around complex configurations is presented. The motion is assumed to consist of constant subsonic or supersonic speed (steady state) and small perturbations around the steady state. Using the finite-element method to discretize the space problem, a set of differential-difference equations in time relating the potential to its normal derivative on the surface of the body was obtained. The aerodynamics transfer function was derived by using standard method of operational calculus.

  17. [Anterior tarsal tunnel syndrome].

    PubMed

    Miliam, Palle B; Basse, Peter N

    2009-03-30

    Anterior tarsal tunnel syndrome is a rare entrapment neuropathy of the deep peroneal nerve beneath the extensor retinaculum of the ankle. It may be rare because it is underrecognized clinically.We present a case regarding a 29-year-old man, drummer, who for one and a half year experienced clinical symptoms of anterior tarsal tunnel syndrome. A surgical decompression of the anterior tarsal tunnel was performed, and at the check three months later the symptoms where gone. One year after, there were still no symptoms.

  18. Subsonic and Supersonic Flutter Analysis of a Highly Tapered Swept-Wing Planform, Including Effects of Density Variation and Finite Wing Thickness, and Comparison with Experiments

    NASA Technical Reports Server (NTRS)

    Yates, Carson, Jr.

    1967-01-01

    The flutter characteristics of several wings with an aspect-ratio of 4.0, a taper ratio of 0.2, and a quarter-chord sweepback of 45 deg. have been investigated analytically for Mach numbers up to 2.0. The calculations were based on the modified-strip-analysis method, the subsonic-kernel-function method, piston theory, and quasi-steady second-order theory. Results of t h e analysis and comparisons with experiment indicated that: (1) Flutter speeds were accurately predicted by the modified strip analysis, although accuracy at t h e highest Mach numbers required the use of nonlinear aerodynamic theory (which accounts for effects of wing thickness) for the calculation of the aerodynamic parameters. (2) An abrupt increase of flutter-speed coefficient with increasing Mach number, observed experimentally in the transonic range, was also indicated by the modified strip analysis. (3) In the low supersonic range for some densities, a discontinuous variation of flutter frequency with Mach number was indicated by the modified strip analysis. An abrupt change of frequency appeared experimentally in the transonic range. (4) Differences in flutter-speed-coefficient levels obtained from tests at low supersonic Mach numbers in two wind tunnels were also predicted by the modified strip analysis and were shown to be caused primarily by differences in mass ratio. (5) Flutter speeds calculated by the subsonic-kernel-function method were in good agreement with experiment and with the results of the modified strip analysis. (6) Flutter speed obtained from piston theory and from quasi-steady second-order theory were higher than experimental values by at least 38 percent.

  19. Subsonic Reynolds Number Effects on a Diamond Wing Configuration

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.; Ghee, T. A.

    2001-01-01

    An advanced diamond-wing configuration was tested at low speeds in the National Transonic Facility (NTF) in air at chord Reynolds numbers from 4.4 million (typical wind-tunnel conditions) to 24 million (nominal flight value). Extensive variations on high-lift rigging were explored as part of a broad multinational program. The analysis for this study is focused on the cruise and landing settings of the wing high-lift systems. Three flow domains were identified from the data and provide a context for the ensuing data analysis. Reynolds number effects were examined in incremental form based upon attached-flow theory. A similar approach showed very little effect of low-speed compressibility.

  20. Optimization of transonic wind tunnel data acquisition and control systems for providing continuous mode tests

    NASA Astrophysics Data System (ADS)

    Petronevich, V. V.

    2016-10-01

    The paper observes the issues related to the increase of efficiency and information content of experimental research in transonic wind tunnels (WT). In particular, questions of optimizing the WT Data Acquisition and Control Systems (DACS) to provide the continuous mode test method are discussed. The problem of Mach number (M number) stabilization in the test section of the large transonic compressor-type wind tunnels at subsonic flow conditions with continuous change of the aircraft model angle of attack is observed on the example of T-128 wind tunnel. To minimize the signals distortion in T-128 DACS measurement channels the optimal MGCplus filter settings of the data acquisition system used in T-128 wind tunnel to measure loads were experimentally determined. As a result of the tests performed a good agreement of the results of balance measurements for pitch/pause and continuous test modes was obtained. Carrying out balance tests for pitch/pause and continuous test methods was provided by the regular data acquisition and control system of T-128 wind tunnel with unified software package POTOK. The architecture and functional abilities of POTOK software package are observed.

  1. Investigation at High Subsonic Speeds of the Static Longitudinal and Lateral Stability Characteristics of Two Canard Airplane Configurations

    NASA Technical Reports Server (NTRS)

    Sleeman, William C., Jr.

    1957-01-01

    The present investigation was conducted in the Langley high-speed 7-by 10-foot tunnel to determine the static longitudinal and lateral stability characteristics at high subsonic speeds of two canard airplane configurations previously tested at supersonic speeds. The Mach number range of this investigation extended from 0.60 to 0.94 and a maximum angle-of-attack range of -2dewg to 24deg was obtained at the lowest test Mach number. Two wing plan forms of equal area were studied in the present tests; one was a 60deg delta wing and the other was a trapezoid wing having an aspect ratio of 3, taper ratio of 0.143, and an unswept 80-percent-chord line. The canard control had a trapezoidal plan form and its area was approximately 11.5 percent of the wing area. The model also had a low-aspect-ratio highly swept vertical tail and twin ventral fins. The longitudinal control characteristics of the models were consistent with past experience at low speed on canard configurations in that stalling of the canard surface occurred at moderate and high control deflections for moderate values of angle of attack. This stalling could impose appreciable limitations on the maximum trim-lift coefficient attainable. The control effectiveness and maximum value of trim-lift was significantly increased by addition of a body flap having a conical shape and located slightly behind the canard surface on the bottom of the body. Addition of the canard surface at 0deg deflection had relatively little effect on overall directional stability of the delta-wing configuration; however, deflection of the canard surface from 0deg to 10deg had a large favorable effect on directional stability at high angles of attack for both the trapezoid- and delta-wing configurations.

  2. Low Speed Model Investigation of Apollo Command Module Configuration in the Langley Spin Tunnel

    NASA Technical Reports Server (NTRS)

    1967-01-01

    Low Speed Dynamic Model Investigation of Apollo Command Module Configuration in the Langley Spin Tunnel. An investigation has been conducted in the Langley spin tunnel to determine the dynamic stability of the Apollo command module at low subsonic speeds, both with and without drogue parachutes. The investigation consisted of tests to determine (1) the dynamic stability of the command module alone, (2) the motion of the command module during the deployment of a drogue parachute, (3) the effect of various drogue-parachute configurations on the stability of the command module, and (4) the effect of modifications to the command module to prevent an apex-forward trim condition. [Entire movie available on DVD from CASI as Doc ID 20070031002. Contact help@sti.nasa.gov

  3. Sonic wind tunnel of the Institute of Fluid Mechanics of Lille

    NASA Technical Reports Server (NTRS)

    Gontier, G.

    1982-01-01

    A 65 hp wind tunnel with a 40 mm by 240 mm airstream is described. This wind tunnel can achieve speeds in the neighborhood of the speed of sound, both subsonic and supersonic. It is useful in studying the transonic bump technique. The test section is 600 mm long. The side walls are made of transparent glass, and both the upper and lower walls are deformable, each through the use of nine jacks with elastic sleeves. So as to avoid condensation, the airstream's temperature is stabilized by an air exchanger at the temperature of the outside air. The first results for supersonic operation, the distribution of Mach numbers within the airstream between the parallel walls, the value of the use factor, and the diffuser's efficiency are all given.

  4. Carpal tunnel release

    MedlinePlus

    ... you are taking. This includes medicines, supplements, or herbs you bought without a prescription. You may be ... gov/pubmed/23026458 . Zhao M, Burke DT. Median neuropathy (carpal tunnel syndrome). In: Frontera WR, Silver JK, ...

  5. Endoscopic cubital tunnel release.

    PubMed

    Cobb, Tyson K

    2010-10-01

    A minimally invasive endoscopic approach has been successfully applied to surgical treatment of cubital tunnel syndrome. This procedure allows for smaller incisions with faster recovery time. This article details relevant surgical anatomy, indications, contraindications, surgical technique, complications, and postoperative management.

  6. Inelastic electron tunneling spectroscopy

    NASA Technical Reports Server (NTRS)

    Khanna, S. K.; Lambe, J.

    1983-01-01

    Inelastic electron tunneling spectroscopy is a useful technique for the study of vibrational modes of molecules adsorbed on the surface of oxide layers in a metal-insulator-metal tunnel junction. The technique involves studying the effects of adsorbed molecules on the tunneling spectrum of such junctions. The data give useful information about the structure, bonding, and orientation of adsorbed molecules. One of the major advantages of inelastic electron tunneling spectroscopy is its sensitivity. It is capable of detecting on the order of 10 to the 10th molecules (a fraction of a monolayer) on a 1 sq mm junction. It has been successfully used in studies of catalysis, biology, trace impurity detection, and electronic excitations. Because of its high sensitivity, this technique shows great promise in the area of solid-state electronic chemical sensing.

  7. World's Largest Wind Tunnel

    NASA Technical Reports Server (NTRS)

    1987-01-01

    NASA's National Full Scale Aerodynamics Complex, which houses two of the world's largest wind tunnels and has been used for testing experimental aircraft since 1944, is presented. This video highlights the structure and instrumentation of the 40 x 80 foot and 80 x 120 foot wind tunnels and documents their use in testing full scale aircraft, NASA's Space Shuttle and the XV-15 Tiltrotor aircraft.

  8. Electron tunnel sensor technology

    NASA Technical Reports Server (NTRS)

    Kenny, T. W.; Waltman, S. B.; Reynolds, J. K.; Kaiser, W. J.

    1991-01-01

    Researchers designed and constructed a novel electron tunnel sensor which takes advantage of the mechanical properties of micro-machined silicon. For the first time, electrostatic forces are used to control the tunnel electrode separation, thereby avoiding the thermal drift and noise problems associated with piezoelectric actuators. The entire structure is composed of micro-machined silicon single crystals, including a folded cantilever spring and a tip. The application of this sensor to the development of a sensitive accelerometer is described.

  9. Full Scale Wind Tunnel

    NASA Technical Reports Server (NTRS)

    1931-01-01

    Construction of motor fairing for the fan motors of the Full-Scale Tunnel (FST). The motors and their supporting structures were enclosed in aerodynamically smooth fairings to minimize resistance to the air flow. Close examination of this photograph reveals the complicated nature of constructing a wind tunnel. This motor fairing, like almost every other structure in the FST, represents a one-of-a-kind installation.

  10. Formal representation of the requirements for an Advanced Subsonic Civil Transport (ASCT) flight control system

    NASA Technical Reports Server (NTRS)

    Frincke, Deborah; Wolber, Dave; Fisher, Gene; Cohen, Gerald C.; Mclees, R. E.

    1992-01-01

    A partial requirement specification for an Advanced Subsonic Civil Transport (ASCT) Flight Control System is described. The example was adopted from requirements given in a NASA Contractor report. The language used to describe the requirements, Requirements Specification Language (RSL), is described in a companion document.

  11. Wing-Alone Aerodynamic Characteristics to High Angles of Attack at Subsonic and Transonic Speeds.

    DTIC Science & Technology

    1982-11-01

    indicators of symmetry since the wings were unbanked within the limits of tolerances and flow angularity. Longitudinal, spanwise, and vertical... unbanked wings at subsonic and transonic speeds from low to high angles of attack. The wing planforms varied in aspect ratio and taper ratio with

  12. Computer program calculates peripheral water injection cooling of axisymmetric subsonic diffuser

    NASA Technical Reports Server (NTRS)

    Grey, J.

    1968-01-01

    Digital computer program calculates the cooling effectiveness and flow characteristics resulting from the mixing of a cool liquid injectant /water/ with a hot sonic or subsonic gas stream /hydrogen/. The output of the program provides pressure, temperature, velocity, density, composition, and Mach number profiles at any location in the mixing duct.

  13. On the circulatory subsonic flow of a compressible fluid past a circular cylinder

    NASA Technical Reports Server (NTRS)

    Bers, Lipman

    1945-01-01

    The circulatory subsonic flow around an infinite circular cylinder is computed using the linearized pressure-volume relation, by a method developed in a previous report. Formulas and graphs are given for the velocity and pressure distributions, the circulation, the lift, and the dependence of the critical Mach number upon the position of the stagnation point.

  14. Method of Making a Composite Panel Having Subsonic Transverse Wave Speed Characteristics

    NASA Technical Reports Server (NTRS)

    Palumbo, Daniel L. (Inventor); Klos, Jacob (Inventor)

    2012-01-01

    A method of making a composite panel having subsonic transverse wave speed characteristics which has first and second sheets sandwiching a core with at least one of the sheets being attached to the core at first regions thereof and unattached to the core at second regions thereof.

  15. A design approach and selected wind tunnel results at high subsonic speeds for wing-tip mounted winglets

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T.

    1976-01-01

    Winglets, which are small, nearly vertical, winglike surfaces, substantially reduce drag coefficients at lifting conditions. The primary winglet surfaces are rearward above the wing tips; secondary surfaces are forward below the wing tips. This report presents a discussion of the considerations involved in the design of the winglets; measured effects of these surfaces on the aerodynamic forces, moments, and loads for a representative first generation, narrow body jet transport wing; and a comparison of these effects with those for a wing tip extension which results in approximately the same increase in bending moment at the wing-fuselage juncture as did the addition of the winglets.

  16. Design Considerations for a UCAV Wing for Subsonic and Transonic Aeroelastic and Flight Mechanic Wind Tunnel Tests

    DTIC Science & Technology

    2007-11-01

    actuation device in the wing will increase the model complexity considerably and very probably stiffen the wing considerably. Figure 6: Desing ...7] http://www.denel.co.za/Aerospace/UAV.asp [8] http://www.aoe.vt.edu/ research /groups/ucav/ [9] Kudva, J.N.: Overview of the DARPA Smart Wing

  17. Evaluation of Wall Interference Effects in a Two-Dimensional Transonic Wind Tunnel by Subsonic Linear Theory,

    DTIC Science & Technology

    1979-02-01

    44 1 (I U 2t1 ),, \\( ~~,I - ‘ ) t 2 tF ~). . (3.41 ) It iv now p~ v siht ~ to ou ict i .iCr. ltl% nautli.,’ co~Ilici~iitv tiit’.tsUi L-d in t he...is - l i number for one part ict ul,ir chordvv ise pos ition (.v .‘ 0- 3~ Corrected data front both nioshs’is of t he N -u - .‘v-001 2 arc plotted...escel ls’nt .igreem neutt (s ’ l l i is ts l t ’ritlg the small di tie rence in angle of u nctsience). I lie f, ict that cI as~ usa I i ills ’ I f

  18. X-24A Detailing Subsonic Control Surface Configuration

    NASA Technical Reports Server (NTRS)

    1971-01-01

    This annotated photo shows a rear view of the X-24A lifting body research vehicle, emphasizing the control surfaces used for the subsonic portions of the aircraft's flights. The X-24 was one of a group of lifting bodies flown by the NASA Flight Research Center (now Dryden Flight Research Center), Edwards, California, in a joint program with the U.S. Air Force at Edwards Air Force Base from 1963 to 1975. The lifting bodies were used to demonstrate the ability of pilots to maneuver and safely land wingless vehicles designed to fly back to Earth from space and be landed like an airplane at a predetermined site. Lifting bodies' aerodynamic lift, essential to flight in the atmosphere, was obtained from their shape. The addition of fins and control surfaces allowed the pilots to stabilize and control the vehicles and regulate their flight paths. Built by Martin Aircraft Company, Maryland, for the U.S. Air Force, the X-24A was a bulbous vehicle shaped like a teardrop with three vertical fins at the rear for directional control. It weighed 6,270 pounds, was 24.5 feet long and 11.5 feet wide (measuring just the fuselage, not the distance between the tips of the outboard fins). Its first unpowered glide flight was on April 17, 1969, with Air Force Maj. Jerauld Gentry at the controls. Gentry also piloted its first powered flight on March 19, 1970. The X-24A was flown 28 times in the program that, like the HL-10, validated the concept that a Space Shuttle vehicle could be landed unpowered. The fastest speed achieved by the X-24A was 1,036 miles per hour (mph--Mach 1.6). Its maximum altitude was 71,400 feet. It was powered by an XLR-11 rocket engine with a maximum theoretical vacuum thrust of 8,480 pounds. The X-24A was later modified into the X-24B. The bulbous shape of the X-24A was converted into a 'flying flatiron' shape with a rounded top, flat bottom, and double delta platform that ended in a pointed nose. The X-24B demonstrated that accurate unpowered reentry vehicle

  19. Novel tunnelling barriers for spin tunnelling junctions

    NASA Astrophysics Data System (ADS)

    Sharma, Manish

    A tunnel junction consists of two metal electrodes separated by an insulating barrier thin enough for electrons to tunnel across. With ferromagnetic electrodes, a spin-dependent tunnelling (SDT) effect, electrons of one spin tunnelling preferentially over those of the other, is observed. When the electrodes are switched from a parallel to an anti-parallel alignment, the tunnelling current changes and gives rise to tunnelling magnetoresistance (TMR). Since 1995, interest in SDT junctions has increased as TMR in excess of 15% has been achieved, making viable their use in non-volatile memory and magnetic sensors applications. In this work, two key issues of SDT junctions are addressed: spin polarization of the electrode and the tunnel barrier. Spin polarization, a measure of electron states of up and down spins, is widely believed to be an intrinsic property of the electrode. In junctions with barriers formed by plasma oxidation of composite Ta/Al films, the surprising effect of the resistance being lower with the electrodes aligned antiparallel was observed. Junctions with Ta/Al barriers and those with Al/Ta barriers behave opposite to each other and exhibit an inversion only when the Ta side of the barrier is biased positive. This demonstrates the spin polarization is also influenced by the barrier material. Half-metallic materials such as magnetite (Fe3O4) have a gap in one of the spins' states at the fermi level, thus having a theoretical spin polarization of 100%. In this work, an ultrathin Fe3O 4 layer was added between the Al2O3 barrier and the NiFe electrode. The TMR increased sharply from 4% to 16% for thicknesses less than 0.5nm. As the tunnel barrier must be thinner than 2nm, choice of the barrier material becomes critical. Presently, Al2O3 is the best known barrier. In looking for alternative materials, AlN and AlON were formed by plasma nitridation and oxy-nitridation of deposited Al films. TMR results of up to 18% and resistance-area products down to 3

  20. The Beginner's Guide to Wind Tunnels with TunnelSim and TunnelSys

    NASA Technical Reports Server (NTRS)

    Benson, Thomas J.; Galica, Carol A.; Vila, Anthony J.

    2010-01-01

    The Beginner's Guide to Wind Tunnels is a Web-based, on-line textbook that explains and demonstrates the history, physics, and mathematics involved with wind tunnels and wind tunnel testing. The Web site contains several interactive computer programs to demonstrate scientific principles. TunnelSim is an interactive, educational computer program that demonstrates basic wind tunnel design and operation. TunnelSim is a Java (Sun Microsystems Inc.) applet that solves the continuity and Bernoulli equations to determine the velocity and pressure throughout a tunnel design. TunnelSys is a group of Java applications that mimic wind tunnel testing techniques. Using TunnelSys, a team of students designs, tests, and post-processes the data for a virtual, low speed, and aircraft wing.

  1. Subsonic Aerodynamic Characteristics of an Airplane Configuration with a 63 deg Sweptback Wing and Twin-Boom Tails

    NASA Technical Reports Server (NTRS)

    Savage, Howard F.; Edwards, George G.

    1959-01-01

    A wind-tunnel investigation has been conducted to determine the effects of an unconventional tail arrangement on the subsonic static longitudinal and lateral stability characteristics of a model having a 63 deg sweptback wing of aspect ratio 3.5 and a fuselage. Tail booms, extending rearward from approximately the midsemispan of each wing panel, supported independent tail assemblies well outboard of the usual position at the rear of the fuselage. The horizontal-tail surfaces had the leading edge swept back 45 deg and an aspect ratio of 2.4. The vertical tail surfaces were geometrically similar to one panel of the horizontal tail. For comparative purposes, the wing-body combination was also tested with conventional fuselage-mounted tail surfaces. The wind-tunnel tests were conducted at Mach numbers from 0.25 to 0.95 with a Reynolds number of 2,000,000, at a Mach number of 0.46 with a Reynolds number of 3,500,000, and at a Mach number of 0.20 with a Reynolds number of 7,000,000. The results of the investigation indicate that longitudinal stability existed to considerably higher lift coefficients for the outboard tail configuration than for the configuration with conventional tail. Wing fences were necessary with both configurations for the elimination of sudden changes in longitudinal stability at lift coefficients between 0.3 and 0.5. Sideslip angles up to 15 deg had only small effects upon the pitching-moment characteristics of the outboard tail configuration. There was an increase in the directional stability for the outboard tail configuration at the higher angles of attack as opposed to a decrease for the conventional tail configuration at most of the Mach numbers and Reynolds numbers of this investigation. The dihedral effect increased rapidly with increasing angle of attack for both the outboard and the conventional tail configurations but the increase was greater for the outboard tail configuration. The data indicate that the outboard tail is an effective

  2. On the Importance of Very Light Internally Subsonic AGN Jets in Radio-mode AGN Feedback

    NASA Astrophysics Data System (ADS)

    Guo, Fulai

    2016-07-01

    Radio-mode active galactic nucleus (AGN) feedback plays a key role in the evolution of galaxy groups and clusters. Its physical origin lies in the kiloparsec-scale interaction of AGN jets with the intracluster medium. Large-scale jet simulations often initiate light internally supersonic jets with density contrast 0.01 < η < 1. Here we argue for the first time for the importance of very light (η < 0.01) internally subsonic jets. We investigated the shapes of young X-ray cavities produced in a suite of hydrodynamic simulations, and found that bottom-wide cavities are always produced by internally subsonic jets, while internally supersonic jets inflate cylindrical, center-wide, or top-wide cavities. We found examples of real cavities with shapes analogous to those inflated in our simulations by internally subsonic and internally supersonic jets, suggesting a dichotomy of AGN jets according to their internal Mach numbers. We further studied the long-term cavity evolution, and found that old cavities resulted from light jets spread along the jet direction, while those produced by very light jets are significantly elongated along the perpendicular direction. The northwestern ghost cavity in Perseus is pancake shaped, providing tentative evidence for the existence of very light jets. Our simulations show that very light internally subsonic jets decelerate faster and rise much slower in the intracluster medium than light internally supersonic jets, possibly depositing a larger fraction of jet energy to cluster cores and alleviating the problem of low coupling efficiencies found previously. The internal Mach number points to the jet’s energy content, and internally subsonic jets are energetically dominated by non-kinetic energy, such as thermal energy, cosmic rays, or magnetic fields.

  3. The Real-Time Wall Interference Correction System of the NASA Ames 12-Foot Pressure Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Ulbrich, Norbert

    1998-01-01

    An improved version of the Wall Signature Method was developed to compute wall interference effects in three-dimensional subsonic wind tunnel testing of aircraft models in real-time. The method may be applied to a full-span or a semispan model. A simplified singularity representation of the aircraft model is used. Fuselage, support system, propulsion simulator, and separation wake volume blockage effects are represented by point sources and sinks. Lifting effects are represented by semi-infinite line doublets. The singularity representation of the test article is combined with the measurement of wind tunnel test reference conditions, wall pressure, lift force, thrust force, pitching moment, rolling moment, and pre-computed solutions of the subsonic potential equation to determine first order wall interference corrections. Second order wall interference corrections for pitching and rolling moment coefficient are also determined. A new procedure is presented that estimates a rolling moment coefficient correction for wings with non-symmetric lift distribution. Experimental data obtained during the calibration of the Ames Bipod model support system and during tests of two semispan models mounted on an image plane in the NASA Ames 12 ft. Pressure Wind Tunnel are used to demonstrate the application of the wall interference correction method.

  4. Experimental Pressure Distributions over Wing Tips at Mach Number 1.9 I : Wing Tip with Subsonic Leading Edge

    NASA Technical Reports Server (NTRS)

    Jagger, James M; Mirels, Harold

    1949-01-01

    An investigation was conducted at a Mach number of 1.91 to determine spanwise pressure distribution over a wing tip in a region influenced by a sharp subsonic leading edge swept back at 70 degrees. Except for pressure distribution on the top surface in the immediate vicinity of the subsonic leading edge, the maximum difference between linearized theory and experimental data was 2 1/2 percent (of free-stream dynamic pressure) for angles of attack up to 4 degrees and 7 percent for angles of attack up to 8 degrees. Pressures on the top surface nearest the subsonic edge indicated local expansions beyond values predicted by linearized theory.

  5. Programmable controller system for wind tunnel diversion vanes

    NASA Technical Reports Server (NTRS)

    King, R. F.

    1982-01-01

    A programmable controller (PC) system automatic sequence control, which acts as a supervisory controller for the servos, selects the proper drives, and automatically sequences the vanes, was developed for use in a subsonic wind tunnel. Tunnel modifications include a new second test section (80 ft x 100 ft with a maximum air speed capability of 110 knots) and an increase in maximum velocity flow from 200 knots to 300 knots. A completely automatic sequence control is necessary in order to allow intricate motion of the 14 triangularly arranged vanes which can be as large as 70 ft high x 35 ft wide and which require precise acceleration and deceleration control. Rate servos on each drive aid in this control, and servo cost was minimized by using four silicon controlled rectifier controllers to control the 20 dc drives. The PC has a programming capacity which facilitated the implementation of extensive logic design. A series of diagrams sequencing the vanes and a block diagram of the system are included.

  6. Exploratory flutter test in a cryogenic wind tunnel

    NASA Technical Reports Server (NTRS)

    Cole, S. R.

    1985-01-01

    A model consisting of a rigid wing with an integral, flexible beam support that was cantilever mounted from the wall in the NASA LaRC 0.3-m transonic cryogenic tunnel was used in a flutter analysis study. The wing had a rectangular planform of aspect ratio 1.5 and a 64A010 airfoil. Various considerations and procedures for conducting flutter tests in a cryogenic wind tunnel were evaluated. Flutter onset conditions were established from extrapolated subcritical response measurements. A flutter boundary was determined at cryogenic temperatures over a Mach number M range from 0.5 to 0.9. Flutter was obtained at two different Reynolds numbers R at M = 0.5 (R = 4.4 and 18.4 x 10 to the 6th power) and at M = 0.8 (R = 5.0 and 10.4 x 10 to the 6th power). Flutter analyses using subsonic lifting surface (kernel function) aerodynamics were made over the range of test conditions. To evaluate the Reynolds number effects at M = 0.5 and 0.8, the experimental results were adjusted using analytical trends to account for differences in the model test temperatures and mass ratios. The adjusted experimental results indicate that increasing Reynolds number from 5.0 to 20.0 x 10 to the 6th power decreased the dynamic pressure by 4.0 to 6.5 percent at M = 0.5 and 0.8.

  7. Single Electron Tunneling

    SciTech Connect

    Ruggiero, Steven T.

    2005-07-25

    Financial support for this project has led to advances in the science of single-electron phenomena. Our group reported the first observation of the so-called ''Coulomb Staircase'', which was produced by tunneling into ultra-small metal particles. This work showed well-defined tunneling voltage steps of width e/C and height e/RC, demonstrating tunneling quantized on the single-electron level. This work was published in a now well-cited Physical Review Letter. Single-electron physics is now a major sub-field of condensed-matter physics, and fundamental work in the area continues to be conducted by tunneling in ultra-small metal particles. In addition, there are now single-electron transistors that add a controlling gate to modulate the charge on ultra-small photolithographically defined capacitive elements. Single-electron transistors are now at the heart of at least one experimental quantum-computer element, and single-electron transistor pumps may soon be used to define fundamental quantities such as the farad (capacitance) and the ampere (current). Novel computer technology based on single-electron quantum dots is also being developed. In related work, our group played the leading role in the explanation of experimental results observed during the initial phases of tunneling experiments with the high-temperature superconductors. When so-called ''multiple-gap'' tunneling was reported, the phenomenon was correctly identified by our group as single-electron tunneling in small grains in the material. The main focus throughout this project has been to explore single electron phenomena both in traditional tunneling formats of the type metal/insulator/particles/insulator/metal and using scanning tunneling microscopy to probe few-particle systems. This has been done under varying conditions of temperature, applied magnetic field, and with different materials systems. These have included metals, semi-metals, and superconductors. Amongst a number of results, we have

  8. Suppression of tunneling rate fluctuations in tunnel field-effect transistors by enhancing tunneling probability

    NASA Astrophysics Data System (ADS)

    Mori, Takahiro; Migita, Shinji; Fukuda, Koichi; Asai, Hidehiro; Morita, Yukinori; Mizubayashi, Wataru; Liu, Yongxun; O’uchi, Shin-ichi; Fuketa, Hiroshi; Otsuka, Shintaro; Yasuda, Tetsuji; Masahara, Meishoku; Ota, Hiroyuki; Matsukawa, Takashi

    2017-04-01

    This paper discusses the impact of the tunneling probability on the variability of tunnel field-effect transistors (TFETs). Isoelectronic trap (IET) technology, which enhances the tunneling current in TFETs, is used to suppress the variability of the ON current and threshold voltage. The simulation results show that suppressing the tunneling rate fluctuations results in suppression of the variability. In addition, a formula describing the relationship between the tunneling rate fluctuations and the electric field strength is derived based on Kane’s band-to-band tunneling model. This formula indicates that the magnitude of the tunneling rate fluctuations is proportional to the magnitude of the fluctuations in the electric field strength and a higher tunneling probability results in a lower variability. The derived relationship is universally valid for any technologies that exploit enhancement of the tunneling probability, including IET technology, channel material engineering, heterojunctions, strain engineering, etc.

  9. Carpal Tunnel Syndrome

    PubMed Central

    Zimmerman, Gregory R.

    1994-01-01

    Carpal tunnel syndrome is a neuropathy resulting from compression of the median nerve as it passes through a narrow tunnel in the wrist on its way to the hand. The lack of precise objective and clinical tests, along with symptoms that are synonymous with other syndromes in the upper extremity, cause carpal tunnel syndrome to appear to be a rare entity in athletics. However, it should not be ruled out as a possible etiology of upper extremity paralysis in the athlete. More typically, carpal tunnel syndrome is the most common peripheral entrapment neuropathy encountered in industry. Treatment may include rest and/or splinting of the involved wrist, ice application, galvanic stimulation, or iontophoresis to reduce inflammation, and then transition to heat modalities and therapeutic exercises for developing flexibility, strength, and endurance. In addition, an ergonomic assessment should be conducted, resulting in modifications to accommodate the carpal tunnel syndrome patient. ImagesFig 3.Fig 4.Fig 5.Fig 6.Fig 7. PMID:16558255

  10. Condensate Mixtures and Tunneling

    SciTech Connect

    Timmermans, E.

    1998-09-14

    The experimental study of condensate mixtures is a particularly exciting application of the recently developed atomic-trap Bose-Einstein condensate (BEC) technology: such multiple condensates represent the first laboratory systems of distinguishable boson superfluid mixtures. In addition, as the authors point out in this paper, the possibility of inter-condensate tunneling greatly enhances the richness of the condensate mixture physics. Not only does tunneling give rise to the oscillating particle currents between condensates of different chemical potentials, such as those studied extensively in the condensed matter Josephson junction experiments, it also affects the near-equilibrium dynamics and stability of the condensate mixtures. In particular, the stabilizing influence of tunneling with respect to spatial separation (phase separation) could be of considerable practical importance to the atomic trap systems. Furthermore, the creation of mixtures of atomic and molecular condensates could introduce a novel type of tunneling process, involving the conversion of a pair of atomic condensate bosons into a single molecular condensate boson. The static description of condensate mixtures with such type of pair tunneling suggests the possibility of observing dilute condensates with the liquid-like property of a self-determined density.

  11. Applying Pressure Sensitive Paint Technology to Rotor Blades

    NASA Technical Reports Server (NTRS)

    Watkins, A. Neal; Leighty, Bradley D.; Lipford, William E.; Goodman, Kyle Z.; Crafton, Jim; Gregory, James W.

    2014-01-01

    This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on rotorcrtaft blades in simulated forward flight at the 14- by 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The basics of the PSP method will be discussed and the modifications that were needed to extend this technology for use on rotor blades. Results from a series of tests will also be presented as well as several areas of improvement that have been identified and are currently being developed for future testing.

  12. A Correlation Between Flight-Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft

    NASA Technical Reports Server (NTRS)

    Sim, Alex G.

    1997-01-01

    Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihood estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5 deg. to 15.7 deg. Data are presented for a subsonic and transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.

  13. A Correlation Between Flight-Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft

    NASA Technical Reports Server (NTRS)

    Sim, Alex G.

    1976-01-01

    Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihooa estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5deg to 15.7deg. Data are presented for a subsonic and a transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.

  14. Femtosecond scanning tunneling microscope

    SciTech Connect

    Taylor, A.J.; Donati, G.P.; Rodriguez, G.; Gosnell, T.R.; Trugman, S.A.; Some, D.I.

    1998-11-01

    This is the final report of a three-year, Laboratory Directed Research and Development (LDRD) project at the Los Alamos National Laboratory (LANL). By combining scanning tunneling microscopy with ultrafast optical techniques we have developed a novel tool to probe phenomena on atomic time and length scales. We have built and characterized an ultrafast scanning tunneling microscope in terms of temporal resolution, sensitivity and dynamic range. Using a novel photoconductive low-temperature-grown GaAs tip, we have achieved a temporal resolution of 1.5 picoseconds and a spatial resolution of 10 nanometers. This scanning tunneling microscope has both cryogenic and ultra-high vacuum capabilities, enabling the study of a wide range of important scientific problems.

  15. Uncooled tunneling infrared sensor

    NASA Technical Reports Server (NTRS)

    Kenny, Thomas W. (Inventor); Kaiser, William J. (Inventor); Podosek, Judith A. (Inventor); Vote, Erika C. (Inventor); Muller, Richard E. (Inventor); Maker, Paul D. (Inventor)

    1995-01-01

    An uncooled infrared tunneling sensor in which the only moving part is a diaphragm which is deflected into contact with a micromachined silicon tip electrode prepared by a novel lithographic process. Similarly prepared deflection electrodes employ electrostatic force to control the deflection of a silicon nitride, flat diaphragm membrane. The diaphragm exhibits a high resonant frequency which reduces the sensor's sensitivity to vibration. A high bandwidth feedback circuit controls the tunneling current by adjusting the deflection voltage to maintain a constant deflection of the membrane. The resulting infrared sensor can be miniaturized to pixel dimensions smaller than 100 .mu.m. An alternative embodiment is implemented using a corrugated membrane to permit large deflection without complicated clamping and high deflection voltages. The alternative embodiment also employs a pinhole aperture in a membrane to accommodate environmental temperature variation and a sealed chamber to eliminate environmental contamination of the tunneling electrodes and undesireable accoustic coupling to the sensor.

  16. View down tank tunnel (tunnel no. 2) showing pipes and ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    View down tank tunnel (tunnel no. 2) showing pipes and walkway of metal grating, side tunnel to tank 3 is on the left - U.S. Naval Base, Pearl Harbor, Diesel Purification Plant, North Road near Pierce Street, Pearl City, Honolulu County, HI

  17. Aorto-ventricular tunnel

    PubMed Central

    McKay, Roxane

    2007-01-01

    Aorto-ventricular tunnel is a congenital, extracardiac channel which connects the ascending aorta above the sinutubular junction to the cavity of the left, or (less commonly) right ventricle. The exact incidence is unknown, estimates ranging from 0.5% of fetal cardiac malformations to less than 0.1% of congenitally malformed hearts in clinico-pathological series. Approximately 130 cases have been reported in the literature, about twice as many cases in males as in females. Associated defects, usually involving the proximal coronary arteries, or the aortic or pulmonary valves, are present in nearly half the cases. Occasional patients present with an asymptomatic heart murmur and cardiac enlargement, but most suffer heart failure in the first year of life. The etiology of aorto-ventricular tunnel is uncertain. It appears to result from a combination of maldevelopment of the cushions which give rise to the pulmonary and aortic roots, and abnormal separation of these structures. Echocardiography is the diagnostic investigation of choice. Antenatal diagnosis by fetal echocardiography is reliable after 18 weeks gestation. Aorto-ventricular tunnel must be distinguished from other lesions which cause rapid run-off of blood from the aorta and produce cardiac failure. Optimal management of symptomatic aorto-ventricular tunnel consists of diagnosis by echocardiography, complimented with cardiac catheterization as needed to elucidate coronary arterial origins or associated defects, and prompt surgical repair. Observation of the exceedingly rare, asymptomatic patient with a small tunnel may be justified by occasional spontaneous closure. All patients require life-long follow-up for recurrence of the tunnel, aortic valve incompetence, left ventricular function, and aneurysmal enlargement of the ascending aorta. PMID:17922908

  18. Aorto-ventricular tunnel.

    PubMed

    McKay, Roxane

    2007-10-08

    Aorto-ventricular tunnel is a congenital, extracardiac channel which connects the ascending aorta above the sinutubular junction to the cavity of the left, or (less commonly) right ventricle. The exact incidence is unknown, estimates ranging from 0.5% of fetal cardiac malformations to less than 0.1% of congenitally malformed hearts in clinico-pathological series. Approximately 130 cases have been reported in the literature, about twice as many cases in males as in females. Associated defects, usually involving the proximal coronary arteries, or the aortic or pulmonary valves, are present in nearly half the cases. Occasional patients present with an asymptomatic heart murmur and cardiac enlargement, but most suffer heart failure in the first year of life. The etiology of aorto-ventricular tunnel is uncertain. It appears to result from a combination of maldevelopment of the cushions which give rise to the pulmonary and aortic roots, and abnormal separation of these structures. Echocardiography is the diagnostic investigation of choice. Antenatal diagnosis by fetal echocardiography is reliable after 18 weeks gestation. Aorto-ventricular tunnel must be distinguished from other lesions which cause rapid run-off of blood from the aorta and produce cardiac failure. Optimal management of symptomatic aorto-ventricular tunnel consists of diagnosis by echocardiography, complimented with cardiac catheterization as needed to elucidate coronary arterial origins or associated defects, and prompt surgical repair. Observation of the exceedingly rare, asymptomatic patient with a small tunnel may be justified by occasional spontaneous closure. All patients require life-long follow-up for recurrence of the tunnel, aortic valve incompetence, left ventricular function, and aneurysmal enlargement of the ascending aorta.

  19. Advanced subsonic Technology Noise Reduction Element Separate Flow Nozzle Tests for Engine Noise Reduction Sub-Element

    NASA Technical Reports Server (NTRS)

    Saiyed, Naseem H.

    2000-01-01

    Contents of this presentation include: Advanced Subsonic Technology (AST) goals and general information; Nozzle nomenclature; Nozzle schematics; Photograph of all baselines; Configurations tests and types of data acquired; and Engine cycle and plug geometry impact on EPNL.

  20. Future tunnelling projects in Iceland

    SciTech Connect

    Jonsson, B. )

    1992-04-01

    More than 300 km of hydro tunnels and 80-90 km of road tunnels could be excavated in Iceland before the year 2050. In order to complete this task, an average of 6-7 km of tunnel per year would have to be driven. This volume of tunnelling is estimated to cost more than $US1 billion, which could be divided as follows: (a) about 100 km of 3.5-m-wide diversion hydro tunnels (unsupported), for a total of $90 million; (b) approx. 100 km of 5-m-wide hydro tunnels (supported), for a total of $210 million; (c) about 100 km of 7.6-m-wide hydro tunnels (supported), for a total of $380 million; and (d) approx. 85 km of road tunnels with 25 m[sup 2] cross-section, for a total of $435 million. 5 refs., 5 figs., 4 tabs.

  1. Carpal Tunnel Syndrome (For Kids)

    MedlinePlus

    ... Dictionary of Medical Words En Español What Other Kids Are Reading Taking Care of Your Ears Taking ... an X-ray Carpal Tunnel Syndrome KidsHealth > For Kids > Carpal Tunnel Syndrome Print A A A What's ...

  2. Instrumentation in wind tunnels

    NASA Technical Reports Server (NTRS)

    Takashima, K.

    1986-01-01

    Requirements in designing instrumentation systems and measurements of various physical quantities in wind tunnels are surveyed. Emphasis is given to sensors used for measuring pressure, temperature, and angle, and the measurements of air turbulence and boundary layers. Instrumentation in wind tunnels require accuracy, fast response, diversity and operational simplicity. Measurements of force, pressure, attitude angle, free flow, pressure distribution, and temperature are illustrated by a table, and a block diagram. The LDV (laser Doppler velocimeter) method for measuring air turbulence and flow velocity and measurement of skin friction and flow fields using laser holograms are discussed. The future potential of these techniques is studied.

  3. Tunneling in axion monodromy

    NASA Astrophysics Data System (ADS)

    Brown, Jon; Cottrell, William; Shiu, Gary; Soler, Pablo

    2016-10-01

    The Coleman formula for vacuum decay and bubble nucleation has been used to estimate the tunneling rate in models of axion monodromy in recent literature. However, several of Coleman's original assumptions do not hold for such models. Here we derive a new estimate with this in mind using a similar Euclidean procedure. We find that there are significant regions of parameter space for which the tunneling rate in axion monodromy is not well approximated by the Coleman formula. However, there is also a regime relevant to large field inflation in which both estimates parametrically agree. We also briefly comment on the applications of our results to the relaxion scenario.

  4. Effects of winglets on a first-generation jet transport wing. 7: Sideslip effects on winglet loads and selected wing loads at subsonic speeds for a full-span model

    NASA Technical Reports Server (NTRS)

    Meyer, Robert R., Jr.; Covell, Peter F.

    1986-01-01

    The effect of sideslip on winglet loads and selected wing loads was investigated at high and low subsonic Mach numbers. The investigation was conducted in two separate wind tunnel facilities, using two slightly different 0.035-scale full-span models. Results are presented which indicate that, in general, winglet loads as a result of sideslip are analogous to wing loads caused by angle of attack. The center-of-pressure locations on the winglets are somewhat different than might be expected for an analogous wing. The spanwise center of pressure for a winglet tends to be more inboard than for a wing. The most notable chordwise location is a forward center-of-pressure location on the winglet at high sideslip angles. The noted differences between a winglet and an analogous wing are the result of the influence of the wing on the winglet.

  5. Scanning tunneling microscope nanoetching method

    DOEpatents

    Li, Yun-Zhong; Reifenberger, Ronald G.; Andres, Ronald P.

    1990-01-01

    A method is described for forming uniform nanometer sized depressions on the surface of a conducting substrate. A tunneling tip is used to apply tunneling current density sufficient to vaporize a localized area of the substrate surface. The resulting depressions or craters in the substrate surface can be formed in information encoding patterns readable with a scanning tunneling microscope.

  6. Application of Pressure-Based Wall Correction Methods to Two NASA Langley Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Iyer, V.; Everhart, J. L.

    2001-01-01

    This paper is a description and status report on the implementation and application of the WICS wall interference method to the National Transonic Facility (NTF) and the 14 x 22-ft subsonic wind tunnel at the NASA Langley Research Center. The method calculates free-air corrections to the measured parameters and aerodynamic coefficients for full span and semispan models when the tunnels are in the solid-wall configuration. From a data quality point of view, these corrections remove predictable bias errors in the measurement due to the presence of the tunnel walls. At the NTF, the method is operational in the off-line and on-line modes, with three tests already computed for wall corrections. At the 14 x 22-ft tunnel, initial implementation has been done based on a test on a full span wing. This facility is currently scheduled for an upgrade to its wall pressure measurement system. With the addition of new wall orifices and other instrumentation upgrades, a significant improvement in the wall correction accuracy is expected.

  7. Atmospheric Effects of Aviation: First Report of the Subsonic Assessment Project

    NASA Technical Reports Server (NTRS)

    Thompson, Anne M. (Editor); Friedl, Randall R. (Editor); Wesoky, Howard L. (Editor)

    1996-01-01

    This document is the first report from the Office of Aeronautics Advanced Subsonic Technology (AST) Program's Subsonic Assessment (SASS) Project. This effort, initiated in late 1993, has as its objective the assessment of the atmospheric effects of the current and predicted future aviation fleet. The two areas of impact are ozone (stratospheric and tropospheric) and radiative forcing. These are driven, respectively, by possible perturbations from aircraft emissions of NOX and soot and/or sulfur-containing particles. The report presents the major questions to which project assessments will be directed (Introduction) and the status of six programmatic elements: Emissions Scenarios, Exhaust Characterization, Near-Field Interactions, Kinetics and Laboratory Studies, Global Modeling, and Atmospheric Observations (field studies).

  8. INTERACTION OF LASER RADIATION WITH MATTER. LASER PLASMA: Subsonic radiation waves in neon

    NASA Astrophysics Data System (ADS)

    Loseva, T. V.; Nemchinov, I. V.

    1989-02-01

    Numerical methods are used to investigate the propagation of plane subsonic radiation waves in neon from an obstacle in the direction opposite to the incident radiation of Nd and CO2 lasers. An analysis is made of the influence of the power density of the incident radiation (in the range 10-100 MW/cm2) and of the initial density of neon (beginning from the normal valuep ρ0 up to 10ρ0) on the various characteristics of subsonic radiation waves. It is shown that waves traveling in neon can provide an effective source of radiation with a continuous spectrum and an efficiency of ~ 12-27% in the ultraviolet range (with a characteristic photon energy ~ 5-10 eV).

  9. Fluctuation diagrams for hot-wire anemometry in subsonic compressible flows

    NASA Technical Reports Server (NTRS)

    Stainback, P. C.; Nagabushana, K. A.

    1991-01-01

    The concept of using 'fluctuation diagrams' for describing basic fluctuations in compressible flows was reported by Kovasznay in the 1950's. The application of this technique, for the most part, was restricted to supersonic flows. Recently, Zinovev and Lebiga published reports where they considered the fluctuation diagrams in subsonic compressible flows. For the above studies, the velocity and density sensitivities of the heated wires were equal. However, there are considerable data, much taken in the 1950's, which indicate that under some conditions the velocity and density sensitivities are not equal in subsonic compressible flows. Therefore, possible fluctuation diagrams are described for the cases where the velocity and density sensitivities are equal and the more general cases where they are unequal.

  10. Generalization of the subsonic kernel function in the s-plane, with applications to flutter analysis

    NASA Technical Reports Server (NTRS)

    Cunningham, H. J.; Desmarais, R. N.

    1984-01-01

    A generalized subsonic unsteady aerodynamic kernel function, valid for both growing and decaying oscillatory motions, is developed and applied in a modified flutter analysis computer program to solve the boundaries of constant damping ratio as well as the flutter boundary. Rates of change of damping ratios with respect to dynamic pressure near flutter are substantially lower from the generalized-kernel-function calculations than from the conventional velocity-damping (V-g) calculation. A rational function approximation for aerodynamic forces used in control theory for s-plane analysis gave rather good agreement with kernel-function results, except for strongly damped motion at combinations of high (subsonic) Mach number and reduced frequency.

  11. Design Sensitivity for a Subsonic Aircraft Predicted by Neural Network and Regression Models

    NASA Technical Reports Server (NTRS)

    Hopkins, Dale A.; Patnaik, Surya N.

    2005-01-01

    A preliminary methodology was obtained for the design optimization of a subsonic aircraft by coupling NASA Langley Research Center s Flight Optimization System (FLOPS) with NASA Glenn Research Center s design optimization testbed (COMETBOARDS with regression and neural network analysis approximators). The aircraft modeled can carry 200 passengers at a cruise speed of Mach 0.85 over a range of 2500 n mi and can operate on standard 6000-ft takeoff and landing runways. The design simulation was extended to evaluate the optimal airframe and engine parameters for the subsonic aircraft to operate on nonstandard runways. Regression and neural network approximators were used to examine aircraft operation on runways ranging in length from 4500 to 7500 ft.

  12. An Analytical Assessment of NASA's N(+)1 Subsonic Fixed Wing Project Noise Goal

    NASA Technical Reports Server (NTRS)

    Berton, Jeffrey J.; Envia, Edmane; Burley, Casey L.

    2010-01-01

    The Subsonic Fixed Wing Project of NASA s Fundamental Aeronautics Program has adopted a noise reduction goal for new, subsonic, single-aisle, civil aircraft expected to replace current 737 and A320 airplanes. These so-called "N+1" aircraft--designated in NASA vernacular as such since they will follow the current, in-service, "N" airplanes--are hoped to achieve certification noise goal levels of 32 cumulative EPNdB under current Stage 4 noise regulations. A notional, N+1, single-aisle, twinjet transport with ultrahigh bypass ratio turbofan engines is analyzed in this study using NASA software and methods. Several advanced noise-reduction technologies are empirically applied to the propulsion system and airframe. Certification noise levels are predicted and compared with the NASA goal.

  13. An Analytical Assessment of NASA's N+1 Subsonic Fixed Wing Project Noise Goal

    NASA Technical Reports Server (NTRS)

    Berton, Jeffrey J.; Envia, Edmane; Burley, Casey L.

    2009-01-01

    The Subsonic Fixed Wing Project of NASA's Fundamental Aeronautics Program has adopted a noise reduction goal for new, subsonic, single-aisle, civil aircraft expected to replace current 737 and A320 airplanes. These so-called 'N+1' aircraft - designated in NASA vernacular as such since they will follow the current, in-service, 'N' airplanes - are hoped to achieve certification noise goal levels of 32 cumulative EPNdB under current Stage 4 noise regulations. A notional, N+1, single-aisle, twinjet transport with ultrahigh bypass ratio turbofan engines is analyzed in this study using NASA software and methods. Several advanced noise-reduction technologies are analytically applied to the propulsion system and airframe. Certification noise levels are predicted and compared with the NASA goal.

  14. Materials and Structures Research for Gas Turbine Applications Within the NASA Subsonic Fixed Wing Project

    NASA Technical Reports Server (NTRS)

    Hurst, Janet

    2011-01-01

    A brief overview is presented of the current materials and structures research geared toward propulsion applications for NASA s Subsonic Fixed Wing Project one of four projects within the Fundamental Aeronautics Program of the NASA Aeronautics Research Mission Directorate. The Subsonic Fixed Wing (SFW) Project has selected challenging goals which anticipate an increasing emphasis on aviation s impact upon the global issue of environmental responsibility. These goals are greatly reduced noise, reduced emissions and reduced fuel consumption and address 25 to 30 years of technology development. Successful implementation of these demanding goals will require development of new materials and structural approaches within gas turbine propulsion technology. The Materials and Structures discipline, within the SFW project, comprise cross-cutting technologies ranging from basic investigations to component validation in laboratory environments. Material advances are teamed with innovative designs in a multidisciplinary approach with the resulting technology advances directed to promote the goals of reduced noise and emissions along with improved performance.

  15. Far-Field Turbulent Vortex-Wake/Exhaust Plume Interaction for Subsonic and HSCT Airplanes

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Adam, Ihab; Wong, Tin-Chee

    1996-01-01

    Computational study of the far-field turbulent vortex-wake/exhaust plume interaction for subsonic and high speed civil transport (HSCT) airplanes is carried out. The Reynolds-averaged Navier-Stokes (NS) equations are solved using the implicit, upwind, Roe-flux-differencing, finite-volume scheme. The two-equation shear stress transport model of Menter is implemented with the NS solver for turbulent-flow calculation. For the far-field study, the computations of vortex-wake interaction with the exhaust plume of a single engine of a Boeing 727 wing in a holding condition and two engines of an HSCT in a cruise condition are carried out using overlapping zonal method for several miles downstream. These results are obtained using the computer code FTNS3D. The results of the subsonic flow of this code are compared with those of a parabolized NS solver known as the UNIWAKE code.

  16. Validation of a reduced-order jet model for subsonic and underexpanded hydrogen jets

    SciTech Connect

    Li, Xuefang; Hecht, Ethan S.; Christopher, David M.

    2016-01-01

    Much effort has been made to model hydrogen releases from leaks during potential failures of hydrogen storage systems. A reduced-order jet model can be used to quickly characterize these flows, with low computational cost. Notional nozzle models are often used to avoid modeling the complex shock structures produced by the underexpanded jets by determining an “effective” source to produce the observed downstream trends. In our work, the mean hydrogen concentration fields were measured in a series of subsonic and underexpanded jets using a planar laser Rayleigh scattering system. Furthermore, we compared the experimental data to a reduced order jet model for subsonic flows and a notional nozzle model coupled to the jet model for underexpanded jets. The values of some key model parameters were determined by comparisons with the experimental data. Finally, the coupled model was also validated against hydrogen concentrations measurements for 100 and 200 bar hydrogen jets with the predictions agreeing well with data in the literature.

  17. Research Data Acquired in World-Class, 60-atm Subsonic Combustion Rig

    NASA Technical Reports Server (NTRS)

    Lee, Chi-Ming; Wey, Changlie

    1999-01-01

    NASA Lewis Research Center's new, world-class, 60-atmosphere (atm) combustor research facility, the Advanced Subsonic Combustion Rig (ASCR), is in operation and producing highly unique research data. Specifically, data were acquired at high pressures and temperatures representative of future subsonic engines from a fundamental flametube configuration with an advanced fuel injector. The data acquired include exhaust emissions as well as pressure and temperature distributions. Results to date represent an improved understanding of nitrous oxide (NOx) formation at high pressures and temperatures and include an NOx emissions reduction greater than 70 percent with an advanced fuel injector at operating pressures to 800 pounds per square inch absolute (psia). ASCR research is an integral part of the Advanced Subsonic Technology (AST) Propulsion Program. This program is developing critical low-emission combustion technology that will result in the next generation of gas turbine engines producing 50 to 70 percent less NOx emissions in comparison to 1996 International Civil Aviation Organization (ICAO) limits. The results to date indicate that the AST low-emission combustor goals of reducing NOx emissions by 50 to 70 percent are feasible. U.S. gas turbine manufacturers have started testing the low-emissions combustors at the ASCR. This collaborative testing will enable the industry to develop low-emission combustors at the high pressure and temperature conditions of future subsonic engines. The first stage of the flametube testing has been implemented. Four GE Aircraft Engines low-emissions fuel injector concepts, three Pratt & Whitney concepts, and two Allison concepts have been tested at Lewis ASCR facility. Subsequently, the flametube was removed from the test stand, and the sector combustor was installed. The testing of low emissions sector has begun. Low-emission combustors developed as a result of ASCR research will enable U.S. engine manufacturers to compete on a

  18. The incorporation of plotting capability into the Unified Subsonic Supersonic Aerodynamic Analysis program, version B

    NASA Technical Reports Server (NTRS)

    Winter, O. A.

    1980-01-01

    The B01 version of the United Subsonic Supersonic Aerodynamic Analysis program is the result of numerous modifications and additions made to the B00 version. These modifications and additions affect the program input, its computational options, the code readability, and the overlay structure. The following are described: (1) the revised input; (2) the plotting overlay programs which were also modified, and their associated subroutines, (3) the auxillary files used by the program, the revised output data; and (4) the program overlay structure.

  19. The drag force on a subsonic projectile in a fluid complex plasma

    SciTech Connect

    Ivlev, A. V.; Zhukhovitskii, D. I.

    2012-09-15

    The incompressible Navier-Stokes equation is employed to describe a subsonic particle flow induced in complex plasmas by a moving projectile. Drag forces acting on the projectile in different flow regimes are calculated. It is shown that, along with the regular neutral gas drag, there is an additional force exerted on the projectile due to dissipation in the surrounding particle fluid. This additional force provides significant contribution to the total drag.

  20. Fundamental Studies of SUBSONIC and Transonic Flow Separation. Part 3. Third Phase Summary Report

    DTIC Science & Technology

    1979-10-01

    Direction in the Immediate Downstream of a Normal Shock Measured by Cone Probe . . . . . . . . . . . . . . . . 34 Typical Oil Flow Over Ramp Compression...Shallow- Cavity Model . . . . . . . . . . . . . 82 A Typical Oil Flow Pattern Upstream and on the Forward-Facing Step .... 83 Variation of Base...Unit Reynolds number range Reynolds number range based on length from leading edge Maximum temperature drop (°C) for i0 sec. Subsonic flow

  1. NASA Subsonic Rotary Wing Project-Multidisciplinary Analysis and Technology Development: Overview

    NASA Technical Reports Server (NTRS)

    Yamauchi, Gloria K.

    2009-01-01

    This slide presentation reviews the objectives of the Multidisciplinary Analysis and Technology Development (MDATD) in the Subsonic Rotary Wing project. The objectives are to integrate technologies and analyses to enable advanced rotorcraft and provide a roadmap to guide Level 1 and 2 research. The MDATD objectives will be met by conducting assessments of advanced technology benefits, developing new or enhanced design tools, and integrating Level 2 discipline technologies to develop and enable system-level analyses and demonstrations.

  2. Automatic computation of Euler-marching and subsonic grids for wing-fuselage configurations

    NASA Technical Reports Server (NTRS)

    Barger, Raymond L.; Adams, Mary S.; Krishnan, Ramki R.

    1994-01-01

    Algebraic procedures are described for the automatic generation of structured, single-block flow computation grids for relatively simple configurations (wing, fuselage, and fin). For supersonic flows, a quasi two-dimensional grid for Euler-marching codes is developed, and some sample results in graphical form are included. A type of grid for subsonic flow calculation is also described. The techniques are algebraic and are based on a generalization of the method of transfinite interpolation.

  3. Influence matrix program for aerodynamic lifting surface theory. [in subsonic flows

    NASA Technical Reports Server (NTRS)

    Medan, R. T.; Ray, K. S.

    1973-01-01

    A users manual is described for a USA FORTRAN 4 computer program which computes an aerodynamic influence matrix and is one of several computer programs used to analyze lifting, thin wings in steady, subsonic flow according to a kernel function method lifting surface theory. The most significant features of the program are that it can treat unsymmetrical wings, control points can be placed on the leading and/or trailing edges, and a stable, efficient algorithm is used to compute the influence matrix.

  4. Unsteady effects of a control surface in two dimensional subsonic and transonic flow

    NASA Technical Reports Server (NTRS)

    Grenon, R.; Desopper, A.; Sides, J.

    1980-01-01

    The experimental results of steady and unsteady pressure measurements, carried out in subsonic and transonic flow on a 16 percent relative thickness supercritical aerofoil, equipped with a trailing edge flap involving 25 percent of the chord, in a sinusoidal motion are given. These experimental results are compared with those obtained by various methods of steady and unsteady inviscid flow calculations. Some calculation results in which viscous effects have been taken into account, for both steady and unsteady flows, are also presented.

  5. Carpal tunnel syndrome

    PubMed Central

    2014-01-01

    Introduction Carpal tunnel syndrome is a collection of clinical symptoms and signs caused by compression of the median nerve within the carpal tunnel. However, the severity of symptoms and signs does not often correlate well with the extent of nerve compression. Methods and outcomes We conducted a systematic review and aimed to answer the following clinical questions: What are the effects of drug treatments, non-drug treatments, and surgical treatments for carpal tunnel syndrome? We searched: Medline, Embase, The Cochrane Library, and other important databases up to October 2013 (Clinical Evidence reviews are updated periodically; please check our website for the most up-to-date version of this review). We included harms alerts from relevant organisations such as the US Food and Drug Administration (FDA) and the UK Medicines and Healthcare products Regulatory Agency (MHRA). Results We found 33 studies that met our inclusion criteria. We performed a GRADE evaluation of the quality of evidence for interventions. Conclusions In this systematic review we present information relating to the effectiveness and safety of the following interventions: carpal tunnel release surgery (open and endoscopic), diuretics, local corticosteroids injection, non-steroidal anti-inflammatory drugs (NSAIDs), therapeutic ultrasound, and wrist splints.

  6. Carpal tunnel syndrome

    PubMed Central

    2010-01-01

    Introduction Carpal tunnel syndrome is a neuropathy caused by compression of the median nerve within the carpal tunnel. However, the severity of symptoms and signs does not often correlate well with the extent of nerve damage. Methods and outcomes We conducted a systematic review and aimed to answer the following clinical questions: What are the effects of drug treatments, non-drug treatments, surgical treatments, and postoperative treatments for carpal tunnel syndrome? We searched: Medline, Embase, The Cochrane Library, and other important databases up to March 2009 (Clinical Evidence reviews are updated periodically; please check our website for the most up-to-date version of this review). We included harms alerts from relevant organisations such as the US Food and Drug Administration (FDA) and the UK Medicines and Healthcare products Regulatory Agency (MHRA). Results We found 53 systematic reviews, RCTs, or observational studies that met our inclusion criteria. We performed a GRADE evaluation of the quality of evidence for interventions. Conclusions In this systematic review we present information relating to the effectiveness and safety of the following interventions: acupuncture, carpal tunnel release surgery (open and endoscopic), diuretics, internal neurolysis, local and systemic corticosteroids, massage therapy, nerve and tendon gliding exercises, non-steroidal anti-inflammatory drugs (NSAIDs), pyridoxine, therapeutic ultrasound, and wrist splints. PMID:21718565

  7. Wind Tunnel Balances

    NASA Technical Reports Server (NTRS)

    Warner, Edward P; Norton, F H

    1920-01-01

    Report embodies a description of the balance designed and constructed for the use of the National Advisory Committee for Aeronautics at Langley Field, and also deals with the theory of sensitivity of balances and with the errors to which wind tunnel balances of various types are subject.

  8. Tunneling path toward spintronics

    NASA Astrophysics Data System (ADS)

    Miao, Guo-Xing; Münzenberg, Markus; Moodera, Jagadeesh S.

    2011-03-01

    The phenomenon of quantum tunneling, which was discovered almost a century ago, has led to many subsequent discoveries. One such discovery, spin polarized tunneling, was made 40 years ago by Robert Meservey and Paul Tedrow (Tedrow and Meservey 1971 Phys. Rev. Lett. 26 192), and it has resulted in many fundamental observations and opened up an entirely new field of study. Until the mid-1990s, this field developed at a steady, low rate, after which a huge increase in activity suddenly occurred as a result of the unraveling of successful spin tunneling between two ferromagnets. In the past 15 years, several thousands of papers related to spin polarized tunneling and transport have been published, making this topic one of the hottest areas in condensed matter physics from both fundamental science and applications viewpoints. Many review papers and book chapters have been written in the past decade on this subject. This paper is not exhaustive by any means; rather, the emphases are on recent progress, technological developments and informing the reader about the current direction in which this topic is moving.

  9. Tunnelling with wormhole creation

    SciTech Connect

    Ansoldi, S.; Tanaka, T.

    2015-03-15

    The description of quantum tunnelling in the presence of gravity shows subtleties in some cases. We discuss wormhole production in the context of the spherically symmetric thin-shell approximation. By presenting a fully consistent treatment based on canonical quantization, we solve a controversy present in the literature.

  10. Full Scale Tunnel model

    NASA Technical Reports Server (NTRS)

    1929-01-01

    Interior view of Full-Scale Tunnel (FST) model. (Small human figures have been added for scale.) On June 26, 1929, Elton W. Miller wrote to George W. Lewis proposing the construction of a model of the full-scale tunnel . 'The excellent energy ratio obtained in the new wind tunnel of the California Institute of Technology suggests that before proceeding with our full scale tunnel design, we ought to investigate the effect on energy ratio of such factors as: 1. small included angle for the exit cone; 2. carefully designed return passages of circular section as far as possible, without sudden changes in cross sections; 3. tightness of walls. It is believed that much useful information can be obtained by building a model of about 1/16 scale, that is, having a closed throat of 2 ft. by 4 ft. The outside dimensions would be about 12 ft. by 25 ft. in plan and the height 4 ft. Two propellers will be required about 28 in. in diameter, each to be driven by direct current motor at a maximum speed of 4500 R.P.M. Provision can be made for altering the length of certain portions, particularly the exit cone, and possibly for the application of boundary layer control in order to effect satisfactory air flow.

  11. Dry wind tunnel system

    NASA Technical Reports Server (NTRS)

    Chen, Ping-Chih (Inventor)

    2013-01-01

    This invention is a ground flutter testing system without a wind tunnel, called Dry Wind Tunnel (DWT) System. The DWT system consists of a Ground Vibration Test (GVT) hardware system, a multiple input multiple output (MIMO) force controller software, and a real-time unsteady aerodynamic force generation software, that is developed from an aerodynamic reduced order model (ROM). The ground flutter test using the DWT System operates on a real structural model, therefore no scaled-down structural model, which is required by the conventional wind tunnel flutter test, is involved. Furthermore, the impact of the structural nonlinearities on the aeroelastic stability can be included automatically. Moreover, the aeroservoelastic characteristics of the aircraft can be easily measured by simply including the flight control system in-the-loop. In addition, the unsteady aerodynamics generated computationally is interference-free from the wind tunnel walls. Finally, the DWT System can be conveniently and inexpensively carried out as a post GVT test with the same hardware, only with some possible rearrangement of the shakers and the inclusion of additional sensors.

  12. The Mystery Tunnel

    ERIC Educational Resources Information Center

    McCormack, Alan J.

    1974-01-01

    Describes a mystery tunnel, constructed by teachers, which provides a variety of non-visual, sensory experiences for children as they crawl through it. It is designed to help primary children develop basic abilities to use their own senses to better observe, discriminate among observations, and describe their own perceptions accurately. (JR)

  13. The Channel Tunnel

    NASA Technical Reports Server (NTRS)

    2006-01-01

    The Channel Tunnel is a 50.5 km-long rail tunnel beneath the English Channel at the Straits of Dover. It connects Dover, Kent in England with Calais, northern France. The undersea section of the tunnel is unsurpassed in length in the world. A proposal for a Channel tunnel was first put forward by a French engineer in 1802. In 1881, a first attempt was made at boring a tunnel from the English side; the work was halted after 800 m. Again in 1922, English workers started boring a tunnel, and advanced 120 m before it too was halted for political reasons. The most recent attempt was begun in 1987, and the tunnel was officially opened in 1994. At completion it was estimated that the project cost around $18 billion. It has been operating at a significant loss since its opening, despite trips by over 7 million passengers per year on the Eurostar train, and over 3 million vehicles per year.

    With its 14 spectral bands from the visible to the thermal infrared wavelength region, and its high spatial resolution of 15 to 90 meters (about 50 to 300 feet), ASTER images Earth to map and monitor the changing surface of our planet.

    ASTER is one of five Earth-observing instruments launched December 18, 1999, on NASA's Terra satellite. The instrument was built by Japan's Ministry of Economy, Trade and Industry. A joint U.S./Japan science team is responsible for validation and calibration of the instrument and the data products.

    The broad spectral coverage and high spectral resolution of ASTER provides scientists in numerous disciplines with critical information for surface mapping, and monitoring of dynamic conditions and temporal change. Example applications are: monitoring glacial advances and retreats; monitoring potentially active volcanoes; identifying crop stress; determining cloud morphology and physical properties; wetlands evaluation; thermal pollution monitoring; coral reef degradation; surface temperature mapping of soils and geology; and measuring

  14. Subsonic flight test evaluation of a performance seeking control algorithm on an F-15 airplane

    NASA Technical Reports Server (NTRS)

    Gilyard, Glenn B.; Orme, John S.

    1992-01-01

    The subsonic flight test evaluation phase of the NASA F-15 (powered by F 100 engines) performance seeking control program was completed for single-engine operation at part- and military-power settings. The subsonic performance seeking control algorithm optimizes the quasi-steady-state performance of the propulsion system for three modes of operation. The minimum fuel flow mode minimizes fuel consumption. The minimum thrust mode maximizes thrust at military power. Decreases in thrust-specific fuel consumption of 1 to 2 percent were measured in the minimum fuel flow mode; these fuel savings are significant, especially for supersonic cruise aircraft. Decreases of up to approximately 100 degree R in fan turbine inlet temperature were measured in the minimum temperature mode. Temperature reductions of this magnitude would more than double turbine life if inlet temperature was the only life factor. Measured thrust increases of up to approximately 15 percent in the maximum thrust mode cause substantial increases in aircraft acceleration. The system dynamics of the closed-loop algorithm operation were good. The subsonic flight phase has validated the performance seeking control technology, which can significantly benefit the next generation of fighter and transport aircraft.

  15. Power-by-Wire Development and Demonstration for Subsonic Civil Transport

    NASA Technical Reports Server (NTRS)

    1996-01-01

    During the last decade, three significant studies by the Lockheed Martin Corporation, the NASA Lewis Research Center, and McDonnell Douglas Corporation have clearly shown operational, weight, and cost advantages for commercial subsonic transport aircraft that use all-electric or more-electric technologies in the secondary electric power systems. Even though these studies were completed on different aircraft, used different criteria, and applied a variety of technologies, all three have shown large benefits to the aircraft industry and to the nation's competitive position. The Power-by-Wire (PBW) program is part of the highly reliable Fly-By-Light/Power-By-Wire (FBL/PBW) Technology Program, whose goal is to develop the technology base for confident application of integrated FBL/PBW systems for transport aircraft. This program is part of the NASA aeronautics strategic thrust in subsonic aircraft/national airspace (Thrust 1) to "develop selected high-leverage technologies and explore new means to ensure the competitiveness of U.S. subsonic aircraft and to enhance the safety and productivity of the national aviation system" (The Aeronautics Strategic Plan). Specifically, this program is an initiative under Thrust 1, Key Objective 2, to "develop, in cooperation with U.S. industry, selected high-payoff technologies that can enable significant improvements in aircraft efficiency and cost."

  16. Nonlinear evolution of subsonic and supersonic disturbances on a compressible free shear layer

    NASA Technical Reports Server (NTRS)

    Leib, S. J.

    1991-01-01

    The effects of a nonlinear-nonequilibrium-viscous critical layer on the spatial evolution of subsonic and supersonic instability modes on a compressible free shear layer is considered. It is shown that the instability wave amplitude is governed by an integrodifferential equation with cubic-type nonlinearity. Numerical and asymptotic solutions to this equation show that the amplitude either ends in a singularity at a finite downstream distance or reaches an equilibrium value, depending on the Prandtl number, viscosity law, viscous parameter and a real parameter which is determined by the linear inviscid stability theory. A necessary condition for the existence of the equilibrium solution is derived, and whether or not this condition is met is determined numerically for a wide range of physical parameters including both subsonic and supersonic disturbances. it is found that no equilibrium solution exists for the subsonic modes unless the temperature ratio of the low-to-high-speed streams exceeds a critical value, while equilibrium solutions for the most rapidly growing supersonic mode exist over most of the parameter range examined.

  17. A New Density Variance-Mach Number Relation for Subsonic and Supersonic Isothermal Turbulence

    NASA Astrophysics Data System (ADS)

    Konstandin, L.; Girichidis, P.; Federrath, C.; Klessen, R. S.

    2012-12-01

    The probability density function of the gas density in subsonic and supersonic, isothermal, driven turbulence is analyzed using a systematic set of hydrodynamical grid simulations with resolutions of up to 10243 cells. We perform a series of numerical experiments with root-mean-square (rms) Mach number {M} ranging from the nearly incompressible, subsonic ( {M}=0.1) to the highly compressible, supersonic ( {M}=15) regime. We study the influence of two extreme cases for the driving mechanism by applying a purely solenoidal (divergence-free) and a purely compressive (curl-free) forcing field to drive the turbulence. We find that our measurements fit the linear relation between the rms Mach number and the standard deviation (std. dev.) of the density distribution in a wide range of Mach numbers, where the proportionality constant depends on the type of forcing. In addition, we propose a new linear relation between the std. dev. of the density distribution σρ and that of the velocity in compressible modes, i.e., the compressible component of the rms Mach number, {M}_{{comp}}. In this relation the influence of the forcing is significantly reduced, suggesting a linear relation between σρ and {M}_{{comp}}, independent of the forcing, and ranging from the subsonic to the supersonic regime.

  18. The Performance of a Subsonic Diffuser Designed for High Speed Turbojet-Propelled Flight

    NASA Technical Reports Server (NTRS)

    Biesiadny, Thomas J. (Technical Monitor); Wendt, Bruce J.

    2004-01-01

    An initial-phase subsonic diffuser has been designed for the turbojet flowpath of the hypersonic x43B flight demonstrator vehicle. The diffuser fit into a proposed mixed-compression supersonic inlet system and featured a cross-sectional shape transitioning flowpath (high aspect ratio rectangular throat-to-circular engine face) and a centerline offset. This subsonic diffuser has been fabricated and tested at the W1B internal flow facility at NASA Glenn Research Center. At an operating throat Mach number of 0.79, baseline Pitot pressure recovery was found to be just under 0.9, and DH distortion intensity was about 0.4 percent. The diffuser internal flow stagnated, but did not separate on the offset surface of this initial-phase subsonic diffuser. Small improvements in recovery (+0.4 percent) and DH distortion (-32 percent) were obtained from using vane vortex generator flow control applied just downstream of the diffuser throat. The optimum vortex generator array patterns produced inflow boundary layer divergence (local downwash) on the offset surface centerline of the diffuser, and an inflow boundary layer convergence (local upwash) on the centerline of the opposite surface.

  19. Effect of winglets on a first-generation jet transport wing. 5: Stability characteristics of a full-span wing with a generalized fuselage at high subsonic speeds

    NASA Technical Reports Server (NTRS)

    Jacobs, P. F.

    1978-01-01

    The effects of winglets on the static aerodynamic stability characteristics of a KC-135A jet transport model at high subsonic speeds are presented. The investigation was conducted in the Langley 8 foot transonic pressure tunnel using 0.035-scale wing panels mounted on a generalized research fuselage. Data were taken over a Mach number range from 0.50 to 0.95 at angles of attack ranging from -12 deg to 20 deg and sideslip angles of 0 deg, 5 deg, and -5 deg. The model was tested at two Reynolds number ranges to achieve a wide angle of attack range and to determine the effect of Reynolds number on stability. Results indicate that adding the winglets to the basic wing configuration produces small increases in both lateral and longitudinal aerodynamic stability and that the model stability increases slightly with Reynolds number. The winglets do increase the wing bending moments slightly, but the buffet onset characteristics of the model are not affected by the winglets.

  20. Subsonic longitudinal and lateral-directional characteristics of a forward-swept-wing fighter configuration at angles of attack up to 47 deg

    NASA Technical Reports Server (NTRS)

    Mann, Michael J.; Huffman, Jarrett K.; Fox, Charles H., Jr.

    1987-01-01

    Subsonic lateral-direction and longitudinal characteristics of a forward-swept-wing fighter configuration were examined in wind-tunnel tests at Mach numbers of 0.2 and 0.5 for angles of attack from -7 to 47 deg. and over a sidelslip range of +/- 15 deg. The effects of a canard, strakes, vertical tail, and leading- and trailing-edge flaps are examined. The canard and strakes both reduce asymmetric moments and side forces at zero sideslip for angles of attack up to about 30 deg. The canard has a small influence on lateral-directional stability; however, strakes produce a substantial reduction in lateral stability for angles of attack greater than about 20 deg. The vertical tail improves directional stability for angles of attack up to 30 deg. Deflection of the leading-edge flap to 20 deg. at high angles of attack on the strake and canard configurations degrades lateral and directional stability. Deflection of the trailing-edge flap to 20 deg. on the canard configuration generally increases lateral and directional stability at high angles of attack. Leading- and trailing-edge flaps on the wing-body and canard configurations are effective for increased lift only for angles of attack up to about 40 deg. The leading-edge flap remains effective on the strake configuration over the entire angle-of-attack range tested.

  1. Low-Subsonic Measurements of the Static and Oscillatory Lateral Stability Derivatives of a Sweptback-Wing Airplane Configuration at Angles of Attack from -10 to 90 deg

    NASA Technical Reports Server (NTRS)

    Hewes, Donald E.

    1959-01-01

    An investigation has been conducted in the Langley free-flight tunnel at low-subsonic speeds to provide some basic information on the stability and control characteristics in the high angle-of-attack range of an airplane configuration typical of current design trends. The investigation consisted of static- and dynamic-force tests over an angle-of- attack range from -10 to 90 deg. The dynamic-force tests, which consisted of both linear- and rotary-oscillation tests, were conducted at values of the reduced-frequency parameter k of 0.10, 0.15, and 0.20. The configuration was directionally unstable for all angles of attack above about 15 deg but maintained positive effective dihedral, control effectiveness, and damping in roll and yaw over most of the angle-of-attack range tested. The effects of frequency on the oscillatory stability derivatives were found to be generally small, but in a few cases the effects were relatively large.

  2. Flight-Determined Subsonic Lift and Drag Characteristics of Seven Lifting-Body and Wing-Body Reentry Vehicle Configurations With Truncated Bases

    NASA Technical Reports Server (NTRS)

    Saltzman, Edwin J.; Wang, K. Charles; Iliff, Kenneth W.

    1999-01-01

    This paper examines flight-measured subsonic lift and drag characteristics of seven lifting-body and wing-body reentry vehicle configurations with truncated bases. The seven vehicles are the full-scale M2-F1, M2-F2, HL-10, X-24A, X-24B, and X-15 vehicles and the Space Shuttle prototype. Lift and drag data of the various vehicles are assembled under aerodynamic performance parameters and presented in several analytical and graphical formats. These formats unify the data and allow a greater understanding than studying the vehicles individually allows. Lift-curve slope data are studied with respect to aspect ratio and related to generic wind-tunnel model data and to theory for low-aspect-ratio planforms. The proper definition of reference area was critical for understanding and comparing the lift data. The drag components studied include minimum drag coefficient, lift-related drag, maximum lift-to-drag ratio, and, where available, base pressure coefficients. The effects of fineness ratio on forebody drag were also considered. The influence of forebody drag on afterbody (base) drag at low lift is shown to be related to Hoerner's compilation for body, airfoil, nacelle, and canopy drag. These analyses are intended to provide a useful analytical framework with which to compare and evaluate new vehicle configurations of the same generic family.

  3. Investigation of the Low-Subsonic Stability and Control Characteristics of a Free-Flying Model of a Thick 70 deg Delta Reentry Configuration

    NASA Technical Reports Server (NTRS)

    Paulson, John W.; Shanks, Robert E.

    1961-01-01

    An investigation of the low-subsonic flight characteristics of a thick 70 deg delta reentry configuration having a diamond cross section has been made in the Langley full-scale tunnel over an angle-of-attack range from 20 to 45 deg. Flight tests were also made at angles of attack near maximum lift (alpha = 40 deg) with a radio-controlled model dropped from a helicopter. Static and dynamic force tests were made over an angle-of-attack range from 0 to 90 deg. The longitudinal stability and control characteristics were considered satisfactory when the model had positive static longitudinal stability. It was possible to fly the model with a small amount of static instability, but the longitudinal characteristics were considered unsatisfactory in this condition. At angles of attack above the stall the model developed a large, constant-amplitude pitching oscillation. The lateral stability characteristics were considered to be only fair at angles of attack from about 20 to 35 deg because of a lightly damped Dutch roll oscillation. At higher angles of attack the oscillation was well damped and the lateral stability was generally satisfactory. The Dutch roll damping at the lower angles of attack was increased to satisfactory values by means of a simple rate-type roll damper. The lateral control characteristics were generally satisfactory throughout the angle- of-attack range, but there was some deterioration in aileron effectiveness in the high angle-of-attack range due mainly to a large increase in damping in roll.

  4. Exploratory study to induce fan noise in the test section of the NASA Langley full-scale wind tunnel

    NASA Technical Reports Server (NTRS)

    Ver, I. L.; Hayden, R. E.; Myles, M. M.; Murray, B. E.

    1975-01-01

    Measures to reduce the intensity of fan noise in the NASA Langley 30 ft x 60 ft subsonic wind tunnel were sought. Measurements were first performed to document existing aerodynamic and acoustic conditions. The purpose of these experiments was to (1) obtain the transfer function between the sound power output of the fan and the sound pressure on the test platform, (2) evaluate the sound attenuation around the tunnel circuit, (3) measure simultaneously the flow profile and the turbulence spectrum of the inflow to the fan and the noise on the test platform, and (4) perform flow observations and identify secondary noise sources. Subsequently, these data were used to predict (1) the relative contribution of the major aerodynamic parameters to total fan noise and (2) the effect of placing a dissipative silencer in the collector duct upstream of the fan. Promising noise control measures were identified and recommendations were made on how to evaluate them.

  5. Combustion of Various Highly Reactive Fuels in a 3.84- by 10-inch Mach 2 Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Allen, Harrison, Jr.; Fletcher, Edward A.

    1959-01-01

    The following fuels and fuel combinations injected from the top wall of a Mach 2 wind tunnel were successfully burned and gave associated pressure rises: aluminum borohydride, pentaborane, mixtures containing up to 41 percent JP-4 fuel in aluminum borohydride, tandem injections of aluminum borohydride, tandem injections of JP-4 fuel and aluminum borohydride, trimethyl aluminum with water injections, and diethyl aluminum hydride with water injections. The following fuels could not be ignited at the tunnel conditions (static pressure, 5.6 in. Hg; static temperature, -148 F): trimethylborane, triethylborane, propylpentaborane, ethyl- decaborane, and vinylsilane. Studies in which the heated region was probed by water injections indicated that the flow downstream of the flame front is subsonic and recirculating.

  6. Tunnel boring machine

    SciTech Connect

    Snyder, L. L.

    1985-07-09

    A tunnel boring machine for controlled boring of a curvilinear tunnel including a rotating cutter wheel mounted on the forward end of a thrust cylinder assembly having a central longitudinal axis aligned with the cutter wheel axis of rotation; the thrust cylinder assembly comprising a cylinder barrel and an extendable and retractable thrust arm received therein. An anchoring assembly is pivotally attached to the rear end of the cylinder barrel for anchoring the machine during a cutting stroke and providing a rear end pivot axis during curved cutting strokes. A pair of laterally extending, extendable and retractable arms are fixedly mounted at a forward portion of the cylinder barrel for providing lateral displacement in a laterally curved cutting mode and for anchoring the machine between cutting strokes and during straight line boring. Forward and rear transverse displacement and support assemblies are provided to facilitate cutting in a transversely curved cutting mode and to facilitate machine movement between cutting strokes.

  7. Uncooled tunneling infrared sensor

    NASA Technical Reports Server (NTRS)

    Kenny, Thomas W. (Inventor); Kaiser, William J. (Inventor); Podosek, Judith A. (Inventor); Vote, Erika C. (Inventor); Rockstad, Howard K. (Inventor); Reynolds, Joseph K. (Inventor)

    1994-01-01

    An uncooled infrared tunneling sensor in which the only moving part is a diaphragm which is deflected into contact with a micromachined silicon tip electrode prepared by a novel lithographic process. Similarly prepared deflection electrodes employ electrostatic force to control the deflection of a silicon nitride, flat diaphragm membrane. The diaphragm exhibits a high resonant frequency which reduces the sensor's sensitivity to vibration. A high bandwidth feedback circuit controls the tunneling current by adjusting the deflection voltage to maintain a constant deflection of the membrane which would otherwise change deflection depending upon incident infrared radiation. The resulting infrared sensor will meet or exceed the performance of all other broadband, uncooled, infrared sensors and can be miniaturized to pixel dimensions smaller than 100 .mu.m. The technology is readily implemented as a small-format linear array suitable for commercial and spacecraft applications.

  8. Possibility of hyperbolic tunneling

    SciTech Connect

    Lobo, Francisco S. N.; Mimoso, Jose P.

    2010-08-15

    Traversable wormholes are primarily useful as 'gedanken experiments' and as a theoretician's probe of the foundations of general relativity. In this work, we analyze the possibility of having tunnels in a hyperbolic spacetime. We obtain exact solutions of static and pseudo-spherically symmetric spacetime tunnels by adding exotic matter to a vacuum solution referred to as a degenerate solution of class A. The physical properties and characteristics of these intriguing solutions are explored, and through the mathematics of embedding it is shown that particular constraints are placed on the shape function, that differ significantly from the Morris-Thorne wormhole. In particular, it is shown that the energy density is always negative, and the radial pressure is positive, at the throat, contrary to the Morris-Thorne counterpart. Specific solutions are also presented by considering several equations of state, and by imposing restricted choices for the shape function or the redshift function.

  9. Tunnel magnetoresistance of diamondoids

    NASA Astrophysics Data System (ADS)

    Matsuura, Yukihito

    2016-10-01

    Tunnel magnetoresistance (TMR) of diamondoids has been predicted by first principles density functional theory. Diamantane was used as a basic molecular proxy for diamondoids because hydrogen atoms in the apical position are easily substituted for a thiol group. The pristine diamantane exhibited a low TMR ratio of 7%, and boron-substitution considerably decreased the TMR ratio. Conversely, nitrogen-substitution enhanced the TMR ratio by up to 20%. Heteroatom-substitution changes the tunneling probabilities by varying the molecular bond lengths. Furthermore, when the spins of the electrodes are parallel, the heteroatoms resulted in transmittance probabilities at an energy range near the Fermi level. Consequently, heteroatom-substitution can control the TMR ratios of diamondoids very well.

  10. Tarsal tunnel syndrome.

    PubMed

    Gould, John S

    2011-06-01

    Tarsal tunnel syndrome, unlike its similar sounding counterpart in the hand, is a significantly misunderstood clinical entity. Confusion concerning the anatomy involved, the presenting symptomatology, the appropriateness and significance of various diagnostic tests, conservative and surgical management, and, finally, the variability of reported results of surgical intervention attests to the lack of consensus surrounding this condition. The terminology involved in various diagnoses for chronic heel pain is also a hodgepodge of poorly understood entities.

  11. Full Scale Tunnel (FST)

    NASA Technical Reports Server (NTRS)

    1930-01-01

    Installation of Full Scale Tunnel (FST) power plant. Virginia Public Service Company could not supply adequate electricity to run the wind tunnels being built at Langley. (The Propeller Research Tunnel was powered by two submarine diesel engines.) This led to the consideration of a number of different ideas for generating electric power to drive the fan motors in the FST. The main proposition involved two 3000 hp and two 1000 hp diesel engines with directly connected generators. Another, proposition suggested 30 Liberty motors driving 600 hp DC generators in pairs. For a month, engineers at Langley were hopeful they could secure additional diesel engines from decommissioned Navy T-boats but the Navy could not offer a firm commitment regarding the future status of the submarines. By mid-December 1929, Virginia Public Service Company had agreed to supply service to the field at the north end of the King Street Bridge connecting Hampton and Langley Field. Thus, new plans for FST powerplant and motors were made. Smith DeFrance described the motors in NACA TR No. 459: 'The most commonly used power plant for operating a wind tunnel is a direct-current motor and motor-generator set with Ward Leonard control system. For the FST it was found that alternating current slip-ring induction motors, together with satisfactory control equipment, could be purchased for approximately 30 percent less than the direct-current equipment. Two 4000-horsepower slip-ring induction motors with 24 steps of speed between 75 and 300 r.p.m. were therefore installed.'

  12. Diffraction as tunneling

    NASA Technical Reports Server (NTRS)

    Nussenzveig, H. M.; Wiscombe, W. J.

    1987-01-01

    A new approximation to the short-wavelength scattering amplitude from an impenetrable sphere is presented. It is uniform in the scattering angle and it is more accurate than previously known approximations (including Fock's theory of diffraction) by up to several orders of magnitude. It remains valid in the transition to long-wavelength scattering. It leads to a new physical picture of diffraction, as tunneling through an inertial barrier.

  13. Resonant Tunneling Spin Pump

    NASA Technical Reports Server (NTRS)

    Ting, David Z.

    2007-01-01

    The resonant tunneling spin pump is a proposed semiconductor device that would generate spin-polarized electron currents. The resonant tunneling spin pump would be a purely electrical device in the sense that it would not contain any magnetic material and would not rely on an applied magnetic field. Also, unlike prior sources of spin-polarized electron currents, the proposed device would not depend on a source of circularly polarized light. The proposed semiconductor electron-spin filters would exploit the Rashba effect, which can induce energy splitting in what would otherwise be degenerate quantum states, caused by a spin-orbit interaction in conjunction with a structural-inversion asymmetry in the presence of interfacial electric fields in a semiconductor heterostructure. The magnitude of the energy split is proportional to the electron wave number. Theoretical studies have suggested the possibility of devices in which electron energy states would be split by the Rashba effect and spin-polarized currents would be extracted by resonant quantum-mechanical tunneling.

  14. Evaluating tunnel kiln performance

    SciTech Connect

    O`Connor, K.R.; Carty, W.M.; Ninos, N.J.

    1997-08-01

    Process improvements in the production of whitewares provide the potential for substantial savings for manufacturers. A typical whiteware manufacturer incurs an annual defective product loss of {approximately}$20 million when accounting for raw materials, energy, labor and waste disposal. Reduction in defective product loss of 1% could result in a savings in excess of $1 million annually. This study was designed to establish benchmarks for two conventional tunnel kilns used to bisque-fire dinnerware at Buffalo China Inc. (Buffalo, NY). The benchmark was established by assessing the current conditions and variability of the two tunnel kilns as a function of the fracture strength of sample bars that were made from production body. Sample bars were fired in multiple locations in both kilns to assess the conditions and variability of firing within each kiln. Comparison of strength results between the two kilns also was assessed. These comparisons were accomplished through applied statistical analysis, wherein significant statistical variations were identified and isolated for both tunnel kilns. The statistical methods and tools used in this analysis are readily accessible to manufacturers, thus allowing implementation of similar analysis, or benchmarking, in-house.

  15. Smart tunnel: Docking mechanism

    NASA Technical Reports Server (NTRS)

    Schliesing, John A. (Inventor); Edenborough, Kevin L. (Inventor)

    1989-01-01

    A docking mechanism is presented for the docking of a space vehicle to a space station comprising a flexible tunnel frame structure which is deployable from the space station. The tunnel structure comprises a plurality of series connected frame sections, one end section of which is attached to the space station and the other end attached to a docking module of a configuration adapted for docking in the payload bay of the space vehicle. The docking module is provided with trunnions, adapted for latching engagement with latches installed in the vehicle payload bay and with hatch means connectable to a hatch of the crew cabin of the space vehicle. Each frame section comprises a pair of spaced ring members, interconnected by actuator-attenuator devices which are individually controllable by an automatic control means to impart relative movement of one ring member to the other in six degrees of freedom of motion. The control means includes computer logic responsive to sensor signals of range and attitude information, capture latch condition, structural loads, and actuator stroke for generating commands to the onboard flight control system and the individual actuator-attenuators to deploy the tunnel to effect a coupling with the space vehicle and space station after coupling. A tubular fluid-impervious liner, preferably fabric, is disposed through the frame sections of a size sufficient to accommodate the passage of personnel and cargo.

  16. Research and test facilities

    NASA Technical Reports Server (NTRS)

    1993-01-01

    A description is given of each of the following Langley research and test facilities: 0.3-Meter Transonic Cryogenic Tunnel, 7-by 10-Foot High Speed Tunnel, 8-Foot Transonic Pressure Tunnel, 13-Inch Magnetic Suspension & Balance System, 14-by 22-Foot Subsonic Tunnel, 16-Foot Transonic Tunnel, 16-by 24-Inch Water Tunnel, 20-Foot Vertical Spin Tunnel, 30-by 60-Foot Wind Tunnel, Advanced Civil Transport Simulator (ACTS), Advanced Technology Research Laboratory, Aerospace Controls Research Laboratory (ACRL), Aerothermal Loads Complex, Aircraft Landing Dynamics Facility (ALDF), Avionics Integration Research Laboratory, Basic Aerodynamics Research Tunnel (BART), Compact Range Test Facility, Differential Maneuvering Simulator (DMS), Enhanced/Synthetic Vision & Spatial Displays Laboratory, Experimental Test Range (ETR) Flight Research Facility, General Aviation Simulator (GAS), High Intensity Radiated Fields Facility, Human Engineering Methods Laboratory, Hypersonic Facilities Complex, Impact Dynamics Research Facility, Jet Noise Laboratory & Anechoic Jet Facility, Light Alloy Laboratory, Low Frequency Antenna Test Facility, Low Turbulence Pressure Tunnel, Mechanics of Metals Laboratory, National Transonic Facility (NTF), NDE Research Laboratory, Polymers & Composites Laboratory, Pyrotechnic Test Facility, Quiet Flow Facility, Robotics Facilities, Scientific Visualization System, Scramjet Test Complex, Space Materials Research Laboratory, Space Simulation & Environmental Test Complex, Structural Dynamics Research Laboratory, Structural Dynamics Test Beds, Structures & Materials Research Laboratory, Supersonic Low Disturbance Pilot Tunnel, Thermal Acoustic Fatigue Apparatus (TAFA), Transonic Dynamics Tunnel (TDT), Transport Systems Research Vehicle, Unitary Plan Wind Tunnel, and the Visual Motion Simulator (VMS).

  17. Microbubble tunneling in gel phantoms

    PubMed Central

    Caskey, Charles F.; Qin, Shengping; Dayton, Paul A.; Ferrara, Katherine W.

    2009-01-01

    Insonified microbubbles were observed in vessels within a gel with a Young’s modulus similar to that of tissue, demonstrating shape instabilities, liquid jets, and the formation of small tunnels. In this study, tunnel formulation occurred in the direction of the propagating ultrasound wave, where radiation pressure directed the contact of the bubble and gel, facilitating the activity of the liquid jets. Combinations of ultrasonic parameters and microbubble concentrations that are relevant for diagnostic imaging and drug delivery and that lead to tunnel formation were applied and the resulting tunnel formation was quantified. PMID:19425620

  18. Magnetic tunnel junction pattern technique

    NASA Astrophysics Data System (ADS)

    Chen, Eugene; Schwarz, Benjamin; Choi, Chang Ju; Kula, Witold; Wolfman, Jerome; Ounadjela, Kamel; Geha, Sam

    2003-05-01

    We have developed a magnetic tunnel junction (MTJ) pattern technique that involves transforming the magnetic layer above the tunnel barrier in unwanted areas into an insulator, thus providing insulation between different MTJ devices without suffering common tunnel barrier shorting problems. With this technique, 90%-100% yielding MTJ devices have been observed. MTJ results using this process are superior to an etching based process. Switching distribution of patterned magnetic bits is also narrower using this novel technique. Process control and the ability to stop on the tunnel barrier have been demonstrated.

  19. Tunneling above the crossover temperature.

    PubMed

    Alvarez-Barcia, Sonia; Flores, Jesús R; Kästner, Johannes

    2014-01-09

    Quantum mechanical tunneling of atoms plays a significant role in many chemical reactions. The crossover temperature between classical and quantum movement is a convenient preliminary indication of the importance of tunneling for a particular reaction. Here we show, using instanton theory, that quantum tunneling is possible significantly above this crossover temperature for specific forms of the potential energy surface. We demonstrate the effect on an analytic potential as well as a chemical system. While protons move asynchronously along a Grotthuss chain in the classical high-temperature range, the onset of tunneling results in a synchronization of their movement.

  20. Ferroelectric tunneling under bias voltages

    NASA Astrophysics Data System (ADS)

    Ma, Z. J.; Chen, G.; Zhou, P.; Mei, Z. H.; Zhang, T. J.

    2017-01-01

    Theoretical investigations of ferroelectric tunneling in a SrRuO3/BaTiO3/Pt junction were conducted, and critical expressions for the surface charge density in the electrodes and the potential distribution across the tunnel junction were derived. It was found that the screening charges associated with the ferroelectric polarization and the charging effect of the capacitor jointly contribute to the charges in the electrodes. A current-voltage study simulating the ‘read’ operation indicated that the tunneling electroresistance effect increases with the ferroelectric thickness, and the tunneling electroresistance values agree well with experimental results.

  1. Improved Performances in Subsonic Flows of an SPH Scheme with Gradients Estimated Using an Integral Approach

    NASA Astrophysics Data System (ADS)

    Valdarnini, R.

    2016-11-01

    In this paper, we present results from a series of hydrodynamical tests aimed at validating the performance of a smoothed particle hydrodynamics (SPH) formulation in which gradients are derived from an integral approach. We specifically investigate the code behavior with subsonic flows, where it is well known that zeroth-order inconsistencies present in standard SPH make it particularly problematic to correctly model the fluid dynamics. In particular, we consider the Gresho-Chan vortex problem, the growth of Kelvin-Helmholtz instabilities, the statistics of driven subsonic turbulence and the cold Keplerian disk problem. We compare simulation results for the different tests with those obtained, for the same initial conditions, using standard SPH. We also compare the results with the corresponding ones obtained previously with other numerical methods, such as codes based on a moving-mesh scheme or Godunov-type Lagrangian meshless methods. We quantify code performances by introducing error norms and spectral properties of the particle distribution, in a way similar to what was done in other works. We find that the new SPH formulation exhibits strongly reduced gradient errors and outperforms standard SPH in all of the tests considered. In fact, in terms of accuracy, we find good agreement between the simulation results of the new scheme and those produced using other recently proposed numerical schemes. These findings suggest that the proposed method can be successfully applied for many astrophysical problems in which the presence of subsonic flows previously limited the use of SPH, with the new scheme now being competitive in these regimes with other numerical methods.

  2. Effect of impinging plate geometry on the self-excitation of subsonic impinging jets

    NASA Astrophysics Data System (ADS)

    Vinoth, B. R.; Rathakrishnan, E.

    2011-11-01

    In the generation of discrete tones by subsonic impinging jets, there exists a difference of opinion as how the feedback is achieved, i.e., the path of the feedback acoustic waves is whether inside the jet or outside the jet? The only available model (Tam and Ahuja model) for the prediction of an average subsonic jet impingement tone frequency assumes that the upstream part of the feedback loop is closed by an upstream propagating neutral wave of the jet. But, there is no information about the plate geometry in the model. The present study aims at understanding the effect of the plate geometry (size and co-axial hole in the plate) on the self-excitation process of subsonic impinging jets and the path of the acoustic feedback to the nozzle exit. The present results show that there is no effect of plate diameter on the frequency of the self-excitation. A new type of tones is generated for plates with co-axial hole (hole diameter is equal to nozzle exit diameter) for Mach numbers 0.9 and 0.95, in addition to the axisymmetric and helical mode tones observed for plates without co-axial hole. The stability results show that the Strouhal number of the least dispersive upstream propagating neutral waves match with the average Strouhal number of the new tones observed in the present experiments. The present study extends the validity of the model of Tam and Ahuja to a plate with co-axial hole (annular plate) and by doing so, we indirectly confirmed that the major acoustic feedback path to the nozzle exit is inside the jet.

  3. Observations of subsonic and supersonic shear flows in laser driven high-energy-density plasmas

    NASA Astrophysics Data System (ADS)

    Harding, E. C.

    2009-11-01

    Shear layers containing strong velocity gradients appear in many high-energy-density (HED) systems and play important roles in mixing and the transition to turbulence. Yet few laboratory experiments have been carried out to study their detailed evolution in this extreme environment where plasmas are compressible, actively ionizing, often involve strong shock waves and have complex material properties. Many shear flows produce the Kelvin-Helmholtz (KH) instability, which initiates the mixing at a fluid interface. We present results from two dedicated shear flow experiments that produced overall subsonic and supersonic flows using novel target designs. In the subsonic case, the Omega laser was used to drive a blast wave along a rippled interface between plastic and foam, shocking both the materials to produce two fluids separated by a sharp shear layer. The interface subsequently rolled-upped into large KH vortices that were accompanied by bubble-like structures of unknown origin. This was the first time the evolution of a well-resolved KH instability was observed in a HED plasma in the laboratory. We have analyzed the properties and dynamics of the plasma based on the data and fundamental models, without resorting to simulated values. In the second, supersonic experiment the Nike laser was used to drive a supersonic flow of Al plasma along a rippled, low-density foam surface. Here again the flowing plasma drove a shock into the second material, so that two fluids were separated by a shear layer. In contrast to the subsonic case, the flow developed shocks around the ripples in response to the supersonic flow of Al. Collaborators: R.P. Drake, O.A. Hurricane, J.F. Hansen, Y. Aglitskiy, T. Plewa, B.A. Remington, H.F. Robey, J.L. Weaver, A.L. Velikovich, R.S. Gillespie, M.J. Bono, M.J. Grosskopf, C.C. Kuranz, A. Visco.

  4. Pressure measurements on a thick cambered and twisted 58 deg delta wing at high subsonic speeds

    NASA Technical Reports Server (NTRS)

    Chu, Julio; Lamar, John E.

    1987-01-01

    A pressure experiment at high subsonic speeds was conducted by a cambered and twisted thick delta wing at the design condition (Mach number 0.80), as well as at nearby Mach numbers (0.75 and 0.83) and over an angle-of-attack range. Effects of twin vertical tails on the wing pressure measurements were also assessed. Comparisons of detailed theoretical and experimental surface pressures and sectional characteristics for the wing alone are presented. The theoretical codes employed are FLO-57, FLO-28, PAN AIR, and the Vortex Lattice Method-Suction Analogy.

  5. The calculation of pressure on slender airplanes in subsonic and supersonic flow

    NASA Technical Reports Server (NTRS)

    Heaslet, Max A; Lomas, Harvard

    1954-01-01

    Under the assumption that a wing, body, or wing-body combination is slender or flying at near sonic velocity, expressions are given which permit the calculation of pressure in the immediate vicinity of the configuration. The disturbance field, in both subsonic and supersonic flight, is shown to consist of two-dimensional disturbance fields extending laterally and a longitudinal field that depends on the streamwise growth of cross-sectional area. A discussion is also given of couplings, between lifting and thickness effects, that necessarily arise as a result of the quadratic dependence of pressure on the induced velocity components. (author)

  6. Three dimensional flow field measurements of a 4:1 aspect ratio subsonic jet

    NASA Technical Reports Server (NTRS)

    Morrison, G. L.; Swan, D. H.

    1989-01-01

    Flow field measurements for a subsonic rectangular cold air jet with an aspect ratio of 4:1 (12.7 x 50.8 mm) at a Mach number of 0.09 and Re of 100,000 have been carried out using a three-dimensional laser Doppler anemometer system. Mean velocity measurements show that the jet width spreads more rapidly along the minor axis than along the major axis. The outward velocities, however, are not significantly different for the two axes, indicating the presence of enhanced mixing along the minor axis. The jet slowly changes from a rectangular jet to a circular jet as the flow progresses downstream.

  7. Development of a shock noise prediction code for high-speed helicopters - The subsonically moving shock

    NASA Technical Reports Server (NTRS)

    Tadghighi, H.; Holz, R.; Farassat, F.; Lee, Yung-Jang

    1991-01-01

    A previously defined airfoil subsonic shock-noise prediction formula whose result depends on a mapping of the time-dependent shock surface to a time-independent computational domain is presently coded and incorporated in the NASA-Langley rotor-noise prediction code, WOPWOP. The structure and algorithms used in the shock-noise prediction code are presented; special care has been taken to reduce computation time while maintaining accuracy. Numerical examples of shock-noise prediction are presented for hover and forward flight. It is confirmed that shock noise is an important component of the quadrupole source.

  8. Effects of high subsonic flow on sound propagation in a variable-area duct

    NASA Technical Reports Server (NTRS)

    Callegari, A. J.; Myers, M. K.

    1976-01-01

    The propagation of sound in a converging-diverging duct containing a quasi-one-dimensional steady flow with a high subsonic throat Mach number was studied. The behavior of linearized acoustic theory at the throat of the duct was shown to be singular. This singularity implies that linearized acoustic theory is invalid. The explicit singular behavior was determined and used to sketch the development (by the method of matched asymptotic expansions) of a nonlinear theory for sound propagation in a sonic throat region.

  9. Subsonic flutter analysis addition to NASTRAN. [for use with CDC 6000 series digital computers

    NASA Technical Reports Server (NTRS)

    Doggett, R. V., Jr.; Harder, R. L.

    1973-01-01

    A subsonic flutter analysis capability has been developed for NASTRAN, and a developmental version of the program has been installed on the CDC 6000 series digital computers at the Langley Research Center. The flutter analysis is of the modal type, uses doublet lattice unsteady aerodynamic forces, and solves the flutter equations by using the k-method. Surface and one-dimensional spline functions are used to transform from the aerodynamic degrees of freedom to the structural degrees of freedom. Some preliminary applications of the method to a beamlike wing, a platelike wing, and a platelike wing with a folded tip are compared with existing experimental and analytical results.

  10. Study of the application of hydrogen fuel to long-range subsonic transport aircraft, volume 2

    NASA Technical Reports Server (NTRS)

    Brewer, G. D.; Morris, R. E.; Lange, R. H.; Moore, J. W.

    1975-01-01

    The feasibility, practicability, and potential advantages/disadvantages of using liquid hydrogen as fuel in long range, subsonic transport aircraft of advanced design were studied. Both passenger and cargo-type aircraft were investigated. To provide a valid basis for comparison, conventional hydrocarbon (Jet A) fueled aircraft were designed to perform identical missions using the same advanced technology and meeting the same operational constraints. The liquid hydrogen and Jet A fueled aircraft were compared on the basis of weight, size, energy utilization, cost, noise, emissions, safety, and operational characteristics. A program of technology development was formulated.

  11. A free-flight experiment of projectiles ranging from high subsonic to high supersonic Mach numbers

    NASA Astrophysics Data System (ADS)

    Oguchi, H.; Funabiki, K.; Sato, S.; Hatakeyama, M.

    1991-09-01

    This paper reports a preliminary experiment with a free-flight range which was designed to be able to be driven by means of a sort of fast-acting valve. The notable characteristics pertinent to this type of range is the pliancy of operation and also the wide coverage of flight speed from the high subsonic to the high supersonic range. For both spherical and vehicle models, flight tests were conducted with shadowgraph observations and flight speed measurements in order to examine the capability of this type of facility.

  12. Development of RTM and powder prepreg resins for subsonic aircraft primary structures

    NASA Technical Reports Server (NTRS)

    Woo, Edmund P.; Groleau, Michael R.; Bertram, James L.; Puckett, Paul M.; Maynard, Shawn J.

    1993-01-01

    Dow developed a thermoset resin which could be used to produce composites via the RTM process. The composites formed are useful at 200 F service temperatures after moisture saturation, and are tough systems that are suitable for subsonic aircraft primary structure. At NASA's request, Dow also developed a modified version of the RTM resin system which was suitable for use in producing powder prepreg. In the course of developing the RTM and powder versions of these resins, over 50 different new materials were produced and evaluated.

  13. 3D CFD modeling of subsonic and transonic flowing-gas DPALs with different pumping geometries

    NASA Astrophysics Data System (ADS)

    Yacoby, Eyal; Sadot, Oren; Barmashenko, Boris D.; Rosenwaks, Salman

    2015-10-01

    Three-dimensional computational fluid dynamics (3D CFD) modeling of subsonic (Mach number M ~ 0.2) and transonic (M ~ 0.9) diode pumped alkali lasers (DPALs), taking into account fluid dynamics and kinetic processes in the lasing medium is reported. The performance of these lasers is compared with that of supersonic (M ~ 2.7 for Cs and M ~ 2.4 for K) DPALs. The motivation for this study stems from the fact that subsonic and transonic DPALs require much simpler hardware than supersonic ones where supersonic nozzle, diffuser and high power mechanical pump (due to a drop in the gas total pressure in the nozzle) are required for continuous closed cycle operation. For Cs DPALs with 5 x 5 cm2 flow cross section pumped by large cross section (5 x 2 cm2) beam the maximum achievable power of supersonic devices is higher than that of the transonic and subsonic devices by only ~ 3% and ~ 10%, respectively. Thus in this case the supersonic operation mode has no substantial advantage over the transonic one. The main processes limiting the power of Cs supersonic DPALs are saturation of the D2 transition and large ~ 60% losses of alkali atoms due to ionization, whereas the influence of gas heating is negligible. For K transonic DPALs both the gas heating and ionization effects are shown to be unimportant. The maximum values of the power are higher than those in Cs transonic laser by ~ 11%. The power achieved in the supersonic and transonic K DPAL is higher than for the subsonic version, with the same resonator and K density at the inlet, by ~ 84% and ~ 27%, respectively, showing a considerable advantaged of the supersonic device over the transonic one. For pumping by rectangular beams of the same (5 x 2 cm2) cross section, comparison between end-pumping - where the laser beam and pump beam both propagate at along the same axis, and transverse-pumping - where they propagate perpendicularly to each other, shows that the output power and optical-to-optical efficiency are not

  14. Applicability of the independence principle to subsonic turbulent flow over a swept rearward-facing step

    NASA Technical Reports Server (NTRS)

    Selby, G. V.

    1983-01-01

    Prandtl (1946) has concluded that for yawed laminar incompressible flows the streamwise flow is independent of the spanwise flow. However, Ashkenas and Riddell (1955) have reported that for turbulent flow the 'independence principle' does not apply to yawed flat plates. On the other hand, it was also found that this principle may be applicable to many turbulent flows. As the sweep angle is increased, a sweep angle is reached which defines the interval over which the 'independence principle' is valid. The results obtained in the present investigation indicate the magnitude of the critical angle for subsonic turbulent flow over a swept rearward-facing step.

  15. Advanced subsonic long-haul transport terminal area compatibility study. Volume 1: Compatibility assessment

    NASA Technical Reports Server (NTRS)

    1974-01-01

    An analysis was made to identify airplane research and technology necessary to ensure advanced transport aircraft the capability of accommodating forecast traffic without adverse impact on airport communities. Projections were made of the delay, noise, and emissions impact of future aircraft fleets on typical large urban airport. Design requirements, based on these projections, were developed for an advanced technology, long-haul, subsonic transport. A baseline aircraft was modified to fulfill the design requirements for terminal area compatibility. Technical and economic comparisons were made between these and other aircraft configured to support the study.

  16. Theoretical methods and design studies for NLF and HLFC swept wings at subsonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Goradia, Suresh H.; Morgan, Harry L., Jr.

    1987-01-01

    Laminarization of the boundary layer on the surface of aircraft wings can be accomplished by the use of concepts such as Natural Laminar Flow (NLF), Laminar-Flow Control (LFC), and Hybrid Laminar-Flow Control (HLFC). Several integral boundary-layer methods were developed for the prediction of laminar, transition, and separating turbulent boundary layers. These methods were developed for use at either subsonic or supersonic speeds, have small computer execution times, and are simple to use. The theoretical equations and assumptions which form the basis of the boundary-layer method, are briefly outlined and the results of several correlation cases with exciting experimental data are presented.

  17. Subsonic Ultra Green Aircraft Research Phase II: N+4 Advanced Concept Development

    NASA Technical Reports Server (NTRS)

    Bradley, Marty K.; Droney, Christopher K.

    2012-01-01

    This final report documents the work of the Boeing Subsonic Ultra Green Aircraft Research (SUGAR) team on Task 1 of the Phase II effort. The team consisted of Boeing Research and Technology, Boeing Commercial Airplanes, General Electric, and Georgia Tech. Using a quantitative workshop process, the following technologies, appropriate to aircraft operational in the N+4 2040 timeframe, were identified: Liquefied Natural Gas (LNG), Hydrogen, fuel cell hybrids, battery electric hybrids, Low Energy Nuclear (LENR), boundary layer ingestion propulsion (BLI), unducted fans and advanced propellers, and combinations. Technology development plans were developed.

  18. An advanced panel method for analysis of arbitrary configurations in unsteady subsonic flow

    NASA Technical Reports Server (NTRS)

    Dusto, A. R.; Epton, M. A.

    1980-01-01

    An advanced method is presented for solving the linear integral equations for subsonic unsteady flow in three dimensions. The method is applicable to flows about arbitrary, nonplanar boundary surfaces undergoing small amplitude harmonic oscillations about their steady mean locations. The problem is formulated with a wake model wherein unsteady vorticity can be convected by the steady mean component of flow. The geometric location of the unsteady source and doublet distributions can be located on the actual surfaces of thick bodies in their steady mean locations. The method is an outgrowth of a recently developed steady flow panel method and employs the linear source and quadratic doublet splines of that method.

  19. Trim and Structural Optimization of Subsonic Transport Wings Using Nonconventional Aeroelastic Tailoring

    NASA Technical Reports Server (NTRS)

    Stanford, Bret K.; Jutte, Christine V.

    2014-01-01

    Several minimum-mass aeroelastic optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic strength and panel buckling constraints are imposed across a variety of trimmed maneuver loads. Tailoring with metallic thickness variations, functionally graded materials, composite laminates, tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.

  20. Design and performance of a parachute for supersonic and subsonic recovery of an 800-lb payload

    SciTech Connect

    Peterson, C.W.; Waye, D.E.; Rollstin, L.R.; Holt, I.T.

    1986-01-01

    The design and development of a parachute system to recover an 800-lb test payload flying at both supersonic and subsonic speeds are presented. Performance data from full-scale flight tests for several parachute configurations were used to define an acceptable parachute system and to gain insight into which design parameters are most critical for supersonic parachute design. The suspension line length, canopy configuration, and forebody wake have a major effect upon parachute performance and stability. A tractor rocket was used to deploy this parachute system successfully at all deployment speeds.

  1. An example of requirements for Advanced Subsonic Civil Transport (ASCT) flight control system using structured techniques

    NASA Technical Reports Server (NTRS)

    Mclees, Robert E.; Cohen, Gerald C.

    1991-01-01

    The requirements are presented for an Advanced Subsonic Civil Transport (ASCT) flight control system generated using structured techniques. The requirements definition starts from initially performing a mission analysis to identify the high level control system requirements and functions necessary to satisfy the mission flight. The result of the study is an example set of control system requirements partially represented using a derivative of Yourdon's structured techniques. Also provided is a research focus for studying structured design methodologies and in particular design-for-validation philosophies.

  2. Planar Doppler Velocimetry for Large-Scale Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    McKenzie, Robert L.

    1997-01-01

    Recently, Planar Doppler Velocimetry (PDV) has been shown by several laboratories to offer an attractive means for measuring three-dimensional velocity vectors everywhere in a light sheet placed in a flow. Unlike other optical means of measuring flow velocities, PDV is particularly attractive for use in large wind tunnels where distances to the sample region may be several meters, because it does not require the spatial resolution and tracking of individual scattering particles or the alignment of crossed beams at large distances. To date, demonstrations of PDV have been made either in low speed flows without quantitative comparison to other measurements, or in supersonic flows where the Doppler shift is large and its measurement is relatively insensitive to instrumental errors. Moreover, most reported applications have relied on the use of continuous-wave lasers, which limit the measurement to time-averaged velocity fields. This work summarizes the results of two previous studies of PDV in which the use of pulsed lasers to obtain instantaneous velocity vector fields is evaluated. The objective has been to quantitatively define and demonstrate PDV capabilities for applications in large-scale wind tunnels that are intended primarily for the production testing of subsonic aircraft. For such applications, the adequate resolution of low-speed flow fields requires accurate measurements of small Doppler shifts that are obtained at distances of several meters from the sample region. The use of pulsed lasers provides the unique capability to obtain not only time-averaged fields, but also their statistical fluctuation amplitudes and the spatial excursions of unsteady flow regions such as wakes and separations. To accomplish the objectives indicated, the PDV measurement process is first modeled and its performance evaluated computationally. The noise sources considered include those related to the optical and electronic properties of Charge-Coupled Device (CCD) arrays and to

  3. Investing American Recovery and Reinvestment Act Funds to Advance Capability, Reliability, and Performance in NASA Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Sydnor, Goerge H.

    2010-01-01

    The National Aeronautics and Space Administration's (NASA) Aeronautics Test Program (ATP) is implementing five significant ground-based test facility projects across the nation with funding provided by the American Recovery and Reinvestment Act (ARRA). The projects were selected as the best candidates within the constraints of the ARRA and the strategic plan of ATP. They are a combination of much-needed large scale maintenance, reliability, and system upgrades plus creating new test beds for upcoming research programs. The projects are: 1.) Re-activation of a large compressor to provide a second source for compressed air and vacuum to the Unitary Plan Wind Tunnel at the Ames Research Center (ARC) 2.) Addition of high-altitude ice crystal generation at the Glenn Research Center Propulsion Systems Laboratory Test Cell 3, 3.) New refrigeration system and tunnel heat exchanger for the Icing Research Tunnel at the Glenn Research Center, 4.) Technical viability improvements for the National Transonic Facility at the Langley Research Center, and 5.) Modifications to conduct Environmentally Responsible Aviation and Rotorcraft research at the 14 x 22 Subsonic Tunnel at Langley Research Center. The selection rationale, problem statement, and technical solution summary for each project is given here. The benefits and challenges of the ARRA funded projects are discussed. Indirectly, this opportunity provides the advantages of developing experience in NASA's workforce in large projects and maintaining corporate knowledge in that very unique capability. It is envisioned that improved facilities will attract a larger user base and capabilities that are needed for current and future research efforts will offer revenue growth and future operations stability. Several of the chosen projects will maximize wind tunnel reliability and maintainability by using newer, proven technologies in place of older and obsolete equipment and processes. The projects will meet NASA's goal of

  4. Computed and Experimental Flutter/LCO Onset for the Boeing Truss-Braced Wing Wind-Tunnel Model

    NASA Technical Reports Server (NTRS)

    Bartels, Robert E.; Scott, Robert C.; Funk, Christie J.; Allen, Timothy J.; Sexton, Bradley W.

    2014-01-01

    This paper presents high fidelity Navier-Stokes simulations of the Boeing Subsonic Ultra Green Aircraft Research truss-braced wing wind-tunnel model and compares the results to linear MSC. Nastran flutter analysis and preliminary data from a recent wind-tunnel test of that model at the NASA Langley Research Center Transonic Dynamics Tunnel. The simulated conditions under consideration are zero angle of attack, so that structural nonlinearity can be neglected. It is found that, for Mach number greater than 0.78, the linear flutter analysis predicts flutter onset dynamic pressure below the wind-tunnel test and that predicted by the Navier-Stokes analysis. Furthermore, the wind-tunnel test revealed that the majority of the high structural dynamics cases were wing limit cycle oscillation (LCO) rather than flutter. Most Navier-Stokes simulated cases were also LCO rather than hard flutter. There is dip in the wind-tunnel test flutter/LCO onset in the Mach 0.76-0.80 range. Conditions tested above that Mach number exhibited no aeroelastic instability at the dynamic pressures reached in the tunnel. The linear flutter analyses do not show a flutter/LCO dip. The Navier-Stokes simulations also do not reveal a dip; however, the flutter/LCO onset is at a significantly higher dynamic pressure at Mach 0.90 than at lower Mach numbers. The Navier-Stokes simulations indicate a mild LCO onset at Mach 0.82, then a more rapidly growing instability at Mach 0.86 and 0.90. Finally, the modeling issues and their solution related to the use of a beam and pod finite element model to generate the Navier-Stokes structure mode shapes are discussed.

  5. Pre-Test Assessment of the Use Envelope of the Normal Force of a Wind Tunnel Strain-Gage Balance

    NASA Technical Reports Server (NTRS)

    Ulbrich, N.

    2016-01-01

    The relationship between the aerodynamic lift force generated by a wind tunnel model, the model weight, and the measured normal force of a strain-gage balance is investigated to better understand the expected use envelope of the normal force during a wind tunnel test. First, the fundamental relationship between normal force, model weight, lift curve slope, model reference area, dynamic pressure, and angle of attack is derived. Then, based on this fundamental relationship, the use envelope of a balance is examined for four typical wind tunnel test cases. The first case looks at the use envelope of the normal force during the test of a light wind tunnel model at high subsonic Mach numbers. The second case examines the use envelope of the normal force during the test of a heavy wind tunnel model in an atmospheric low-speed facility. The third case reviews the use envelope of the normal force during the test of a floor-mounted semi-span model. The fourth case discusses the normal force characteristics during the test of a rotated full-span model. The wind tunnel model's lift-to-weight ratio is introduced as a new parameter that may be used for a quick pre-test assessment of the use envelope of the normal force of a balance. The parameter is derived as a function of the lift coefficient, the dimensionless dynamic pressure, and the dimensionless model weight. Lower and upper bounds of the use envelope of a balance are defined using the model's lift-to-weight ratio. Finally, data from a pressurized wind tunnel is used to illustrate both application and interpretation of the model's lift-to-weight ratio.

  6. The influence of subsonic mission segments on the use of variable-sweep wings for high speed civil transport configurations

    NASA Technical Reports Server (NTRS)

    Martin, Glenn L.; Beissner, Fred L., Jr.; Domack, Christopher S.; Shields, E. William

    1988-01-01

    A Mach-3.0, 250-passenger, 6500-n. mi. range SST configuration's alternative use of fixed-planform or variable-sweep wings is presently evaluated, with a view to effects on aerodynamics, mission performance, and sizing. After preliminary design, the fixed and variable-wing configurations were resized to perform missions incorporating subsonic cruise segments of as much as 4000 n. mi.; the effect of subsonic segment length on design gross weight and block time was then ascertained. Due to the reduced supersonic efficiency of the variable-sweep aircraft, over one-half of the 6500-n. mi. mission would have to be flown subsonically for its sizing to reach a lower ramp weight than that of its fixed-geometry counterpart.

  7. A conservative type-dependent full potential method for the treatment of supersonic flows with embedded subsonic regions

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Szema, K.-Y.; Osher, S.

    1983-01-01

    A nonlinear method based on the full potential equation in conservation form, cast in an arbitrary coordinate system, has been developed to treat predominantly supersonic flows with embedded subsonic regions. This type of flow field occurs frequently near the fuselage-canopy junction area and wing leading edge regions for a moderately swept fighter configuration. The method uses the theory of characteristics to accurately monitor the type-dependent flow field. A conservative switching scheme is developed to transition from the supersonic marching algorithm to a subsonic relaxation procedure, and vice versa. An implicit approximate factorization scheme is employed to solve the finite-differenced equation. Results are shown for a few configurations, including a wing-body-wake realistic fighter model having embedded subsonic regions.

  8. Carpal tunnel syndrome

    PubMed Central

    Aroori, Somaiah; Spence, Roy AJ

    2008-01-01

    Carpal tunnel syndrome is one of the most common peripheral neuropathies. It affects mainly middle aged women. In the majority of patients the exact cause and pathogenesis of CTS is unclear. Although several occupations have been linked to increased incidence and prevalence of CTS the evidence is not clear. Occupational CTS is uncommon and it is essential to exclude all other causes particularly the intrinsic factors such as obesity before attributing it to occupation. The risk of CTS is high in occupations involving exposure to high pressure, high force, repetitive work, and vibrating tools. The classic symptoms of CTS include nocturnal pain associated with tingling and numbness in the distribution of median nerve in the hand. There are several physical examination tests that will help in the diagnosis of CTS but none of these tests are diagnostic on their own. The gold standard test is nerve conduction studies. However, they are also associated with false positive and false negative results. The diagnosis of CTS should be based on history, physical examination and results of electrophysiological studies. The patient with mild symptoms of CTS can be managed with conservative treatment, particularly local injection of steroids. However, in moderate to severe cases, surgery is the only treatment that provides cure. The basic principle of surgery is to increase the volume of the carpal tunnel by dividing transverse carpal ligament to release the pressure on the median nerve. Apart from early recovery and return to work there is no significant difference in terms of early and late complications and long-term pain relief between endoscopic and open carpal tunnel surgery. PMID:18269111

  9. Full Scale Tunnel (FST)

    NASA Technical Reports Server (NTRS)

    1930-01-01

    Construction of Full Scale Tunnel (FST). In November 1929, Smith DeFrance submitted his recommendations for the general design of the Full Scale Wind Tunnel. The last on his list concerned the division of labor required to build this unusual facility. He believed the job had five parts and described them as follows: 'It is proposed that invitations be sent out for bids on five groups of items. The first would be for one contract on the complete structure; second the same as first, including the erection of the cones but not the fabrication, since this would be more of a shipyard job; third would cover structural steel, cover, sash and doors, but not cones or foundation; fourth, foundations; an fifth, fabrication of cones.' DeFrance's memorandum prompted the NACA to solicit estimates from a large number of companies. Preliminary designs and estimates were prepared and submitted to the Bureau of the Budget and Congress appropriated funds on February 20, 1929. The main construction contract with the J.A. Jones Company of Charlotte, North Carolina was signed one year later on February 12, 1930. It was a peculiar structure as the building's steel framework is visible on the outside of the building. DeFrance described this in NACA TR No. 459: 'The entire equipment is housed in a structure, the outside walls of which serve as the outer walls of the return passages. The over-all length of the tunnel is 434 feet 6 inches, the width 222 feet, and the maximum height 97 feet. The framework is of structural steel....' (pp. 292-293)

  10. Full Scale Tunnel (FST)

    NASA Technical Reports Server (NTRS)

    1930-01-01

    Construction of Full-Scale Tunnel (FST). In November 1929, Smith DeFrance submitted his recommendations for the general design of the Full Scale Wind Tunnel. The last on his list concerned the division of labor required to build this unusual facility. He believed the job had five parts and described them as follows: 'It is proposed that invitations be sent out for bids on five groups of items. The first would be for one contract on the complete structure; second the same as first, including the erection of the cones but not the fabrication, since this would be more of a shipyard job; third would cover structural steel, cover, sash and doors, but not cones or foundation; fourth, foundations; and fifth, fabrication of cones.' DeFrance's memorandum prompted the NACA to solicit estimates from a large number of companies. Preliminary designs and estimates were prepared and submitted to the Bureau of the Budget and Congress appropriated funds on February 20, 1929. The main construction contract with the J.A. Jones Company of Charlotte, North Carolina was signed one year later on February 12, 1930. It was a peculiar structure as the building's steel framework is visible on the outside of the building. DeFrance described this in NACA TR No. 459: 'The entire equipment is housed in a structure, the outside walls of which serve as the outer walls of the return passages. The over-all length of the tunnel is 434 feet 6 inches, the width 222 feet, and the maximum height 97 feet. The framework is of structural steel....' (pp. 292-293).

  11. Application of Rapid Prototyping Methods to High-Speed Wind Tunnel Testing

    NASA Technical Reports Server (NTRS)

    Springer, A. M.

    1998-01-01

    This study was undertaken in MSFC's 14-Inch Trisonic Wind Tunnel to determine if rapid prototyping methods could be used in the design and manufacturing of high speed wind tunnel models in direct testing applications, and if these methods would reduce model design/fabrication time and cost while providing models of high enough fidelity to provide adequate aerodynamic data, and of sufficient strength to survive the test environment. Rapid prototyping methods utilized to construct wind tunnel models in a wing-body-tail configuration were: fused deposition method using both ABS plastic and PEEK as building materials, stereolithography using the photopolymer SL-5170, selective laser sintering using glass reinforced nylon, and laminated object manufacturing using plastic reinforced with glass and 'paper'. This study revealed good agreement between the SLA model, the metal model with an FDM-ABS nose, an SLA nose, and the metal model for most operating conditions, while the FDM-ABS data diverged at higher loading conditions. Data from the initial SLS model showed poor agreement due to problems in post-processing, resulting in a different configuration. A second SLS model was tested and showed relatively good agreement. It can be concluded that rapid prototyping models show promise in preliminary aerodynamic development studies at subsonic, transonic, and supersonic speeds.

  12. Computational design of low aspect ratio wing-winglet configurations for transonic wind-tunnel tests

    NASA Technical Reports Server (NTRS)

    Kuhlman, John M.; Brown, Christopher K.

    1989-01-01

    Computational designs were performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three configurations was selected to be constructed as a wind tunnel model for testing in the NASA LaRC 8-foot transonic pressure tunnel. A design point of M = 0.8, C(sub L) is approximate or = to 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 deg and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. The design process and the predicted transonic performance are summarized for each configuration. In addition, a companion low-speed design study was conducted, using one of the transonic design wing-winglet planforms but with different camber and thickness distributions. A low-speed wind tunnel model was constructed to match this low-speed design geometry, and force coefficient data were obtained for the model at speeds of 100 to 150 ft/sec. Measured drag coefficient reductions were of the same order of magnitude as those predicted by numerical subsonic performance predictions.

  13. Transonic Dynamics Tunnel Force and Pressure Data Acquired on the HSR Rigid Semispan Model

    NASA Technical Reports Server (NTRS)

    Schuster, David M.; Rausch, Russ D.

    1999-01-01

    This report describes the aerodynamic data acquired on the High Speed Research Rigid Semispan Model (HSR-RSM) during NASA Langley Transonic Dynamics Tunnel (TDT) Test 520 conducted from 18 March to 4 April, 1996. The purpose of this test was to assess the aerodynamic character of a rigid high speed civil transport wing. The wing was fitted with a single trailing edge control surface which was both steadily deflected and oscillated during the test to investigate the response of the aerodynamic data to steady and unsteady control motion. Angle-of-attack and control surface deflection polars at subsonic, transonic and low-supersonic Mach numbers were obtained in the tunnel?s heavy gas configuration. Unsteady pressure and steady loads data were acquired on the wing, while steady pressures were measured on the fuselage. These data were reduced using a variety of methods, programs and computer systems. The reduced data was ultimately compiled onto a CD-ROM volume which was distributed to HSR industry team members in July, 1996. This report documents the methods used to acquire and reduce the data, and provides an assessment of the quality, repeatability, and overall character of the aerodynamic data measured during this test.

  14. Comparison of Ares I-X Wind-Tunnel Derived Buffet Environment with Flight Data

    NASA Technical Reports Server (NTRS)

    Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.

    2011-01-01

    The Ares I-X Flight Test Vehicle (FTV), launched in October 2009, carried with it over 243 buffet verification pressure sensors and was one of the most heavily instrumented launch vehicle flight tests. This flight test represented a unique opportunity for NASA and its partners to compare the wind-tunnel derived buffet environment with that measured during the flight of Ares I-X. It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. Ares I-X buffet forcing functions were obtained via wind-tunnel testing of a rigid buffet model (RBM) instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. This paper discusses the comparison of RBM and FTV buffet environments, including fluctuating pressure coefficient and normalized sectional buffet forcing function root-mean-square magnitudes, frequency content of power-spectral density functions, and force magnitudes of an alternating flow phenomena. Comparison of wind-tunnel model and flight test vehicle buffet environments show very good agreement with root-mean-square magnitudes of buffet forcing functions at the majority of vehicle stations. Spectra proved a challenge to compare because of different wind-tunnel and flight test conditions and data acquisition rates. However, meaningful and promising comparisons of buffet spectra are presented. Lastly, the buffet loads resulting from the transition of subsonic separated flow to supersonic attached flow were significantly over-predicted by wind-tunnel results.

  15. Recurrent tarsal tunnel syndrome.

    PubMed

    Gould, John S

    2014-09-01

    Recurrence of tarsal tunnel syndrome after surgery may be due to inadequate release, lack of understanding or appreciation of the actual anatomy involved, variations in the anatomy of the nerve(s), failure to execute the release properly, bleeding with subsequent scarring, damage to the nerve and branches, persistent hypersensitivity of the nerves, and preexisting intrinsic damage to the nerve. Approaches include more thorough release, use of barrier materials to decrease adherence of the nerve to surrounding tissues to avoid traction neuritis, excisions of neuromas using conduits, and consideration of nerve stimulators and systemic medications to deal with persistent neural pain.

  16. Icing Research Tunnel

    NASA Technical Reports Server (NTRS)

    Chennault, Jonathan

    2004-01-01

    The Icing Research Tunnel in Building 11 at the NASA Glenn Research Center is committed to researching the effects of in flight icing on aircraft and testing ways to stop the formation of hazardous icing conditions on planes. During this summer, I worked here with Richard DelRosa, the lead engineer for this area. address one of the major concerns of aviation: icing conditions. During the war, many planes crashed (especially supply planes going over the.Himalayas) because ice built up in their wings and clogged the engines. To this day, it remains the largest ice tunnel in the world, with a test section that measures 6 feet high, 9 feet long, and 20 feet wide. It can simulate airspeeds from 50 to 300 miles per hour at temperatures as low as -50 Fahrenheit. Using these capabilities, IRT can simulate actual conditions at high altitudes. The first thing I did was creating a cross reference in Microsoft Excel. It lists commands for the DPU units that control the pressure and temperature variations in the tunnel, as well as the type of command (keyboard, multiplier, divide, etc). The cross reference also contains the algorithm for every command, and which page it is listed in on the control sheet (visual Auto-CAD graphs, which I helped to make). I actually spent most of the time on the computer using Auto-CAD. I drew a diagram of the entire icing tunnel and then drew diagrams of its various parts. Between my mentor and me, we have drawings of every part of it, from the spray bars to the thermocouples, power cabinets, input-output connectors for power systems, and layouts of various other machines. I was also responsible for drawing schematics for the Escort system (which controls the spray bars), the power system, DPUs, and other electrical systems. In my spare time, I am attempting to build and program the "toddler". Toddler is a walking robot that I have to program in PBASIC language. When complete, it should be able to walk on level terrain while avoiding obstacles in

  17. Effect of wind tunnel acoustic modes on linear oscillating cascade aerodynamics

    NASA Technical Reports Server (NTRS)

    Buffum, D. H.; Fleeter, S.

    1994-01-01

    The aerodynamics of a biconvex airfoil cascade oscillating in torsion is investigated using the unsteady aerodynamic influence coefficient technique. For subsonic flow and reduced frequencies as large as 0.9, airfoil surface unsteady pressures resulting from oscillation of one of the airfoils are measured using flush-mounted high-frequency-response pressure transducers. The influence coefficient data are examined in detail and then used to predict the unsteady aerodynamics of a cascade oscillating at various interblade phase angles. These results are correlated with experimental data obtained in the traveling-wave mode of oscillation and linearized analysis predictions. It is found that the unsteady pressure disturbances created by an oscillating airfoil excite wind tunnel acoustic modes, which have detrimental effects on the experimental results. Acoustic treatment is proposed to rectify this problem.

  18. Effect of wind tunnel acoustic modes on linear oscillating cascade aerodynamics

    NASA Technical Reports Server (NTRS)

    Buffum, Daniel H.; Fleeter, Sanford

    1993-01-01

    The aerodynamics of a biconvex airfoil cascade oscillating in torsion is investigated using the unsteady aerodynamic influence coefficient technique. For subsonic flow and reduced frequencies as large as 0.9, airfoil surface unsteady pressures resulting from oscillation of one of the airfoils are measured using flush-mounted high-frequency-response pressure transducers. The influence coefficient data are examined in detail and then used to predict the unsteady aerodynamics of a cascade oscillating at various interblade phase angles. These results are correlated with experimental data obtained in the traveling-wave mode of oscillation and linearized analysis predictions. It is found that the unsteady pressure disturbances created by an oscillating airfoil excite wind tunnel acoustic modes which have detrimental effects on the experimental data. Acoustic treatment is proposed to rectify this problem.

  19. The Langley Wind Tunnel Enterprise

    NASA Technical Reports Server (NTRS)

    Paulson, John W., Jr.; Kumar, Ajay; Kegelman, Jerome T.

    1998-01-01

    After 4 years of existence, the Langley WTE is alive and growing. Significant improvements in the operation of wind tunnels have been demonstrated and substantial further improvements are expected when we are able to truly address and integrate all the processes affecting the wind tunnel testing cycle.

  20. Supersonic Wind Tunnel Test Section

    NASA Technical Reports Server (NTRS)

    1957-01-01

    8ft x 6ft Supersonic Wind Tunnel Test-Section showing changes made in Stainless Steel walls with 17 inch inlet model installation. The model is the ACN Nozzle model used for aircraft engines. The Supersonic Wind Tunnel is located in the Lewis Flight Propulsion Laboratory, now John H. Glenn Research Center

  1. Early Childhood: Funnels and Tunnels.

    ERIC Educational Resources Information Center

    Fowlkes, Mary Anne

    1985-01-01

    Suggests using funnels and tunnels in combination with water, blocks, transportation toys, and other materials to help teach preschoolers to make predictions. Many examples are included for using funnels to understand properties of liquids and for using tunnels to predict order. (DH)

  2. A comparison of the experimental subsonic pressure distributions about several bodies of revolution with pressure distributions computed by means of the linearized theory

    NASA Technical Reports Server (NTRS)

    Matthews, Clarence W

    1953-01-01

    An analysis is made of the effects of compressibility on the pressure coefficients about several bodies of revolution by comparing experimentally determined pressure coefficients with corresponding pressure coefficients calculated by the use of the linearized equations of compressible flow. The results show that the theoretical methods predict the subsonic pressure-coefficient changes over the central part of the body but do not predict the pressure-coefficient changes near the nose. Extrapolation of the linearized subsonic theory into the mixed subsonic-supersonic flow region fails to predict a rearward movement of the negative pressure-coefficient peak which occurs after the critical stream Mach number has been attained. Two equations developed from a consideration of the subsonic compressible flow about a prolate spheroid are shown to predict, approximately, the change with Mach number of the subsonic pressure coefficients for regular bodies of revolution of fineness ratio 6 or greater.

  3. Two tunnels to inflation

    SciTech Connect

    Aguirre, Anthony; Johnson, Matthew C.

    2006-06-15

    We investigate the formation via tunneling of inflating (false-vacuum) bubbles in a true-vacuum background, and the reverse process. Using effective potentials from the junction condition formalism, all true- and false-vacuum bubble solutions with positive interior and exterior cosmological constant, and arbitrary mass are catalogued. We find that tunneling through the same effective potential appears to describe two distinct processes: one in which the initial and final states are separated by a wormhole (the Farhi-Guth-Guven mechanism), and one in which they are either in the same hubble volume or separated by a cosmological horizon. In the zero-mass limit, the first process corresponds to the creation of an inhomogenous universe from nothing, while the second mechanism is equivalent to the nucleation of true- or false-vacuum Coleman-De Luccia bubbles. We compute the probabilities of both mechanisms in the WKB approximation using semiclassical Hamiltonian methods, and find that--assuming both process are allowed--neither mechanism dominates in all regimes.

  4. Carpal tunnel syndrome.

    PubMed

    Chammas, M

    2014-04-01

    Carpal tunnel syndrome is the commonest entrapment neuropathy and is due to combined compression and traction on the median nerve at the wrist. It is often idiopathic. Although spontaneous resolution is possible, the usual natural evolution is slow progression. Diagnosis is mainly clinical depending on symptoms and provocative tests. An electromyogram is recommended preoperatively and in cases of work-related disease. Medical treatment is indicated early on or in cases with no deficit and consists of steroid injection in the canal or a night splint in neutral wrist position. Surgical treatment is by section of the flexor retinaculum and is indicated in resistance to medical treatment, in deficit or acute cases. Mini-invasive techniques such as endoscopic and mini-open approaches to carpal tunnel release with higher learning curves are justified by the shorter functional recovery time compared to classical surgery, but with identical long-term results. The choice depends on the surgeon's preference, patient information, stage of severity, etiology and availability of material. Results are satisfactory in 90% of cases. Nerve recovery depends on the stage of severity as well as general patient factors. Recovery of force takes about 2-3 months after the disappearance of 'pillar pain'. This operation has a benign reputation with a 0.2-0.5% reported neurovascular complication rate.

  5. Tunneling magnetic force microscopy

    NASA Technical Reports Server (NTRS)

    Burke, Edward R.; Gomez, Romel D.; Adly, Amr A.; Mayergoyz, Isaak D.

    1993-01-01

    We have developed a powerful new tool for studying the magnetic patterns on magnetic recording media. This was accomplished by modifying a conventional scanning tunneling microscope. The fine-wire probe that is used to image surface topography was replaced with a flexible magnetic probe. Images obtained with these probes reveal both the surface topography and the magnetic structure. We have made a thorough theoretical analysis of the interaction between the probe and the magnetic fields emanating from a typical recorded surface. Quantitative data about the constituent magnetic fields can then be obtained. We have employed these techniques in studies of two of the most important issues of magnetic record: data overwrite and maximizing data-density. These studies have shown: (1) overwritten data can be retrieved under certain conditions; and (2) improvements in data-density will require new magnetic materials. In the course of these studies we have developed new techniques to analyze magnetic fields of recorded media. These studies are both theoretical and experimental and combined with the use of our magnetic force scanning tunneling microscope should lead to further breakthroughs in the field of magnetic recording.

  6. Tunneling in Superconductors

    NASA Astrophysics Data System (ADS)

    Giaever, Ivar

    2002-03-01

    It has been said that Thomas Edison's greatest invention was that of the "Research Laboratory" as a social institution. My greatest discovery was when I learned at 29 years of age that it was possible to work in such an institution and get paid for doing research. I had become interested in physics, gotten a job at General Electric Research Laboratory and found a great mentor in John C. Fischer, who besides instructing me in physics told me that sooner or later we all would become historians of science. I guess for me that time is now, because I have been asked to tell you about my second greatest discovery: Tunneling in superconductors. My great fortune was to be at the right place at the right time, where I had access to outstanding and helpful (not necessary an oxymoron) physicists. Hopefully I will be able to convey to you some of the fun and excitement of that area in this recollection. If you become real interested you may find a written version in my Nobel Prize talk: "Electron Tunneling and Superconductivity" Les Prix Nobel en 1973 or Science 183, 1253-1258 1974 or Reviews of Modern Physics 46 (2), 245-250 1974

  7. 15-Foot Spin Tunnel

    NASA Technical Reports Server (NTRS)

    1935-01-01

    A researcher is launching a model into the tunnel airstream of the 15-Foot Spin Tunnel. Charles Zimmerman wrote in NASA TR No. 557: 'After the observations have been made, the model is lowered into a net held in the air stream by one of the operators or into a large bowl-shaped net at the bottom of the test section. When lowered into the large net, the model is retrieved with a long- handled clamp.' (p. 267) 'The models used are generally 1/10 to 1/16 scale. The size of the models is limited by the wing span and the wing loading. The maximum allowable span is about 36 inches; the maximum wing loading is about 1.3 pounds per square foot.' (p. 266) 'Balsa wood is the usual structural material because of its low density. It is necessary to hollow out the after portion of the fuselage and to cut out a large portion of the wood in the wings to permit proper mass distribution. The wing cut-outs are covered with silk tissue paper. The leading and trailing edges and tips of the wings are fitted with strips of spruce, pattern pine, or bamboo inset into the edge of the balsa to prevent disfigurement from accidental blows or from striking the safety netting. Lead is used for ballast.' (p. 266)

  8. Natural Rolling Responses of a Delta Wing in Transonic and Subsonic Flows

    NASA Technical Reports Server (NTRS)

    Menzies, Margaret A.; Kandil, Osama A.

    1996-01-01

    The unsteady, three-dimensional, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the natural rolling response of slender delta wings of zero thickness at moderate to high angles of attack, to transonic and subsonic flows. The governing equations of fluid flow and dynamics of the present multi-disciplinary problem are solved using the time-accurate solution of the NS equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The main focus is to analyze the effect of Mach number and angle of attack on the leading edge vortices and their breakdown, the resultant rolling motion, and overall aerodynamic response of the wing. Three cases demonstrate the natural response of a 65 deg swept, cropped delta wing in a transonic flow with breakdown of the leading edge vortices and an 80 deg swept delta wing in a subsonic flow undergoing either damped or self-excited limit-cycle rolling oscillations as a function of angle of attack. Comparisons with an experimental investigation completes this study, validating the analysis and illustrating the complex details afforded by computational investigations.

  9. Calculation of subsonic and supersonic steady and unsteady aerodynamic forces using velocity potential aerodynamic elements

    NASA Technical Reports Server (NTRS)

    Haviland, J. K.; Yoo, Y. S.

    1976-01-01

    Expressions for calculation of subsonic and supersonic, steady and unsteady aerodynamic forces are derived, using the concept of aerodynamic elements applied to the downwash velocity potential method. Aerodynamic elements can be of arbitrary out of plane polygon shape, although numerical calculations are restricted to rectangular elements, and to the steady state case in the supersonic examples. It is suggested that the use of conforming, in place of rectangular elements, would give better results. Agreement with results for subsonic oscillating T tails is fair, but results do not converge as the number of collocation points is increased. This appears to be due to the form of expression used in the calculations. The methods derived are expected to facilitate automated flutter analysis on the computer. In particular, the aerodynamic element concept is consistent with finite element methods already used for structural analysis. The method is universal for the complete Mach number range, and, finally, the calculations can be arranged so that they do not have to be repeated completely for every reduced frequency.

  10. Effect of Initial Condition on Subsonic Jet Noise from Two Rectangular Nozzles

    NASA Technical Reports Server (NTRS)

    Zaman, K. B. M. Q.

    2012-01-01

    Differences in jet noise data from two small 8:1 aspect ratio nozzles are investigated experimentally. The interiors of the two nozzles are identical but one has a thin-lip at the exit while the has a perpendicular face at the exit (thick-lip). It is found that the thin-lip nozzle is substantially noisier throughout the subsonic Mach number range. As much as 5dB difference in OASPL is noticed around Mj =0.96. Hot-wire measurements are carried out for the characteristics of the exit boundary layer and, overall, the noise difference can be ascribed to differences in the boundary layer state. The boundary layer of the quieter (thick-lip) nozzle goes through transition around M(sub j) =0.25 and at higher M(sub j) it remains "nominally turbulent". In comparison, the boundary layer of the thin-lip nozzle is found to remain "nominally laminar". at high subsonic conditions. The nominally laminar state involves significantly larger turbulence intensities commensurate with the higher radiated noise.

  11. Electron collection theory for a D-region subsonic blunt electrostatic probe

    NASA Technical Reports Server (NTRS)

    Wai-Kwong Lai, T.

    1974-01-01

    Blunt probe theory for subsonic flow in a weakly ionized and collisional gas is reviewed, and an electron collection theory for the relatively unexplored case, Deybye length approximately 1, which occurs in the lower ionosphere (D-region), is developed. It is found that the dimensionless Debye length is no longer an electric field screening parameter, and the space charge field effect can be negelected. For ion collection, Hoult-Sonin theory is recognized as a correct description of the thin, ion density-perturbed layer adjacent the blunt probe surface. The large volume with electron density perturbed by a positively biased probe renders the usual thin boundary layer analysis inapplicable. Theories relating free stream conditions to the electron collection rate for both stationary and moving blunt probes are obtained. A model based on experimental nonlinear electron drift velocity data is proposed. For a subsonically moving probe, it is found that the perturbed region can be divided into four regions with distinct collection mechanisms.

  12. Technologies and Concepts for Reducing the Fuel Burn of Subsonic Transport Aircraft

    NASA Technical Reports Server (NTRS)

    Nickol, Craig L.

    2012-01-01

    There are many technologies under development that have the potential to enable large fuel burn reductions in the 2025 timeframe for subsonic transport aircraft relative to the current fleet. This paper identifies a potential technology suite and analyzes the fuel burn reduction potential of these technologies when integrated into advanced subsonic transport concepts. Advanced tube-and-wing concepts are developed in the single aisle and large twin aisle class, and a hybrid-wing-body concept is developed for the large twin aisle class. The resulting fuel burn reductions for the advanced tube-and-wing concepts range from a 42% reduction relative to the 777-200 to a 44% reduction relative to the 737-800. In addition, the hybrid-wingbody design resulted in a 47% fuel burn reduction relative to the 777-200. Of course, to achieve these fuel burn reduction levels, a significant amount of technology and concept maturation is required between now and 2025. A methodology for capturing and tracking concept maturity is also developed and presented in this paper.

  13. Engine Seal Technology Requirements to Meet NASA's Advanced Subsonic Technology Program Goals

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.; Hendricks, Robert C.

    1994-01-01

    Cycle studies have shown the benefits of increasing engine pressure ratios and cycle temperatures to decrease engine weight and improve performance of commercial turbine engines. NASA is working with industry to define technology requirements of advanced engines and engine technology to meet the goals of NASA's Advanced Subsonic Technology Initiative. As engine operating conditions become more severe and customers demand lower operating costs, NASA and engine manufacturers are investigating methods of improving engine efficiency and reducing operating costs. A number of new technologies are being examined that will allow next generation engines to operate at higher pressures and temperatures. Improving seal performance - reducing leakage and increasing service life while operating under more demanding conditions - will play an important role in meeting overall program goals of reducing specific fuel consumption and ultimately reducing direct operating costs. This paper provides an overview of the Advanced Subsonic Technology program goals, discusses the motivation for advanced seal development, and highlights seal technology requirements to meet future engine performance goals.

  14. On fluttering modes for aircraft wing model in subsonic air flow.

    PubMed

    Shubov, Marianna A

    2014-12-08

    The paper deals with unstable aeroelastic modes for aircraft wing model in subsonic, incompressible, inviscid air flow. In recent author's papers asymptotic, spectral and stability analysis of the model has been carried out. The model is governed by a system of two coupled integrodifferential equations and a two-parameter family of boundary conditions modelling action of self-straining actuators. The Laplace transform of the solution is given in terms of the 'generalized resolvent operator', which is a meromorphic operator-valued function of the spectral parameter λ, whose poles are called the aeroelastic modes. The residues at these poles are constructed from the corresponding mode shapes. The spectral characteristics of the model are asymptotically close to the ones of a simpler system, which is called the reduced model. For the reduced model, the following result is shown: for each value of subsonic speed, there exists a radius such that all aeroelastic modes located outside the circle of this radius centred at zero are stable. Unstable modes, whose number is always finite, can occur only inside this 'circle of instability'. Explicit estimate of the 'instability radius' in terms of model parameters is given.

  15. On fluttering modes for aircraft wing model in subsonic air flow

    PubMed Central

    Shubov, Marianna A.

    2014-01-01

    The paper deals with unstable aeroelastic modes for aircraft wing model in subsonic, incompressible, inviscid air flow. In recent author’s papers asymptotic, spectral and stability analysis of the model has been carried out. The model is governed by a system of two coupled integrodifferential equations and a two-parameter family of boundary conditions modelling action of self-straining actuators. The Laplace transform of the solution is given in terms of the ‘generalized resolvent operator’, which is a meromorphic operator-valued function of the spectral parameter λ, whose poles are called the aeroelastic modes. The residues at these poles are constructed from the corresponding mode shapes. The spectral characteristics of the model are asymptotically close to the ones of a simpler system, which is called the reduced model. For the reduced model, the following result is shown: for each value of subsonic speed, there exists a radius such that all aeroelastic modes located outside the circle of this radius centred at zero are stable. Unstable modes, whose number is always finite, can occur only inside this ‘circle of instability’. Explicit estimate of the ‘instability radius’ in terms of model parameters is given. PMID:25484610

  16. Source characterization of a subsonic jet by using near-field acoustical holography.

    PubMed

    Lee, Moohyung; Bolton, J Stuart

    2007-02-01

    In the present study, patch near-field acoustical holography was used in conjunction with a multireference, cross-spectral sound pressure measurement to visualize the sound field emitted by a subsonic jet and to predict its farfield radiation pattern. A strategy for microphone array design is described that accounts for the low spatial coherence of aeroacoustic sources and for microphone self-noise resulting from entrained flow near the jet. In the experiments, a 0.8-cm-diameter burner was used to produce a subsonic, turbulent jet with a Mach number of 0.26. Six fixed, linear arrays holding eight reference microphones apiece were disposed circumferentially around the jet, and a circular array holding sixteen, equally spaced field microphones was traversed along the jet axis to measure the sound field on a 30-cm-diameter cylindrical surface enclosing the jet. The results revealed that the jet could be modeled as a combination of eleven uncorrelated dipole-, quadrupole-, and octupole-like sources, and the contribution of each source type to the total radiated sound power could be identified. Both the total sound field reconstructed in a three-dimensional space and the farfield radiation directivity obtained by using the latter model were successfully validated by comparisons to directly measured results.

  17. Validation of a reduced-order jet model for subsonic and underexpanded hydrogen jets

    DOE PAGES

    Li, Xuefang; Hecht, Ethan S.; Christopher, David M.

    2016-01-01

    Much effort has been made to model hydrogen releases from leaks during potential failures of hydrogen storage systems. A reduced-order jet model can be used to quickly characterize these flows, with low computational cost. Notional nozzle models are often used to avoid modeling the complex shock structures produced by the underexpanded jets by determining an “effective” source to produce the observed downstream trends. In our work, the mean hydrogen concentration fields were measured in a series of subsonic and underexpanded jets using a planar laser Rayleigh scattering system. Furthermore, we compared the experimental data to a reduced order jet modelmore » for subsonic flows and a notional nozzle model coupled to the jet model for underexpanded jets. The values of some key model parameters were determined by comparisons with the experimental data. Finally, the coupled model was also validated against hydrogen concentrations measurements for 100 and 200 bar hydrogen jets with the predictions agreeing well with data in the literature.« less

  18. Acoustic radiation damping of flat rectangular plates subjected to subsonic flows

    NASA Technical Reports Server (NTRS)

    Lyle, Karen Heitman

    1993-01-01

    The acoustic radiation damping for various isotropic and laminated composite plates and semi-infinite strips subjected to a uniform, subsonic and steady flow has been predicted. The predictions are based on the linear vibration of a flat plate. The fluid loading is characterized as the perturbation pressure derived from the linearized Bernoulli and continuity equations. Parameters varied in the analysis include Mach number, mode number and plate size, aspect ratio and mass. The predictions are compared with existing theoretical results and experimental data. The analytical results show that the fluid loading can significantly affect realistic plate responses. Generally, graphite/epoxy and carbon/carbon plates have higher acoustic radiation damping values than similar aluminum plates, except near plate divergence conditions resulting from aeroelastic instability. Universal curves are presented where the acoustic radiation damping normalized by the mass ratio is a linear function of the reduced frequency. A separate curve is required for each Mach number and plate aspect ratio. In addition, acoustic radiation damping values can be greater than or equal to the structural component of the modal critical damping ratio (assumed as 0.01) for the higher subsonic Mach numbers. New experimental data were acquired for comparison with the analytical results.

  19. Tunneling in thin MOS structures

    NASA Technical Reports Server (NTRS)

    Maserjian, J.

    1974-01-01

    Recent results on tunneling in thin MOS structures are described. Thermally grown SiO2 films in the thickness range of 22-40 A have been shown to be effectively uniform on an atomic scale and exhibit an extremely abrupt oxide-silicon interface. Resonant reflections are observed at this interface for Fowler-Nordheim tunneling and are shown to agree with the exact theory for a trapezoidal barrier. Tunneling at lower fields is consistent with elastic tunneling into the silicon direct conduction band and, at still lower fields, inelastic tunneling into the indirect conduction band. Approximate dispersion relations are obtained over portions of the silicon-dioxide energy gap and conduction band.

  20. Modification of the Ames 40- by 80-foot wind tunnel for component acoustic testing for the second generation supersonic transport

    NASA Technical Reports Server (NTRS)

    Schmitz, F. H.; Allmen, J. R.; Soderman, P. T.

    1994-01-01

    The development of a large-scale anechoic test facility where large models of engine/airframe/high-lift systems can be tested for both improved noise reduction and minimum performance degradation is described. The facility development is part of the effort to investigate economically viable methods of reducing second generation high speed civil transport noise during takeoff and climb-out that is now under way in the United States. This new capability will be achieved through acoustic modifications of NASA's second largest subsonic wind tunnel: the 40-by 80-Foot Wind Tunnel at the NASA Ames Research Center. Three major items are addressed in the design of this large anechoic and quiet wind tunnel: a new deep (42 inch (107 cm)) test section liner, expansion of the wind tunnel drive operating envelope at low rpm to reduce background noise, and other promising methods of improving signal-to-noise levels of inflow microphones. Current testing plans supporting the U.S. high speed civil transport program are also outlined.