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  • Inorganic Chemistry  (3,890)
  • Aerodynamics
  • Ertrag
  • LUNAR AND PLANETARY EXPLORATION
  • 2005-2009  (341)
  • 1930-1934  (3,908)
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  • 101
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    Multibody Modeling and Simulation for the Mars Phoenix Lander Entry, Descent and Landing (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A multi-body flight simulation for the Phoenix Mars Lander has been developed that includes high fidelity six degree-of-freedom rigid-body models for the parachute and lander system. The simulation provides attitude and rate history predictions of all bodies throughout the flight, as well as loads on each of the connecting lines. In so doing, a realistic behavior of the descending parachute/lander system dynamics can be simulated that allows assessment of the Phoenix descent performance and identification of potential sensitivities for landing. This simulation provides a complete end-to-end capability of modeling the entire entry, descent, and landing sequence for the mission. Time histories of the parachute and lander aerodynamic angles are presented. The response of the lander system to various wind models and wind shears is shown to be acceptable. Monte Carlo simulation results are also presented.
    Keywords: Aerodynamics
    Type: AIAA Atmospheric Flight Mechanics Conference; Aug 18, 2008 - Aug 21, 2008; Honolulu, HI; United States
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  • 102
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    Control Laws for a Dual-Spin Stabilized Platform (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes two attitude control laws suitable for atmospheric flight vehicles with a steady angular momentum bias in the vehicle yaw axis. This bias is assumed to be provided by an internal flywheel, and is introduced to enhance roll and pitch stiffness. The first control law is based on Lyapunov stability theory, and stability proofs are given. The second control law, which assumes that the angular momentum bias is large, is based on a classical PID control. It is shown that the large yaw-axis bias requires that the PI feedback component on the roll and pitch angle errors be cross-fed. Both control laws are applied to a vehicle simulation in the presence of disturbances for several values of yaw-axis angular momentum bias. It is seen that both control laws provide a significant improvement in attitude performance when the bias is sufficiently large, but the nonlinear control law is also able to provide improved performance for a small value of bias. This is important because the smaller bias corresponds to a smaller requirement for mass to be dedicated to the flywheel.
    Keywords: Aerodynamics
    Type: AIAA Guidance, Navigation and Control Conference and Exhibit; Aug 18, 2008 - Aug 21, 2008; Honolulu, HI; United States
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  • 103
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    Unsteady Analysis of Blade and Tip Heat Transfer as Influenced by the Upstream Momentum and Thermal Wakes (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effect of the upstream wake on the blade heat transfer has been numerically examined. The geometry and the flow conditions of the first stage turbine blade of GE s E3 engine with a tip clearance equal to 2 percent of the span was utilized. Based on numerical calculations of the vane, a set of wake boundary conditions were approximated, which were subsequently imposed upon the downstream blade. This set consisted of the momentum and thermal wakes as well as the variation in modeled turbulence quantities of turbulence intensity and the length scale. Using a one-blade periodic domain, the distributions of unsteady heat transfer rate on the turbine blade and its tip, as affected by the wake, were determined. Such heat transfer coefficient distribution was computed using the wall heat flux and the adiabatic wall temperature to desensitize the heat transfer coefficient to the wall temperature. For the determination of the wall heat flux and the adiabatic wall temperatures, two sets of computations were required. The results were used in a phase-locked manner to compute the unsteady or steady heat transfer coefficients. It has been found that the unsteady wake has some effect on the distribution of the time averaged heat transfer coefficient on the blade and that this distribution is different from the distribution that is obtainable from a steady computation. This difference was found to be as large as 20 percent of the average heat transfer on the blade surface. On the tip surface, this difference is comparatively smaller and can be as large as four percent of the average.
    Keywords: Aerodynamics
    Type: NASA/TM-2008-215257 , GT2008-51242 , E-16520 , 2008 Expo 2008 Gas Turbine Technical Congress and Exposition; Jun 01, 2008; Berlin; Germany
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  • 104
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    Parametric Study of High Frequency Pulse Detonation Tubes (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes development of high frequency pulse detonation tubes similar to a small pulse detonation engine (PDE). A high-speed valve injects a charge of a mixture of fuel and air at rates of up to 1000 Hz into a constant area tube closed at one end. The reactants detonate in the tube and the products exit as a pulsed jet. High frequency pressure transducers are used to monitor the pressure fluctuations in the device and thrust is measured with a balance. The effects of injection frequency, fuel and air flow rates, tube length, and injection location are considered. Both H2 and C2H4 fuels are considered. Optimum (maximum specific thrust) fuel-air compositions and resonant frequencies are identified. Results are compared to PDE calculations. Design rules are postulated and applications to aerodynamic flow control and propulsion are discussed.
    Keywords: Aerodynamics
    Type: 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 20, 2008 - Jul 23, 2008; Hartford, CT; United States
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  • 105
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    Overview of CFD Validation Experiments for Circulation Control Applications at NASA (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Circulation control is a viable active flow control approach that can be used to meet the NASA Subsonic Fixed Wing project s Cruise Efficient Short Take Off and Landing goals. Currently, circulation control systems are primarily designed using empirical methods. However, large uncertainty in our ability to predict circulation control performance has led to the development of advanced CFD methods. This paper provides an overview of a systematic approach to developing CFD tools for basic and advanced circulation control applications. This four-step approach includes "Unit", "Benchmar", "Subsystem", and "Complete System" experiments. The paper emphasizes the ongoing and planned 2-D and 3-D physics orientated experiments with corresponding CFD efforts. Sample data are used to highlight the challenges involved in conducting circulation control computations and experiments.
    Keywords: Aerodynamics
    Type: 2008 International Powered Lift Conference; Jul 22, 2008 - Jul 24, 2008; London; United Kingdom
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  • 106
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    Wind Tunnel Testing of Powered Lift, All-Wing STOL Model (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: 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.
    Keywords: Aerodynamics
    Type: 2008 International Powered Lift Conference; Jul 22, 2008 - Jul 24, 2008; London; United Kingdom
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  • 107
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    Detection and Characterization of Boundary-Layer Transition in Flight at Supersonic Conditions Using Infrared Thermography (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Infrared thermography is a powerful tool for investigating fluid mechanics on flight vehicles. (Can be used to visualize and characterize transition, shock impingement, separation etc.). Updated onboard F-15 based system was used to visualize supersonic boundary layer transition test article. (Tollmien-Schlichting and cross-flow dominant flow fields). Digital Recording improves image quality and analysis capability. (Allows accurate quantitative (temperature) measurements, Greater enhancement through image processing allows analysis of smaller scale phenomena).
    Keywords: Aerodynamics
    Type: 13th International Symposium on Flow Visualization; Jul 01, 2008 - Jul 04, 2008; Nice; France
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  • 108
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    Volume Dynamics Propulsion System Modeling for Supersonics Vehicle Research (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Under the NASA Fundamental Aeronautics Program, the Supersonics Project is working to overcome the obstacles to supersonic commercial flight. The proposed vehicles are long slim body aircraft with pronounced aero-servo-elastic modes. These modes can potentially couple with propulsion system dynamics; leading to performance challenges such as aircraft ride quality and stability. Other disturbances upstream of the engine generated from atmospheric wind gusts, angle of attack, and yaw can have similar effects. In addition, for optimal propulsion system performance, normal inlet-engine operations are required to be closer to compressor stall and inlet unstart. To study these phenomena an integrated model is needed that includes both airframe structural dynamics as well as the propulsion system dynamics. This paper covers the propulsion system component volume dynamics modeling of a turbojet engine that will be used for an integrated vehicle Aero-Propulso-Servo-Elastic model and for propulsion efficiency studies.
    Keywords: Aerodynamics
    Type: NASA/TM--2008-215172 , E-16415 , GT2008-50524 , ASME Turbo Expo; Jun 09, 2008 - Jun 13, 2008; Berlin; Germany
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  • 109
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    Modeling State-Space Aeroelastic Systems Using a Simple Matrix Polynomial Approach for the Unsteady Aerodynamics (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A simple matrix polynomial approach is introduced for approximating unsteady aerodynamics in the s-plane and ultimately, after combining matrix polynomial coefficients with matrices defining the structure, a matrix polynomial of the flutter equations of motion (EOM) is formed. A technique of recasting the matrix-polynomial form of the flutter EOM into a first order form is also presented that can be used to determine the eigenvalues near the origin and everywhere on the complex plane. An aeroservoelastic (ASE) EOM have been generalized to include the gust terms on the right-hand side. The reasons for developing the new matrix polynomial approach are also presented, which are the following: first, the "workhorse" methods such as the NASTRAN flutter analysis lack the capability to consistently find roots near the origin, along the real axis or accurately find roots farther away from the imaginary axis of the complex plane; and, second, the existing s-plane methods, such as the Roger s s-plane approximation method as implemented in ISAC, do not always give suitable fits of some tabular data of the unsteady aerodynamics. A method available in MATLAB is introduced that will accurately fit generalized aerodynamic force (GAF) coefficients in a tabular data form into the coefficients of a matrix polynomial form. The root-locus results from the NASTRAN pknl flutter analysis, the ISAC-Roger's s-plane method and the present matrix polynomial method are presented and compared for accuracy and for the number and locations of roots.
    Keywords: Aerodynamics
    Type: RTO-AVT-154 , NATO RTO Specialists Meeting AVT-154 on Advanced Methods in Aeroelasticity; May 05, 2008 - May 07, 2008; Norway; Norway
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  • 110
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    Development of Advanced Computational Aeroelasticity Tools at NASA Langley Research Center (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA Langley Research Center has continued to develop its long standing computational tools to address new challenges in aircraft and launch vehicle design. This paper discusses the application and development of those computational aeroelastic tools. Four topic areas will be discussed: 1) Modeling structural and flow field nonlinearities; 2) Integrated and modular approaches to nonlinear multidisciplinary analysis; 3) Simulating flight dynamics of flexible vehicles; and 4) Applications that support both aeronautics and space exploration.
    Keywords: Aerodynamics
    Type: AVT-154-003 , NATO RTO Specialists'' Meeting AVT-154 on Advanced Methods in Aeroelasticity; May 05, 2008 - May 07, 2008; Norway; Norway
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  • 111
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    Aeroacoustic Simulations of Tandem Cylinders with Subcritical Spacing (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Tandem cylinders are being studied because they model a variety of component level interactions of landing gear. The present effort is directed at the case of two identical cylinders with their centroids separated in the streamwise direction by 1.435 diameters. Experiments in the Basic Aerodynamic Research Tunnel and Quiet Flow Facility at NASA Langley Research Center have provided an extensive experimental database of the nearfield flow and radiated noise. The measurements were conducted at a Mach number of 0.1285 and Reynolds number of 1.66x10(exp 5) based on the cylinder diameter. A trip was used on the upstream cylinder to insure a fully turbulent flow separation and, hence, to simulate a major aspect of high Reynolds number flow. The parallel computational effort uses the three-dimensional Navier-Stokes solver CFL3D with a hybrid, zonal turbulence model that turns off the turbulence production term everywhere except in a narrow ring surrounding solid surfaces. The experiments exhibited an asymmetry in the surface pressure that was persistent despite attempts to eliminate it through small changes in the configuration. To model the asymmetry, the simulations were run with the cylinder configuration at a nonzero but small angle of attack. The computed results and experiments are in general agreement that vortex shedding for the spacing studied herein is weak relative to that observed at supercritical spacings. Although the shedding was subdued in the simulations, it was still more prominent than in the experiments. Overall, the simulation comparisons with measured near-field data and the radiated acoustics are reasonable, especially if one is concerned with capturing the trends relative to larger cylinder spacings. However, the flow details of the 1.435 diameter spacing have not been captured in full even though very fine grid computations have been performed. Some of the discrepancy may be associated with the simulation s inexact representation of the experimental configuration, but numerical and flow modeling errors are also likely contributors to the observed differences.
    Keywords: Aerodynamics
    Type: AIAA Paper 2008-2862 , 14th AIAA/CEAS Aeroacoustics Conference; May 05, 2008 - May 07, 2008; Vancouver; Canada
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  • 112
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    X-43A Flight-Test-Determined Aerodynamic Force and Moment Characteristics at Mach 7.0 (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets, frequency sweeps, and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80-0.95 and at altitudes from 92,000 ft mean sea level to sea level. The dynamic pressure varied from 1300 to 400 psf with the angle of attack ranging from 0 to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel test data. The X-43A flight-derived axial force was found to be 10-15%higher than prediction. Underpredictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.
    Keywords: Aerodynamics
    Type: AIAA Paper-2006-8028 , 14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference; Nov 06, 2006 - Nov 09, 2006; Canberra; Australia|Journal of Spacecraft and Rockets; 45; 3; 472-484
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  • 113
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    Initial Experiments and Analysis of Blunt-Edge Vortex Flows (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A review is presented of the initial experimental results and analysis that formed the basis the Vortex Flow Experiment 2 (VFE-2). The focus of this work was to distinguish the basic effects of Reynolds number, Mach number, angle of attack, and leading edge bluntness on separation-induced leading-edge vortex flows that are common to slender wings. Primary analysis is focused on detailed static surface pressure distributions, and the results demonstrate significant effects regarding the onset and progression of leading-edge vortex separation.
    Keywords: Aerodynamics
    Type: AIAA Paper-2008-0378 , 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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  • 114
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    What Was Learned from the New VFE-2 Experiments? (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In the present paper the main results of the new experiments from Vortex Flow Experiment (VFE-2) are summarized. These include some force and moment results, surface and off-body measurements, as well as steady and fluctuating quantities. Some critical remarks are added, and an outlook for future investigations is given.
    Keywords: Aerodynamics
    Type: AIAA Paper-2008-0383 , 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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  • 115
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    Fluorescence Imaging Study of Impinging Underexpanded Jets (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An experiment was designed to create a simplified simulation of the flow through a hole in the surface of a hypersonic aerospace vehicle and the subsequent impingement of the flow on internal structures. In addition to planar laser-induced fluorescence (PLIF) flow visualization, pressure measurements were recorded on the surface of an impingement target. The PLIF images themselves provide quantitative spatial information about structure of the impinging jets. The images also help in the interpretation of impingement surface pressure profiles by highlighting the flow structures corresponding to distinctive features of these pressure profiles. The shape of the pressure distribution along the impingement surface was found to be double-peaked in cases with a sufficiently high jet-exit-to-ambient pressure ratio so as to have a Mach disk, as well as in cases where a flow feature called a recirculation bubble formed at the impingement surface. The formation of a recirculation bubble was in turn found to depend very sensitively upon the jet-exit-to-ambient pressure ratio. The pressure measured at the surface was typically less than half the nozzle plenum pressure at low jet pressure ratios and decreased with increasing jet pressure ratios. Angled impingement cases showed that impingement at a 60deg angle resulted in up to a factor of three increase in maximum pressure at the plate compared to normal incidence.
    Keywords: Aerodynamics
    Type: 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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  • 116
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    NASA Perspective on Requirements for Development of Advanced Methods Predicting Unsteady Aerodynamics and Aeroelasticity (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Over the past three years, the National Aeronautics and Space Administration (NASA) has initiated design, development, and testing of a new human-rated space exploration system under the Constellation Program. Initial designs within the Constellation Program are scheduled to replace the present Space Shuttle, which is slated for retirement within the next three years. The development of vehicles for the Constellation system has encountered several unsteady aerodynamics challenges that have bearing on more traditional unsteady aerodynamic and aeroelastic analysis. This paper focuses on the synergy between the present NASA challenges and the ongoing challenges that have historically been the subject of research and method development. There are specific similarities in the flows required to be analyzed for the space exploration problems and those required for some of the more nonlinear unsteady aerodynamic and aeroelastic problems encountered on aircraft. The aggressive schedule, significant technical challenge, and high-priority status of the exploration system development is forcing engineers to implement existing tools and techniques in a design and application environment that is significantly stretching the capability of their methods. While these methods afford the users with the ability to rapidly turn around designs and analyses, their aggressive implementation comes at a price. The relative immaturity of the techniques for specific flow problems and the inexperience with their broad application to them, particularly on manned spacecraft flight system, has resulted in the implementation of an extensive wind tunnel and flight test program to reduce uncertainty and improve the experience base in the application of these methods. This provides a unique opportunity for unsteady aerodynamics and aeroelastic method developers to test and evaluate new analysis techniques on problems with high potential for acquisition of test and even flight data against which they can be evaluated. However, researchers may be required to alter the geometries typically used in their analyses, the types of flows analyzed, and even the techniques by which computational tools are verified and validated. This paper discusses these issues and provides some perspective on the potential for new and innovative approaches to the development of methods to attack problems in nonlinear unsteady aerodynamics.
    Keywords: Aerodynamics
    Type: AVT-154-002 , NATO/RTO AVT-154 Advanced Methods in Aeroelasticity; May 05, 2008 - May 08, 2008; Oslo; Norway
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  • 117
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    Assessment of 3D Codes for Predicting Liner Attenuation in Flow Ducts (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents comparisons of seven propagation codes for predicting liner attenuation in ducts with flow. The selected codes span the spectrum of methods available (finite element, parabolic approximation, and pseudo-time domain) and are collectively representative of the state-of-art in the liner industry. These codes are included because they have two-dimensional and three-dimensional versions and can be exported to NASA's Columbia Supercomputer. The basic assumptions, governing differential equations, boundary conditions, and numerical methods underlying each code are briefly reviewed and an assessment is performed based on two predefined metrics. The two metrics used in the assessment are the accuracy of the predicted attenuation and the amount of wall clock time to predict the attenuation. The assessment is performed over a range of frequencies, mean flow rates, and grazing flow liner impedances commonly used in the liner industry. The primary conclusions of the study are (1) predicted attenuations are in good agreement for rigid wall ducts, (2) the majority of codes compare well to each other and to approximate results from mode theory for soft wall ducts, (3) most codes compare well to measured data on a statistical basis, (4) only the finite element codes with cubic Hermite polynomials capture extremely large attenuations, and (5) wall clock time increases by an order of magnitude or more are observed for a three-dimensional code relative to the corresponding two-dimensional version of the same code.
    Keywords: Aerodynamics
    Type: 14th AIAA/CEAS Aeroacoustics Conference; May 05, 2008 - May 08, 2008; Vancouver, BC; Canada
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  • 118
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    Infrared Imaging of Boundary Layer Transition Flight Experiments (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Hypersonic Thermodynamic Infrared Measurement (HYTHIRM) project is presently focused on near term support to the Shuttle program through the development of an infrared imaging capability of sufficient spatial and temporal resolution to augment existing on-board Orbiter instrumentation. Significant progress has been made with the identification and inventory of relevant existing optical imaging assets and the development, maturation, and validation of simulation and modeling tools for assessment and mission planning purposes, which were intended to lead to the best strategies and assets for successful acquisition of quantitative global surface temperature data on the Shuttle during entry. However, there are longer-term goals of providing global infrared imaging support to other flight projects as well. A status of HYTHIRM from the perspective of how two NASA-sponsored boundary layer transition flight experiments could benefit by infrared measurements is provided. Those two flight projects are the Hypersonic Boundary layer Transition (HyBoLT) flight experiment and the Shuttle Boundary Layer Transition Flight Experiment (BLT FE), which are both intended for reducing uncertainties associated with the extrapolation of wind tunnel derived transition correlations for flight application. Thus, the criticality of obtaining high quality flight data along with the impact it would provide to the Shuttle program damage assessment process are discussed. Two recent wind tunnel efforts that were intended as risk mitigation in terms of quantifying the transition process and resulting turbulent wedge locations are briefly reviewed. Progress is being made towards finalizing an imaging strategy in support of the Shuttle BLT FE, however there are no plans currently to image HyBoLT.
    Keywords: Aerodynamics
    Type: AIAA 2008-4027 , 40th AIAA Thermophysics Conference; Jun 23, 2008 - Jun 26, 2008; Seattle, WA; United States
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  • 119
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    Assessment and Mission Planning Capability For Quantitative Aerothermodynamic Flight Measurements Using Remote Imaging (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: High resolution calibrated infrared imagery of vehicles during hypervelocity atmospheric entry or sustained hypersonic cruise has the potential to provide flight data on the distribution of surface temperature and the state of the airflow over the vehicle. In the early 1980 s NASA sought to obtain high spatial resolution infrared imagery of the Shuttle during entry. Despite mission execution with a technically rigorous pre-planning capability, the single airborne optical system for this attempt was considered developmental and the scientific return was marginal. In 2005 the Space Shuttle Program again sponsored an effort to obtain imagery of the Orbiter. Imaging requirements were targeted towards Shuttle ascent; companion requirements for entry did not exist. The engineering community was allowed to define observation goals and incrementally demonstrate key elements of a quantitative spatially resolved measurement capability over a series of flights. These imaging opportunities were extremely beneficial and clearly demonstrated capability to capture infrared imagery with mature and operational assets of the US Navy and the Missile Defense Agency. While successful, the usefulness of the imagery was, from an engineering perspective, limited. These limitations were mainly associated with uncertainties regarding operational aspects of data acquisition. These uncertainties, in turn, came about because of limited pre-flight mission planning capability, a poor understanding of several factors including the infrared signature of the Shuttle, optical hardware limitations, atmospheric effects and detector response characteristics. Operational details of sensor configuration such as detector integration time and tracking system algorithms were carried out ad hoc (best practices) which led to low probability of target acquisition and detector saturation. Leveraging from the qualified success during Return-to-Flight, the NASA Engineering and Safety Center sponsored an assessment study focused on increasing the probability of returning spatially resolved scientific/engineering thermal imagery. This paper provides an overview of the assessment task and the systematic approach designed to establish confidence in the ability of existing assets to reliably acquire, track and return global quantitative surface temperatures of the Shuttle during entry. A discussion of capability demonstration in support of a potential Shuttle boundary layer transition flight test is presented. Successful demonstration of a quantitative, spatially resolved, global temperature measurement on the proposed Shuttle boundary layer transition flight test could lead to potential future applications with hypersonic flight test programs within the USAF and DARPA along with flight test opportunities supporting NASA s project Constellation.
    Keywords: Aerodynamics
    Type: AIAA 2008-4022 , 40th AIAA Thermophysics Conference; Jun 23, 2008 - Jun 26, 2008; Seattle, WA; United States
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  • 120
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    Boundary Layer Transition on X-43A (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The successful Mach 7 and 10 flights of the first fully integrated scramjet propulsion systems by the Hyper-X (X-43A) program have provided the means with which to verify the original design methodologies and assumptions. As part of Hyper-X s propulsion-airframe integration, the forebody was designed to include a spanwise array of vortex generators to promote boundary layer transition ahead of the engine. Turbulence at the inlet is thought to provide the most reliable engine design and allows direct scaling of flight results to groundbased data. Pre-flight estimations of boundary layer transition, for both Mach 7 and 10 flight conditions, suggested that forebody boundary layer trips were required to ensure fully turbulent conditions upstream of the inlet. This paper presents the results of an analysis of the thermocouple measurements used to infer the dynamics of the transition process during the trajectories for both flights, on both the lower surface (to assess trip performance) and the upper surface (to assess natural transition). The approach used in the analysis of the thermocouple data is outlined, along with a discussion of the calculated local flow properties that correspond to the transition events as identified in the flight data. The present analysis has confirmed that the boundary layer trips performed as expected for both flights, providing turbulent flow ahead of the inlet during critical portions of the trajectory, while the upper surface was laminar as predicted by the pre-flight analysis.
    Keywords: Aerodynamics
    Type: AIAA-2008-3736 , L-6068 , 40th AIAA Thermophysics Conference; Jun 23, 2008 - Jun 26, 2008; Seattle, WA; United States
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    Grid Quality and Resolution Issues from the Drag Prediction Workshop Series (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The drag prediction workshop series (DPW), held over the last six years, and sponsored by the AIAA Applied Aerodynamics Committee, has been extremely useful in providing an assessment of the state-of-the-art in computationally based aerodynamic drag prediction. An emerging consensus from the three workshop series has been the identification of spatial discretization errors as a dominant error source in absolute as well as incremental drag prediction. This paper provides an overview of the collective experience from the workshop series regarding the effect of grid-related issues on overall drag prediction accuracy. Examples based on workshop results are used to illustrate the effect of grid resolution and grid quality on drag prediction, and grid convergence behavior is examined in detail. For fully attached flows, various accurate and successful workshop results are demonstrated, while anomalous behavior is identified for a number of cases involving substantial regions of separated flow. Based on collective workshop experiences, recommendations for improvements in mesh generation technology which have the potential to impact the state-of-the-art of aerodynamic drag prediction are given.
    Keywords: Aerodynamics
    Type: 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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    RANS Analyses of Turbofan Nozzles with Wedge Deflectors for Noise Reduction (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Computational fluid dynamics (CFD) was used to evaluate a promising concept for reducing the noise at take-off of dual-stream, turbofan nozzles. The concept, offset stream technology, reduces the jet noise observed on the ground by diverting (offsetting) the majority of the fan flow below the core flow, thickening this layer between the high velocity core flow and the ground observers. In this study a wedge placed in the internal fan stream is used as the diverter. Wind, a Reynolds Averaged Navier-Stokes (RANS) code, was used to analyze the flowfield of the exhaust plume and to calculate nozzle performance. Results showed that the wedge effectively diverts the fan flow and the turbulent kinetic energy on the observer side of the nozzle is reduced. The reduction in turbulent kinetic energy should correspond to a reduction in noise. The blockage due to the wedge reduces the fan massflow proportional to its blockage and the overall thrust is consequently reduced. The CFD predictions are in very good agreement with experimental data. This noise reduction concept shows promise for reduced jet noise at a small reduction in thrust. It has been demonstrated that RANS CFD can be used to optimize this concept.
    Keywords: Aerodynamics
    Type: 46th AIAA Aerospace Sciences Meeting; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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    Comparison of Far-Field Noise for Three Significantly Different Model Turbofans (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Far-field noise sound power level (PWL) spectra and overall sound pressure level (OASPL) directivities were compared for three significantly different model fan stages which were tested in the NASA Glenn 9x15 Low Speed Wind Tunnel. The test fans included the Advanced Ducted Propulsor (ADP) Fan1, the baseline Source Diagnostic Test (SDT) fan, and the Quiet High Speed Fan2 (QHSF2) These fans had design rotor tangential tip speeds from 840 to 1474 ft/s and stage pressure ratios from 1.29 to 1.82. Additional parameters included rotor-stator spacing, stator sweep, and downstream support struts. Acoustic comparison points were selected on the basis of stage thrust. Acoustic results for the low tip speed/low pressure ratio fan (ADP Fan1) were thrust-adjusted to show how a geometrically-scaled version of this fan might compare at the higher design thrust levels of the other two fans. Lowest noise levels were typically observed for ADP Fan1 (which had a radial stator) and for the intermediate tip speed fan (Source Diagnostics Test, SDT, R4 rotor) with a swept stator. Projected noise levels for the ADP fan to the SDT swept stator configuration at design point conditions showed the fans to have similar noise levels. However, it is possible that the ADP fan could be 2 to 3 dB quieter with incorporation of a swept stator. Benefits of a scaled ADP fan include avoidance of multiple pure tones associated with transonic and higher blade tip speeds. Penalties of a larger size ADP fan would include increased nacelle size and drag.
    Keywords: Aerodynamics
    Type: 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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    Detailed Uncertainty Analysis for Ares I Ascent Aerodynamics Wind Tunnel Database (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A detailed uncertainty analysis for the Ares I ascent aero 6-DOF wind tunnel database is described. While the database itself is determined using only the test results for the latest configuration, the data used for the uncertainty analysis comes from four tests on two different configurations at the Boeing Polysonic Wind Tunnel in St. Louis and the Unitary Plan Wind Tunnel at NASA Langley Research Center. Four major error sources are considered: (1) systematic errors from the balance calibration curve fits and model + balance installation, (2) run-to-run repeatability, (3) boundary-layer transition fixing, and (4) tunnel-to-tunnel reproducibility.
    Keywords: Aerodynamics
    Type: AIAA Paper-2008-4259 , 26th AIAA Aerodynamic Measurement Technology & Ground Testing Conference; Jun 23, 2008 - Jun 26, 2008; Seattle, WA; United States
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    Aerothermodynamic Testing of the Crew Exploration Vehicle in the LaRC 20-Inch Mach 6 and 31-Inch Mach 10 Tunnels (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An experimental wind tunnel program is being conducted in support of a NASA wide effort to develop a Space Shuttle replacement and to support the Agency s long term objective of returning to the Moon and Mars. This report documents experimental measurements made on several scaled ceramic heat transfer models of the proposed Crew Exploration Vehicle Crew Module. The experimental data highlighted in this test report are to be used to assess numerical tools that will be used to generate the flight aerothermodynamic database. Global heat transfer images and heat transfer distributions were obtained over a range of freestream Reynolds numbers and angles of attack with the phosphor thermography technique. Heat transfer data were measured on the forebody and afterbody and were used to infer the heating on the vehicle as well as the boundary layer state on the forebody surface. Several model support configurations were assessed to minimize potential support interference. In addition, the ability of the global phosphor thermography method to provide quantitative heating measurements in the low temperature environment of the capsule base region was assessed. While naturally fully developed turbulent levels were not obtained on the forebody, the use of boundary layer trips generated fully developed turbulent flow. Laminar and turbulent computational results were shown to be in good agreement with the data. Backshell testing demonstrated the ability to obtain data in the low temperature region as well as demonstrating the lack of significant model support hardware influence on heating.
    Keywords: Aerodynamics
    Type: 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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    Aerodynamic Analysis of Simulated Heat Shield Recession for the Orion Command Module (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The aerodynamic effects of the recession of the ablative thermal protection system for the Orion Command Module of the Crew Exploration Vehicle are important for the vehicle guidance. At the present time, the aerodynamic effects of recession being handled within the Orion aerodynamic database indirectly with an additional safety factor placed on the uncertainty bounds. This study is an initial attempt to quantify the effects for a particular set of recessed geometry shapes, in order to provide more rigorous analysis for managing recession effects within the aerodynamic database. The aerodynamic forces and moments for the baseline and recessed geometries were computed at several trajectory points using multiple CFD codes, both viscous and inviscid. The resulting aerodynamics for the baseline and recessed geometries were compared. The forces (lift, drag) show negligible differences between baseline and recessed geometries. Generally, the moments show a difference between baseline and recessed geometries that correlates with the maximum amount of recession of the geometry. The difference between the pitching moments for the baseline and recessed geometries increases as Mach number decreases (and the recession is greater), and reach a value of -0.0026 for the lowest Mach number. The change in trim angle of attack increases from approx. 0.5deg at M = 28.7 to approx. 1.3deg at M = 6, and is consistent with a previous analysis with a lower fidelity engineering tool. This correlation of the present results with the engineering tool results supports the continued use of the engineering tool for future work. The present analysis suggests there does not need to be an uncertainty due to recession in the Orion aerodynamic database for the force quantities. The magnitude of the change in pitching moment due to recession is large enough to warrant inclusion in the aerodynamic database. An increment in the uncertainty for pitching moment could be calculated from these results and included in the development of the aerodynamic database uncertainty for pitching moment.
    Keywords: Aerodynamics
    Type: 46th AIAA Aerospace Sciences Meeting and Exhibit; Jan 07, 2008 - Jan 10, 2008; Reno, Nevada; United States
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    The Inception of OMA in the Development of Modal Testing Technology for Wind Turbines (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Wind turbines are immense, flexible structures with aerodynamic forces acting on the rotating blades at harmonics of the turbine rotational frequency, which are comparable to the modal frequencies of the structure. Predicting and experimentally measuring the modal frequencies of wind turbines has been important to their successful design and operation. Performing modal tests on wind turbine structures over 100 meters tall is a substantial challenge, which has inspired innovative developments in modal test technology. For wind turbines, a further complication is that the modal frequencies are dependent on the turbine rotation speed. The history and development of a new technique for acquiring the modal parameters using output-only response data, called the Natural Excitation Technique (NExT), will be reviewed, showing historical tests and techniques. The initial attempts at output-only modal testing began in the late 1980's with the development of NExT in the 1990's. NExT was a predecessor to OMA, developed to overcome these challenges of testing immense structures excited with environmental inputs. We will trace the difficulties and successes of wind turbine modal testing from 1982 to the present. Keywords: OMA, Modal Analysis, NExT, Wind Turbines, Wind Excitation
    Keywords: Aerodynamics
    Type: JSC-CN-16440 , To be published in Mechanical Systems and Signal Processing, August 2008
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    Application of Synthetic Jets to Reduce Stator Flow Separation in a Low Speed Axial Compressor (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Flow control using synthetic jet injection has been applied in a low speed axial compressor. The synthetic jets were applied from the suction surface of a stator vane via a span-wise row of slots pitched in the streamwise direction. Actuation was provided externally from acoustic drivers coupled to the vane tip via flexible tubing. The acoustic resonance characteristics of the system, and the resultant jet velocities were obtained. The effects on the separated flow field for various jet velocities and frequencies were explored. Total pressure loss reductions across the vane passage were measured. The effect of synthetic jet injection was shown to be comparable to that of pulsatory injection with mass addition for stator vanes which had separated flow. While only a weak dependence of the beneficial effect was noted based on the excitation frequency, a strong dependence on the amplitude was observed at all frequencies.
    Keywords: Aerodynamics
    Type: NASA/TM-2008-215145 , AIAA Paper-2008-0602 , E-16308
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    Aerothermodynamic Testing and Boundary Layer Trip Sizing of the HIFiRE Flight 1 Vehicle (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An experimental wind tunnel test was conducted in the NASA Langley Research Center s 20-Inch Mach 6 Air Tunnel in support of the Hypersonic International Flight Research Experimentation Program. The information in this article is focused on the Flight 1 configuration, the first in a series of flight experiments. The article documents experimental measurements made over a Reynolds numbers range of 2.1x10(exp 6)/ft to 5.6x10(exp 6)/ft and angles of attack of -5 to +5 deg on several scaled ceramic heat transfer models of the Flight 1 configuration. Global heat transfer was measured using phosphor thermography and the resulting images and heat transfer distributions were used to infer the state of the boundary layer on the vehicle windside and leeside surfaces. Boundary layer trips were used to force the boundary layer turbulent and the experimental data highlighted in this article were used to size and place the boundary layer trip for the flight vehicle. The required height of the flight boundary layer trip was determined to be 0.079 in and the trip was moved from the design location of 7.87 in to 20.47 in to ensure that augmented heating would not impact the laminar side of the vehicle. Allowable roughness was selected to be 3.2x10(exp -3) in.
    Keywords: Aerodynamics
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    Fluorescence Imaging and Streamline Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Only a few of the models survived repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2- inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various model configurations and NO seeding methods were used, including a new streamwise visualization method based on PLIF. Virtual Diagnostics Interface (ViDI) technology, developed at NASA Langley Research Center, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images.
    Keywords: Aerodynamics
    Type: LF99-5899
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    Spline-Based Smoothing of Airfoil Curvatures (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Constrained fitting for airfoil curvature smoothing (CFACS) is a splinebased method of interpolating airfoil surface coordinates (and, concomitantly, airfoil thicknesses) between specified discrete design points so as to obtain smoothing of surface-curvature profiles in addition to basic smoothing of surfaces. CFACS was developed in recognition of the fact that the performance of a transonic airfoil is directly related to both the curvature profile and the smoothness of the airfoil surface. Older methods of interpolation of airfoil surfaces involve various compromises between smoothing of surfaces and exact fitting of surfaces to specified discrete design points. While some of the older methods take curvature profiles into account, they nevertheless sometimes yield unfavorable results, including curvature oscillations near end points and substantial deviations from desired leading-edge shapes. In CFACS as in most of the older methods, one seeks a compromise between smoothing and exact fitting. Unlike in the older methods, the airfoil surface is modified as little as possible from its original specified form and, instead, is smoothed in such a way that the curvature profile becomes a smooth fit of the curvature profile of the original airfoil specification. CFACS involves a combination of rigorous mathematical modeling and knowledge-based heuristics. Rigorous mathematical formulation provides assurance of removal of undesirable curvature oscillations with minimum modification of the airfoil geometry. Knowledge-based heuristics bridge the gap between theory and designers best practices. In CFACS, one of the measures of the deviation of an airfoil surface from smoothness is the sum of squares of the jumps in the third derivatives of a cubicspline interpolation of the airfoil data. This measure is incorporated into a formulation for minimizing an overall deviation- from-smoothness measure of the airfoil data within a specified fitting error tolerance. CFACS has been extensively tested on a number of supercritical airfoil data sets generated by inverse design and optimization computer programs. All of the smoothing results show that CFACS is able to generate unbiased smooth fits of curvature profiles, trading small modifications of geometry for increasing curvature smoothness by eliminating curvature oscillations and bumps (see figure).
    Keywords: Aerodynamics
    Type: LAR-17227-1 , NASA Tech Briefs, September 2008; 40-41
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    Aerodynamic Effects of a 24-foot Multisegmented Telescoping Nose Boom on an F-15B Airplane (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An experimental multisegmented telescoping nose boom has been installed on an F-15B airplane to be tested in a flight environment. The experimental nose boom is representative of one that could be used to tailor the sonic boom signature of an airplane such as a supersonic business jet. The nose boom consists of multiple sections and could be extended during flight to a length of 24 ft. The preliminary analyses indicate that the addition of the experimental nose boom could adversely affect vehicle flight characteristics and air data systems. Before the boom was added, a series of flights was conducted to update the aerodynamic model and characterize the air data systems of the baseline airplane. The baseline results have been used in conjunction with estimates of the nose boom's influence to prepare for a series of research flights conducted with the nose boom installed. Data from these flights indicate that the presence of the experimental boom reduced the static pitch and yaw stability of the airplane. The boom also adversely affected the static-position error of the airplane but did not significantly affect angle-of-attack or angle-of-sideslip measurements. The research flight series has been successfully completed.
    Keywords: Aerodynamics
    Type: NASA/TM-2008-214634 , H-2809
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    Effects of a Rotating Aerodynamic Probe on the Flow Field of a Compressor Rotor (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An investigation of distortions of the rotor exit flow field caused by an aerodynamic probe mounted in the rotor is described in this paper. A rotor total pressure Kiel probe, mounted on the rotor hub and extending up to the mid-span radius of a rotor blade channel, generates a wake that forms additional flow blockage. Three types of high-response aerodynamic probes were used to investigate the distorted flow field behind the rotor. These probes were: a split-fiber thermo-anemometric probe to measure velocity and flow direction, a total pressure probe, and a disk probe for in-flow static pressure measurement. The signals acquired from these high-response probes were reduced using an ensemble averaging method based on a once per rotor revolution signal. The rotor ensemble averages were combined to construct contour plots for each rotor channel of the rotor tested. In order to quantify the rotor probe effects, the contour plots for each individual rotor blade passage were averaged into a single value. The distribution of these average values along the rotor circumference is a measure of changes in the rotor exit flow field due to the presence of a probe in the rotor. These distributions were generated for axial flow velocity and for static pressure.
    Keywords: Aerodynamics
    Type: NASA/CR-2008-215215 , E-16503
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    Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds (2008)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: The data contained on this CD are a supplement to NASA/TP-2001-210629 published in February 2001. This CD replaces a web-site database search and retrieval system - noted as reference 36 in the NASA/TP - that was to supply the aeronautical community with access to the flight data. Unfortunately, this web-site was only available for a short time after the publication of NASA/TP-2001-21068 due to software and support issues. The contents of this CD are organized into five folders containing data from the flight test and reference 1. In particular, the following are provided: (1) tabular data of the Flight Conditions from Table 5; (2) boundary layer data from Table 12 for three flights in multiple formats; (3) skin-friction data - xmgr format (ref. 3) - used to generate Figure 26; (4) surface pressure data with a listing of the parameters; and (5) tuft-images from three cameras in two formats.
    Keywords: Aerodynamics
    Type: NASA/TP-2001-210629/SUPPL , L19482
    Format: text
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    Successes and Challenges for Flow Control Simulations (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A survey is made of recent computations published for synthetic jet flow control cases from a CFD workshop held in 2004. The three workshop cases were originally chosen to represent different aspects of flow control physics: nominally 2-D synthetic jet into quiescent air, 3-D circular synthetic jet into turbulent boundary-layer crossflow, and nominally 2-D flow-control (both steady suction and oscillatory zero-net-mass-flow) for separation control on a simple wall-mounted aerodynamic hump shape. The purpose of this survey is to summarize the progress as related to these workshop cases, particularly noting successes and remaining challenges for computational methods. It is hoped that this summary will also by extension serve as an overview of the state-of-the-art of CFD for these types of flow-controlled flow fields in general.
    Keywords: Aerodynamics
    Type: LF99-7374
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    Fuel Savings and Aerodynamic Drag Reduction from Rail Car Covers (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The potential for energy savings by reducing the aerodynamic drag of rail cars is significant. A previous study of aerodynamic drag of coal cars suggests that a 25% reduction in drag of empty cars would correspond to a 5% fuel savings for a round trip [1]. Rail statistics for the United States [2] report that approximately 5.7 billion liters of diesel fuel were consumed for coal transportation in 2002, so a 5% fuel savings would total 284 million liters. This corresponds to 2% of Class I railroad fuel consumption nationwide. As part of a DOE-sponsored study, the aerodynamic drag of scale rail cars was measured in a wind tunnel. The goal of the study was to measure the drag reduction of various rail-car cover designs. The cover designs tested yielded an average drag reduction of 43% relative to empty cars corresponding to an estimated round-trip fuel savings of 9%.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN112 , AfricaRail 2008; Jun 02, 2008 - Jun 06, 2008; Johannesburg; South Africa
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    Overview of the LaNCETS Flight Experiment and CFD Analysis. Supplemental Movies (2008)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: This presentation focuses on nearfield airborne pressure signatures from the Lift and Nozzle Change Effect on Tail Shocks (LaNCETS) flight test experiment. The primary motivation for nearfield probing in the supersonic regime is to measure the shock structure of aircraft in an ongoing effort to overcome the overland sonic boom barrier for commercial supersonic transportation. LaNCETS provides the opportunity to investigate lift distribution and engine plume effects. During Phase 1 flight testing an F-15B was used to probe the F-15 LaNCETS aircraft in order to validate CFD and pre-flight prediction tools. A total of 29 probings were taken at 40,000 ft. altitude at Machs 1.2, 1.4 and 1.6. LaNCETS Phase 1 flight data are presented as a detailed pressure signature superimposed over a picture of the LaNCETS aircraft. The attenuation of the Inlet-Canard shocks with distance as well as its forward propagation and the coalescence of the noseboom shock are illustrated. A detailed CFD study on a simplified LaNCETS aircraft jet nozzle was performed providing the ability to more accurately capture the shocks from the propulsion system and emphasizing how under- and over-expanding the nozzle affects the existence of shock trains inside the jet plume. With Phase 1 being a success preparations are being made to move forward to Phase 2. Phase 2 will fly similar flight conditions, but this time changing the aircraft's lift distribution by biasing the canard positions, and changing the plume shape by under- and over-expanding the nozzle. Nearfield probing will again be completed in the same manner as in Phase 1. An additional presentation focuses on LaNCETS CFD solution methodology. Discussions highlight grid preprocessing, grid shearing and stretching, flow solving and contour plots. Efforts are underway to better capture the flow features via grid modification and flow solution methodology, which will help to achieve better agreement with flight data. An included CD-ROM provides animations of the nearfield probing procedure and of real data from one of the probings integrated with GPS positional and velocity data. An additional in-flight video from the rear seat of the probing aircraft is also provided.
    Keywords: Aerodynamics
    Type: Industry Panel Presentation at the University of Southern California; Nov 03, 2017; Los Angeles, CA; United States|Fundamental Aeronautics 2008 Annual Meeting; Oct 07, 2008 - Oct 09, 2008; Atlanta, GA; United States
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    CFD Results for an Axisymmetric Isentropic Relaxed Compression Inlet (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The OVERFLOW code was used to calculate the flow field for a family of five relaxed compression inlets, which were part of a screening study to determine a configuration most suited to the application of microscale flow control technology as a replacement for bleed. Comparisons are made to experimental data collected for each of the inlets in the 1- by 1-Foot Supersonic Wind Tunnel at the NASA Glenn Research Center (GRC) to help determine the suitability of computational fluid dynamics (CFD) as a tool for future studies of these inlets with flow control devices. Effects on the wind tunnel results of the struts present in a high subsonic flow region accounted for most of the inconsistency between the results. Based on the level of agreement in the present study, it is expected that CFD can be used as a tool to aid in the design of a study of this class of inlets with flow control.
    Keywords: Aerodynamics
    Type: NASA/TM-2008-215416 , AIAA Paper-2008-0092 , E-16580 , 46th AIAA Aerospace Sciences Meeting; Jan 07, 2008 - Jan 10, 2008; Reno, NV; United States
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    Adjoint-Based Methodology for Time-Dependent Optimization (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a discrete adjoint method for a broad class of time-dependent optimization problems. The time-dependent adjoint equations are derived in terms of the discrete residual of an arbitrary finite volume scheme which approximates unsteady conservation law equations. Although only the 2-D unsteady Euler equations are considered in the present analysis, this time-dependent adjoint method is applicable to the 3-D unsteady Reynolds-averaged Navier-Stokes equations with minor modifications. The discrete adjoint operators involving the derivatives of the discrete residual and the cost functional with respect to the flow variables are computed using a complex-variable approach, which provides discrete consistency and drastically reduces the implementation and debugging cycle. The implementation of the time-dependent adjoint method is validated by comparing the sensitivity derivative with that obtained by forward mode differentiation. Our numerical results show that O(10) optimization iterations of the steepest descent method are needed to reduce the objective functional by 3-6 orders of magnitude for test problems considered.
    Keywords: Aerodynamics
    Type: 12th AIAA/ISSMO Multidisciplinary Analysis and Optimization Conference; Sep 10, 2008 - Sep 12, 2008; Victoria; Canada
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    Wetterderivate als Instrument im Risikomanagement landwirtschaftlicher Betriebe (2007)
    In:  Benutzerhandbuch fur @RISK Risikoanalysen- und Simulations- Add-InRisikoanalysen Add-In für Microsoft® Excel Version 5.7, September, 2010 Palisade Corporation 798 Cascadilla Street Ithaca, NY 14850 USA
    Details
    Publication Date: 2007
    Description: Korrelationsanalyse zwischen Wettervariablen über mehrere Perioden und dem Zuckerrübenertrag an verschiedenen Standorten KATASTER-BESCHREIBUNG: Bildung von Wetterindexsummen über mehrere Monate im Vergleich zur Betrachtung einzelner Monate ergab keine höheren Korrelationen, höchste positive Korrelationskoeffizienten von 0,61 zwischen der Niederschlagssumme des Zeitraumes Juni bis August für den Standort Dedelow, in der Uckermark Brandesburgs. Einflusses der Temperatur auf den Zuckerertrag ergab negative Korrelationskoeffizienten, der höchste (negative) ermittelte Wert liegt bei -0,78 für die Periode Juli bis September, Dedelow. Positiven Einfluss hoher Temperaturen auf den Rübenertrag am Standort Kiel und Düse KATASTER-DETAIL:
    Keywords: Deutschland, teilw. einzelne Regionen ; 1958-2006 ; Zuckerrüben ; Ertrag ; Korrelationsmethode ; Niederschlag ; Temperatur
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    Kulturpflanzen von Südtirol, Nordtirol und Graubünden Literaturstudie alpiner Kulturpflanzen (2007)
    In:  2 -NAP 02-231 Bericht IV Interreg IIIA Literaturstudie alpine Kulturpflanzen Vs. 3.0 070425.
    Details
    Publication Date: 2007
    Description: Sammlung historischer Informationen und Dokumentation des bäuerlichen Erfahrungswissens Kulturpflanzen von der Prähistorie - 20. Jahrhundert KATASTER-BESCHREIBUNG: KATASTER-DETAIL:
    Keywords: Südtirol, Nordtirol und GraubündenSüdtirol, Nordtirol und Graubünden ; Kartoffeln ; Anbautermine ; Boden ; Ertrag ; Getreide ; Hafer ; Klima ; Landwirtschaft ; Mais ; Niederschlag ; Pflanzenkrankheit ; Pflanzenschädling ; Roggen ; Temperatur ; Trockenheit ; Vegetationsperiode ; Weizen ; Wetterbeobachtung ; Witterung ; Düngung ; Hackfrüchte
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  • 142
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    Simulation of a Periodic Jet in a Crossflow with a RANS Solver Using an Unstructured Grid (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: A second-order unstructured-grid code, developed and used primarily for steady aerodynamic simulations, is applied to the synthetic jet in a cross flow. The code, FUN3D, is a vertex-centered finite-volume method originally developed by Anderson[1, 2], and is currently supported by members of the Fast Adaptive Aerospace Tools team at NASA Langley. Used primarily for design[3] and analysis[4] of steady aerodynamic configurations, FUN3D incorporates a discrete adjoint capability, and supports parallel computations using MPI. A detailed description of the FUN3D code can be found in the references given above. The code is under continuous development and contains a variety of flux splitting algorithms for the inviscid terms, two methods for computing gradients, several turbulence models, and several solution methodologies; all in varying states of development. Only the most robust and reliable components, based on experiences with steady aerodynamic simulations, were employed in this work. As applied in this work, FUN3D solves the Reynolds averaged Navier-Stokes equations using the one equation turbulence model of Spalart and Allmaras[5]. The spatial discretization is formed on unstructured meshes using a vertex-centered approach. The inviscid terms are evaluated by a flux-difference splitting formulation using least-squares reconstruction and Roe-type approximate Riemann fluxes. Green-Gauss gradient evaluations are used for viscous and turbulence modeling terms. The discrete spatial operator is combined with a backward time operator which is then solved iteratively using point or line Gauss-Seidel and local time stepping in a pseudo time. For steady flows, the physical time step is set to infinity and the pseudo time step is ramped up with the iteration count. A second-order backward in time operator is used for time accurate flows with 20 to 50 steps in the pseudo time applied at each physical time step. For this effort, FUN3D was modified to support spatially varying boundary and initial conditions, and unsteady boundary conditions. Also, a specialized in/out flow boundary condition was implemented to model the action of the diaphragm. This boundary condition is described below in more detail. The grids were generated using the internally developed codes GridEX[6] for meshing the surfaces and inviscid regions of the domain, and for CAD access; and MesherX[7] for meshing the viscous regions. Grid spacing in on the surfaces and in the inviscid regions are indirectly controlled by specifying sources. The viscous layers are generated using an advancing layer technique. MeshersX allows the user to control the spatial variation of the first step off the surface, growth rates, and the termination criterion by providing small problem dependent subroutines.
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 2.6.1 - 2.6.5; NASA/CP-2007-214874
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  • 143
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    On the Prediction of Separation Bubbles Using a Modified Chen-Thyson Model (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: The prediction of separation bubbles on NACA 65-213 and NACA 0012 using a modified Chen-Thyson transition model is presented. The contents include: 1) Background; 2) Analysis of NACA 65-213 separation bubble using cebeci's viscous-inviscid interaction method; 3) Analysis of NACA 0012 separation bubble using navier-stokes method; and 4) Comparison with experiment.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 269-281; NASA/CP-2007-214667
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  • 144
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    Experiments in Transitional Boundary Layers With Emphasis on High Free-Stream Disturbance Level, Surface Concave Curvature and Strong Favorable Streamwise Pressure Gradient Effects (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: Experiments on boundary layer transition with flat, concave and convex walls and various levels of free-stream disturbance and with zero and strong streamwise acceleration have been conducted. Measurements of both fluid mechanics and heat transfer processes were taken. Examples are profiles of mean velocity and temperature; Reynolds normal and shear stresses; turbulent streamwise and cross-stream heat fluxed; turbulent Prandtl number; and streamwise variations of wall skin friction and heat transfer coefficient values. Free-stream turbulence levels were varied over the range from about 0.3 percent to about 8 percent. The effects of curvature on the onset of transition under low disturbance conditions are clear; concave curvature leads to an earlier and more rapid transition and the opposite is true for convex curvature This was previously known but little documentation of the transport processes in the flow was available
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 373-388; NASA/CP-2007-214667
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  • 145
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    Some Scenarios for Transition on Turbomachinery Blading (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: Measurements on transition under different levels of adverse pressure gradient and free-stream turbulence level are described. This extensive series of investigations, which was predicated on intermittency measurement techniques, has resulted in correlations for transition length and turbulent spot formation rate. These correlations rae intended to be used in conjunction with boundary layer prediction methods and examples are given of such predictions. More effective predictions of the transition region, especially under conditions of variable pressure gradient, are dependent on a more comprehensive understanding of transition and spot behavior. It is expected that this will result in improved transition modeling.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 311-318; NASA/CP-2007-214667
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  • 146
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    Transition in Separating-Reattaching Boundary Layer Flows (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: Experimental work with leading edge separation bubbles is presented to clarify the issues regarding transition in separated regions. Hot-wire measurements, in the form of oscilloscope traces, turbulence intermittency and conditionally sampled velocity distributions are given. The resulting points of transition onset and spot production rates are compared to existing correlations.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 421-429; NASA/CP-2007-214667
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  • 147
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    Active Control of Transition Using the Lorentz Force (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: A new concept and technique has been developed to directly control boundary-layer transition and turbulence. Near-wall vertical motions are directly suppressed through the application of Lorentz force. Current (j) and magnetic (b) fields are applied parallel to the boundary and normal to each other to produce a Lorentz force (j x B) normal to the boundary. This approach is called magnetic turbulence control (MTC). Experiments have been performed on flat-plate transitional and turbulent boundary layers in water seeded with a weak electrolyte.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 51-59; NASA/CP-2007-214667
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  • 148
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    Characteristics of Boundary Layer Transition in a Multi-Stage Low-Pressure Turbine (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: An experimental investigation of boundary layer transition in a multi-stage turbine has been completed using surface-mounted hot-film sensors. Tests were carried out using the two-stage Low Speed Research Turbine of the Aerodynamics Research Laboratory of GE Aircraft Engines. Blading in this facility models current, state-of-the-art low pressure turbine configurations. The instrumentation technique involved arrays of densely-packed hot-film sensors on the surfaces of second stage rotor and nozzle blades. The arrays were located at mid-span on both the suction and pressure surfaces. Boundary layer measurements were acquired over a complete range of relevant Reynolds numbers. Data acquisition capabilities provided means for detailed data interrogation in both time and frequency domains. Data indicate that significant regions of laminar and transitional boundary layer flow exist on the rotor and nozzle suction surfaces. Evidence of relaminarization both near the leading edge of the suction surface and along much of the pressure surface was observed. Measurements also reveal the nature of the turbulent bursts occuring within and between the wake segments convecting through the blade row. The complex character of boundary layer transition resulting from flow unsteadiness due to nozzle/nozzle, rotor/nozzle, and nozzle/rotor wake interactions are elucidated using these data. These measurements underscore the need to provide turbomachinery designers with models of boundary layer transition to facilitate accurate prediction of aerodynamic loss and heat transfer.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 1-2; NASA/CP-2007-214667
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  • 149
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    Development of Hairpin Vortices in Turbulent Spots and End-Wall Transition (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: The end-stage phase of boundary layer transition is characterized by the development of hairpin-like vortices which evolve rapidly into patches of turbulent behavior. In general, the characteristics of the evolution form this hairpin stage to the turbulent stage is poorly understood, which has prompted the present experimental examination of hairpin vortex development and growth processes. Two topics of particular relevance to the workshop focus will be covered: 1) the growth of turbulent spots through the generatio and amalgamation of hairpin-like vortices, and 2) the development of hairpin vortices during transition in an end-wall junction flow. Brief summaries of these studies are described below. Using controlled generation of hairpin vortices by surface injection in a critical laminar boundary layer, detailed flow visualization studies have been done of the phases of growth of single hairpin vortices, from the initial hairgin generation, through the systematic generation of secondary hairpin-like flow structures, culminating in the evolution to a turbulent spot. The key to the growth process is strong vortex-surface interactions, which give rise to strong eruptive events adjacent to the surface, which results in the generation of subsequent hairpin vortex structures due to inviscid-viscuous interactions between the eruptive events and the free steam fluid. The general process of vortex-surface fluid interaction, coupled with subsequent interactions and amalgamation of the generated multiple hairpin-type vortices, is demonstrated as a physical mechanism for the growth and development of turbulent spots. When a boundary layer flow along a surface encounters a bluff body obstruction extending from the surface (such as cylinder or wing), the strong adverse pressure gradients generated by these types of flows result in the concentration of the impinging vorticity into a system of discrete vortices near the end-wall juncture of the obstruction, with the extensions of the vortices engirdling the obstruction to form "necklace" or "horseshoe" vortices. Recent hydrogen bubble and particle image visualization have shown that as Reynolds number is increased for a laminar approach flow, the flow will become critical, and a destabilization of the necklace vortices results in the development of an azimuthal waviness, or "kinks", in the vortices. These vortex kinks are accentuated by Biot-Savart effects, causing portions of a distorted necklace vortex to make a rapid approach to the surface, precipitating processes of localized, three-dimensional surface interactions. These interactions result in the rapid generation, focussing, and ejection of thin tongues of surface fluid, which rapidly roll-over and appear as hairpin vortices in the junction region. Subsequent amalgamation of these hairpin vortices with the necklace vortices produces a complex transitional-type flow. A presentation of key results from both these studies will be done, emphasizing both the ubiquity of such hairpin-type flow structures in manifold transitional-type flows, and the importance of vortex-surface interactions n the development of hairpin vortices.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 79-89; NASA/CP-2007-214667
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  • 150
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    Turbulence Modeling for the Simulation of Transition in Wall Shear Flows (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Our research involves study of the behavior of k-epsilon turbulence models for simulation of bypass-level transition over flat surfaces and turbine blades. One facet of the research has been to assess the performance of a multitude of k-epsilon models in what we call "natural transition", i.e. no modifications to the k-e models. The study has been to ascertain what features in the dynamics of the model affect the start and end of the transition. Some of the findings are in keeping with those reported by others (e.g. ERCOFTAC). A second facet of the research has been to develop and benchmark a new multi-time scale k-epsilon model (MTS) for use in simulating bypass-level transition. This model has certain features of the published MTS models by Hanjalic, Launder, and Schiestel, and by Kim and his coworkers. The major new feature of our MTS model is that it can be used to compute wall shear flows as a low-turbulence Reynolds number type of model, i.e. there is no required partition with patching a one-equation k model in the near-wall region to a two-equation k-epsilon model in the outer part of the flow. Our MTS model has been studied extensively to understand its dynamics in predicting the onset of transition and the end-stage of the transition. Results to date indicate that it far superior to the standard unmodified k-epsilon models. The effects of protracted pressure gradients on the model behavior are currently being investigated.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 495-514; NASA/CP-2007-214667
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  • 151
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    Intermittent Turbulence in the Attachment Line Flow Formed on an Infinite Swept Wing (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: The transition process which takes place in the attachment-line boundary layer in the presence of gross contamination is an issue of considerable interest to wing designers. It is well known that this flow is very sensitive to the presence of isolated roughness and that transition can be initiated at a very low value of the local medium thickness Reynolds number.Moreover, once the attachment line is turbulent, the flow over the whole wing chords, top and bottom surface, will be turbulent and this has major implications for wind drag.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 327-337; NASA/CP-2007-214667
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  • 152
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    Stability of the Boundary Layer and the Spot (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: The similarity among turbulent spots observed in various transition experiments, and the rate in which they contaminate the surrounding laminar boundary layer is only cursory. The shape of the spot depends on the Reynolds number of the surrounding boundary layer and on the pressure gradient to which it and the surrounding laminar flow are exposed. The propagation speeds of the spot boundaries depend, in addition, on the location from which the spot originated and do not simply scale with the local free stream velocity. The understanding of the manner in which the turbulent manner in which the turbulent spot destabilizes the surrounding, vortical fluid is a key to the understanding of the transition process. We therefore turned to detailed observations near the spot boundaries in general and near the spanwise tip of the spot in particular.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 285-309; NASA/CP-2007-214667
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  • 153
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    The Many Worlds of Transition Research. (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: The transition from laminar to turbulent flow in a boundary layer is a complex phenomenon that may take different routes, each involving distinct stages governed by different, often not-yet unraveled dynamical principles. There are, surprisingly, questions concerning virtually every stage in the process, beginning with receptivity to external disturbances, the linear stability of spatially developing flows, different possible nonlinear end games, the formation and propagation of turbulent spots and the emergence of fully developed turbulent flow. There seems no doubt that the flow has to be seen as a forced, nonlinear spatio-temporal system, but the system is so complex that to extract simple insights is still very difficult.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 3-10; NASA/CP-2007-214667
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  • 154
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    The Evolution of Modulated Wavetrains Into Turbulent Spots (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Experiment are being carried out to study the process by which th almost periodic disturbance waves generated naturally by the freestream evolve into turbulence. The boundary layer on a flat plate has been used for this study. The novelty of the approach is in the form of artificial excitation that is used. In this work the flow is excited artificially by deterministic white noise. The weak T-S wave created develops down stream, becomes nonlinear and blows up locally onto a highly distorted flow. These large local distortions of the mean flow allow very high frequency disturbances to grow and form into small turbulent spots. The spots arise from the excitation, and if the same noise sequence is repeated a spot will form at the same position and time instant relative to the excitation.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 39-49; NASA/CP-2007-214667
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  • 155
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    A Research Program for Improving Heat Transfer Prediction Capability for the Laminar to Turbulent Transition Region of Turbine Vanes/Blades (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: A program sponsored by the National Aeronautics and Space Administration (NASA) for the investigation of the heat transfer in the transition region of turbine vanes and blades with the object of improving the capability for predicting heat transfer is described,. The accurate prediction of gas-side heat transfer is important to the determination of turbine longevity, engine performance and developmental costs. The need for accurate predictions will become greater as the operating temperatures and stage loading levels of advanced turbine engines increase. The present methods for predicting transition shear stress and heat transfer on turbine blades are based on incomplete knowledge and are largely empirical. To meet the objectives of the NASA program, a team approach consisting of researchers from government, universities, a research institute, and a small business is presented. The research is divided into areas of experimentation, direct numerical simulation (DNS) and turbulence modeling. A summary of the results to date is given for the above research areas in a high-disturbance environment (bypass transition) with a discussion of the model development necessary for use in numerical codes.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 235-267; NASA/CP-2007-214667
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  • 156
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    From Disturbances to Instabilities, to Breakdown to Turbulence: The Physics of Transition in Boundary Layers (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: In order to understand the end-stages of boundary layer transition in low as well as high disturbance environments it is desirable to establish a unified view of the sequences of physico-mathematical phenomena that lead from laminar flow to self-sustained "bursting" in wall turbulence. The dominant driving disturbances: oncoming free turbulence, unsteady pressure fields, inhomogeneous density fields, inhomogeneities in wall geometry, all force disturbed motions within the boundary layer via multiple competitive receptivity mechanisms. For small disturbances, a sequence of instabilities then leads to sporadic local bursting very near the wall which can sustain turbulence. The local seeds of turbulence then somehow propagate to engulf quite rapidly the surrounding disturbed but still laminar regions.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 11-21; NASA/CP-2007-214667
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  • 157
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    Boundary Layer Transition on an Axial Compressor Stator Blade-Wake Passing and Freestream Turbulence Effects (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Quantitative observations of transitional boundary layers in regions of strong flow deceleration on an axial compressor stator blade are reported. Measurements are obtained at a fixed chordwise position, and the blade incidence was varied by changing the compressor throughflow so as to move the transition region relative to the stationary probe. It was thus possible to observe typical boundary layer behavior at various stages of transition in the turbomachine environment. The range of observations covers separating laminar flow at transition onset, and reattachment of intermittently turbulent periodically separated shear layers.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 163-173; NASA/CP-2007-214667
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  • 158
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    Intermittency Models and Spot Measurements (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: Experimental work at the University of Oxford Osney Lab has demonstrated characteristics of the late-stage transition process by the use of thin-film heat transfer gauges. The development of turbulent spots has been observed in a range of environments, including flat plates, turbine blade cascade tests and wake-passing experiments. These results were taken at Mach/Reynolds numbers and gas-to-wall temperature ratios representative of gas turbines. Analyses of the spot characteristics are consistent with measurements taken in low speed experiments, and support the Schubauer and Klebanoff type of turbulent spots. The addition of simulated wakes from upstream stages has been observed to be primarily superpositional for these tests.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 149-162; NASA/CP-2007-214667
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  • 159
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    Hairpin Vortices and the Final Stages of Transition (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: A spatially developing direct numerical simulation has been performed for flow over a flat plate that is subjected to a one-time fluid injection through an elongated slit in the wall. The flow parameters have been chosen to closely approximate the experimental conditions of Haidari, Taylow, and Smith (AIAA-89-0964). A hairpin vortex quickly develops near the upstream end of the slit, and a pair of necklace vortices form around the slow-moving injection fluid. As seen in the experiments and reported in Haidari and Smith (in review, JFM), the hairpin vortex spawns both in-line and sidelobe secondary vortices. However, no subdsidiary vortices (those formed by the inviscid deformation of a vortex-line bundle) are observed. At later times, a set of three different types of vortices are identified: hairpin vortex structures with heads that rise away from the wall horseshoe-shaped vortices that do not rise out of the boundary layer, and quasi-streamwise vortices. These structures interact with each other and with the wall layer to generate new vortices that are similar in structure to those mentioned above, although a particular parent vortex may have an offspring that more nearly resembles another member of the set. Perturbation velocity and vertical vorticity contours reveal an arrowhead shape of the highly disturbed region that is reminiscent of a turbulent spot. Spatially averaged velocity profiles in the highly disturbed area are nonlaminar, but as yet do not show typical low-of-the-wall behavior.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 91-114; NASA/CP-2007-214667
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  • 160
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    Turbulent-Spot Growth Characteristics: Wind-Tunnel and Flight Measurements of Natural Transition at High Reynolds and Mach Numbers (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: A series of experiments are described which examine the growth of turbulent spots on a flat plate at Reynolds and Mach numbers typical of gas-turbine blading. A short-duration piston tunnel is employed and rapid-response miniature surface-heat-transfer gauges are used to asses the state of the boundary layer. The leading- and trailing-edge velocities of spots are reported for different external pressure gradients and Mach numbers. Also, the lateral spreading angle is determined from the heat-transfer signals which demonstrate dramatically the reduction in spot growth associated with favorable pressure gradients. An associated experiment on the development of turbulent wedges is also reported where liquid-crystal heat-transfer techniques are employed in low-speed wind tunnel to visualize and measure the wedge characteristics. Finally, both liquid crystal techniques and hot-film measurements from flight tests at Mach number of 0.6 are presented.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 319-325; NASA/CP-2007-214667
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  • 161
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    Initiation of Turbulent Spots in a Laminar Boundary Layer by Rigid Falling Particulates (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: A transitional laminar boundary layer is developed on a 1m wide km long flat plate in a 0.6m deep water channel with a freestream velocity of 15-50 cm/s. A particulate dispenser under computer control ejects individual particles having diameters of 1/3 delta into the free stream. The particulates are introduced with an initial velocity of U(sub infinity) in the direction of the free stream. They have differing specific gravities of 1.03-2.7 which introduces an additional non-dimensional parameter relating the time taken to traverse the boundary layer to the convective time scale. The particulates produce a wake in the upper region of the boundary layer as they sink towards the wall. Visualization data taken over the range 5 x 10(exp 4) less than Re(sub x) less than 5 x 10(exp 5) indicate that turbulent spots are produced by the disturbances due to the wake rather than by the particulates themselves. This suggests that the spot formation process in this case may be inviscid in nature and may not be strongly influenced by the presence of the wall.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 23-30; NASA/CP-2007-214667
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  • 162
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    The Role of Separation Bubbles on the Aerodynamic Characteristics of Airfoils, Including Stall and Post-Stall, at Low Reynolds Numbers (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Airfoils at high Reynolds numbers, in general, have small separation bubbles that are usually confined to the leading edge. Since the Reynolds number is large, the turbulence model for the transition region between the laminar and turbulent flow is not important. Furthermore, the onset of transition occurs either at separation or prior to separation and can be predicted satisfactorily by empirical correlations when the incident angle is small and can be assumed to correspond to laminar separation when the correlations do not apply, i.e., at high incidence angles.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 339-356; NASA/CP-2007-214667
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  • 163
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    Transition Zone Modeling (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: This lecture reviews current practice as well as new modeling ideas for the calculation of at least skin friction and heat transfer between the onset and end of transition.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 431-471; NASA/CP-2007-214667
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  • 164
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    Simulations of Boundary-Layer Transition (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: For incompressible benchmark flows, we have demonstrated the capability of the parabolized stability equations (PSE) to simulate the transition process in excellent agreement with microscopic experiments and direct Navier-Stokes simulations at modest computational cost. Encouraged by these results, we have developed the PSE methodology of three-dimensional boundary-layers in general curvilinear coordinates for the range from low to hypersonic speeds, and for both linear and nonlinear problems. For given initial and boundary conditions, the approach permits simulations from receptivity through linear and secondary instabilities into the late stages of transition where significant changes in skin friction and heat transfer coefficients occur. We have performed transition simulations for a variety of two- and three-dimensional similarity solutions and for realistic flows over swept wings at subsonic and supersonic speeds, the pressure ans suction side of turbine blades at low and medium turbulence levels, and over a blunt cone at Mach number Ma = 8. We present selected results for different transition mechanisms with emphasis on the late stage of transition and the evolution of wall-shear stress and heat transfer.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 473-487; NASA/CP-2007-214667
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    Computations of Flow Over the Hump Model Using Higher-Order Method With Turbulence Modeling (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: The flow over the two-dimensional hump model is computed by solving the RANS equations with kappa-omega (SST) model. The governing equations, the flow equations and the turbulent equations, are solved using the 5th order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using explicit third order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The WENO and the TVD methods and the formulas are explained in [1] and the application of ENO method to N-S equations is given in [2]. The solution method implemented in this computation is described in detail in [3].
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 3.15.1 - 3.15.5; NASA/CP-2007-214874
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  • 166
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    Two-Dimensional Flow Control Analysis on the Hump Model (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: Computational analyses have been conducted on the Wall-mounted Glauert-Goldschmied type body ("hump" model) with the Full Unstructured Navier-Stokes 2-D (FUN2D) flow solver developed at NASA LaRC. This investigation uses the time-accurate Reynolds-averaged Navier- Stokes (RANS) approach to predict aerodynamic performance of the active flow control experimental database for the hump model. The workshop is designed to assess the current capabilities of different classes of turbulent flow solution methodologies, such as RANS, to predict flow fields induced by synthetic jets and separation control geometries. The hump model being studied is geometrically similar to that previously tested both experimentally and computationally at NASA LaRC [ref. 1 and 2, respectively].
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 3.10.1 - 3.10.5; NASA/CP-2007-214874
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  • 167
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    Transition Models for Engineering Calculations (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: While future theoretical and conceptual developments may promote a better understanding of the physical processes involved in the latter stages of boundary layer transition, the designers of rotodynamic machinery and other fluid dynamic devices need effective transition models now. This presentation will therefore center around the development of of some transition models which have been developed as design aids to improve the prediction codes used in the performance evaluation of gas turbine blading. All models are based on Narasimba's concentrated breakdown and spot growth.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 133-148; NASA/CP-2007-214667
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  • 168
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    Aspects of Transition in Turbomachines (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: This talk provides a description of several types of transition encountered in turbomachines. It is based largely on personal experience of the detection of transition in turbomachines. Examples are taken from axial compressors, axial turbines and radial turbines. The illustrations are concerned with transition in steady and unsteady boundary layers that develop under the influence of two-dimensional and three-dimensional flow fields.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 115-132; NASA/CP-2007-214667
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  • 169
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    Boundary Layer Development on a Turbine Blade in a Linear Cascade (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Several different boundary-layer development patterns for flow over the suction surface of a turbine airfoil in a linear cascade were studied and documented using a sliding surface hot-film sensor. The state of the boundary layer, whether laminar, transitional or turbulent, was determined at numerous locations along the airfoil suction surface from leading to trailing edge. Boundary-layer transition from laminar to turbulent flow through laminar separation and turbulent reattachment, or through a combination of bypass transition and strong and weak separation and turbulent reattachment, or through solely bypass transition without separation, was observed and benchmark data were recorded. Surface flow visualization and numerical boundary-layer analysis results are consistent with the hot-film data. Flow and geometry information necessary for nmerical code operation is available.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 207-232; NASA/CP-2007-214667
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  • 170
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    On the Evolution of Localized Disturbances and their Spanwise Interactions Leading to Breakdown (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Localized disturbances in a laminar boundary layer represent a more realistic model of transition than the extensively studies, two or quasi three-dimensional perturbations regardless of the fact if they evolve in a linear manner or not. Localized disturbances can originate by surface imperfections, insects or dust. The disturbances can be harmonic (i.e. containing a single frequency and a complete set of spanwise wave numbers) or Pulsed (i.e. containing a band of streamwise and spanwise wave numbers). At sufficiently low amplitudes localized disturbances behave according to a linear stability model. It is highly probably that in a natural transition process such localized disturbances will overslap and interact. These interactions could either delay transition because of a partial wave cancellation resulting in an attenuation of the disturbance, or adversely enhance it by promoting nonlinear interactions. The nonlinearity could be simply amplitude dependent or cause a triad resonance. Nonlinear processes in a wave packet lead to breakdown and to the formation of turbulent spots. When the amplitude of the harmonic disturbance saturates, nonlinear processes widen the band of the lower amplified frequencies adjacent to the frequency of excitation. Experimental results describing the spanwise interactions of harmonic and pulsed localized disturbances leading to breakdown will be presented and discussed. A comparison to the evolution and breakdown of a single localized disturbance will be provided.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 390-419; NASA/CP-2007-214667
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  • 171
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    Numerical Simulations of the Late Stages of Transition to Turbulence (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: This viewgraph presentation reviews direct numerical simulation in the late stages of the transition process to turbulence.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 489-493; NASA/CP-2007-214667
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  • 172
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    Fluid Mechanics and Heat Transfer in Transitional Boundary Layers (2007)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: Experiments have been performed to investigate the effects of elevated free-stream turbulence and streamwise acceleration on flow and thermal structures in transitional boundary layers. The free-stream turbulence ranges from 0.5 to 6.4% and the streamwise acceleration ranges from K = 0 to 0.8 x 10(exp -6). The onset of transition, transition length and the turbulent spot formation rate are determined. The statistical results and conditionally sampled results of th streamwise and cross-stream velocity fluctuations, temperature fluctuations, Reynolds stress and Reynolds heat fluxes are presented.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 175-205; NASA/CP-2007-214667
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  • 173
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    Nonlinear Theory and Breakdown (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: The main points of recent theoretical and computational studies on boundary-layer transition and turbulence are to be highlighted. The work is based on high Reynolds numbers and attention is drawn to nonlinear interactions, breakdowns and scales. The research focuses in particular on truly nonlinear theories, i.e. those for which the mean-flow profile is completely altered from its original state. There appear to be three such theories dealing with unsteady nonlinear pressure-displacement interactions (I), with vortex/wave interactions (II), and with Euler-scale flows (III). Specific recent findings noted for these three, and in quantitative agreement with experiments, are the following. Nonlinear finite-time break-ups occur in I, leading to sublayer eruption and vortex formation; here the theory agrees with experiments (Nishioka) regarding the first spike. II gives rise to finite-distance blowup of displacement thickness, then interaction and break-up as above; this theory agrees with experiments (Klebanoff, Nishioka) on the formation of three-dimensional streets. III leads to the prediction of turbulent boundary-layer micro-scale, displacement-and stress-sublayer-thicknesses.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 69-78; NASA/CP-2007-214667
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  • 174
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    Role of Detuning in the Final Stage of Subharmonic Mode Transition in Boundary Layers (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: This work involves mechanisms for transition to turbulence in a Blasius boundary layer through resonant interactions between a plane Tollmien-Schlichting Wave and pairs of oblique waves with equal-but-opposite wave angles. When the frequency of the TS wave is exactly twice that of the oblique waves, we have a "tuned" subharmonic resonance. This leads to the enhanced growth of the oblique modes. Following this, other nonlinear interactions lead to the growth of other 3-D modes which are harmonically based, along with a 3-D mean flow distortion. In the final stage of this process, a gradual spectral filling occurs which we have traced to the growth of fundamental and subharmonic side-band modes. To simulate this with controlled inputs, we introduced the oblique wave pairs at the same conditions, but shifted the frequency of the plane TS mode (by as much as 12 percent) so that it was not exactly twice that of the 3-D modes. These "detuned" conditions also lead to the enhanced growth of the oblique modes, as well as discrete side-band modes which come about through sum and difference interactions. Other interactions quickly lead to a broad band of discrete modes. Of particular importance is the lowest difference frequency which produces a low frequency modulation similar to what has been seen in past experiments with natural 3-D mode input. Cross-bispectral analysis of time series allows us to trace the origin and development of the different modes. Following these leads to a scenario which we believe is more relevant to conditions of "natural" transitions, where low amplitude background disturbances either lead to the gradual detuning of exact fundamental/subharmonic resonance, or in which 3-D mode resonance is detuned from the onset. The results contrast the two conditions, and document the propensity of the 2-D/3-D mode interactions to become detuned.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 61-67; NASA/CP-2007-214667
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  • 175
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    Late Stage Hypersonic Boundary Layer Transition (2007)
    In:  CASI
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    Publication Date: 2018-06-06
    Description: Our knowledge of late-stage hypersonic boundary layer transition is very limited, since most theoretical and experimental work has concentrated on the linear disturbance amplification regime. Although experiments show linear higher harmonics beginning at approximately one-half the transition Reynolds number, there is no experimental evidence for subharmonics, in contrast to subsonic boundary layer transition. A practical definition of transition is the location where mean surface heat transfer first begins to rise above laminar values. Hot wire spectra show that prior to transition, spectral dispersion occurs, with second mode energy decreasing, and energy at neighboring frequencies increasing. Near the transition point, disturbance energy begins to spread from the critical layer toward the wall. Greater emphasis on the breakdown region is planned for future experiments.
    Keywords: Aerodynamics
    Type: Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition; 357-369; NASA/CP-2007-214667
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  • 176
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    Separation Control Over A Wall-Mounted Hump (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: Separation control by means of steady suction or zero efflux oscillatory jets is known to be effective in a wide variety of flows under different flow conditions. Control is effective when applied in a nominally two-dimensional manner, for example, at the leading-edge of a wing or at the shoulder of a deflected flap. Despite intuitive understanding of the flow, at present there is no accepted theoretical model that can adequately explain or describe the observed effects of the leading parameters such as reduced suction-rate, or frequency and momentum input. This difficulty stems partly from the turbulent nature of the flows combined with superimposed coherent structures, which are usually driven by at least one instability mechanism. The ever increasing technological importance of these flows has spurned an urgent need to develop turbulence models with a predictive capability. Present attempts to develop such models are hampered in one way or another by incomplete data sets, uncertain or undocumented inflow and boundary conditions, or inadequate flow-field measurements. This paper attempts to address these issues by conducting an experimental investigation of a lowspeed separated flow over a wall-mounted hump model. The model geometry was designed by Seifert & Pack, who measured static and dynamic pressures on the model for a wide range of Reynolds and Mach numbers and control conditions. This paper describes the present experimental setup, as well as the types and range of data acquired. Sample data is presented and future work is discussed.
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 3.1.1 - 3.1.5; NASA/CP-2007-214874
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  • 177
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    Synthetic Jets in Quiescent Air (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: An oscillatory jet with zero net mass flow is generated by a cavity-pumping actuator. Among the three test cases selected for the Langley CFD validation workshop to assess the current CFD capabilities to predict unsteady flow fields, this basic oscillating jet flow field is the least complex and is selected as the first test case. Increasing in complexity, two more cases studied include jet in cross flow boundary layer and unsteady flow induced by suction and oscillatory blowing with separation control geometries. In this experiment, velocity measurements from three different techniques, hot-wire anemometry, Laser Doppler Velocimetry (LDV) and Particle Image Velocimetry (PIV), documented the synthetic jet flow field. To provide boundary conditions for computations, the experiment also monitored the actuator operating parameters including diaphragm displacement, internal cavity pressure, and internal cavity temperature.
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 1.1.1 - 1.1.5; NASA/CP-2007-214874
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  • 178
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    Experimental Evaluation of an Isolated Synthetic Jet IN Crossflow (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: The second case for this workshop builds upon the isolated synthetic jet of Case 1 by adding a crossflow, with no streamwise pressure gradient, for the developing jet to interact with. Formally, Case 2 examines the interaction of a single, isolated, synthetic jet and a fully turbulent zero-pressure gradient boundary layer. The resulting flow has many of the characteristics that need to be modeled with fidelity if the results of the calculations are to serve as the basis for research and design with active flow control devices. These include the turbulence in the boundary layer, the time-evolution of the large vortical structure emanating from the jet orifice and its subsequent interaction with and distortion by the boundary layer turbulence, and the effect of the suction cycle on the boundary layer flow. In a synthetic jet, the flow through the orifice and out into the outer flowfield alternates between an exhaust and a suction cycle, driven by the contraction and expansion of a cavity internal to the actuator. In the present experiment, the volume changes in the internal cavity are accomplished by replacing one of the rigid walls of the cavity, the wall opposite the orifice exit, with a deformable wall. This flexible wall is driven by a bottom-mounted moveable piston. The piston is driven electro-mechanically. The synthetic jet issues into the external flow through a circular orifice. In the present experiment, this orifice has a diameter of 0.250 inches (6.35 mm). The flow is conceptually similar to that documented in Schaeffler [1]. To document the flow, several measurement techniques were utilized. The upstream boundary conditions (in-flow conditions), and several key phase-averaged velocity profiles were measured with a 3-component laser-Doppler velocimetry system. Phase-averaged velocity field measurements were made with both stereo digital particle image velocimetry and 2-D digital particle image velocimetry as the primary measurement system. Surface pressure measurements were made utilizing an electronically scanned pressure system.
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 2.1.1 - 2.1.8; NASA/CP-2007-214874
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  • 179
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    Simulation of a Synthetic Jet in Quiescent Air Using TLNS3D Flow Code (2007)
    In:  CASI
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    Publication Date: 2018-06-11
    Description: Although the actuator geometry is highly three-dimensional, the outer flowfield is nominally two-dimensional because of the high aspect ratio of the rectangular slot. For the present study, this configuration is modeled as a two-dimensional problem. A multi-block structured grid available at the CFDVAL2004 website is used as a baseline grid. The periodic motion of the diaphragm is simulated by specifying a sinusoidal velocity at the diaphragm surface with a frequency of 450 Hz, corresponding to the experimental setup. The amplitude is chosen so that the maximum Mach number at the jet exit is approximately 0.1, to replicate the experimental conditions. At the solid walls zero slip, zero injection, adiabatic temperature and zero pressure gradient conditions are imposed. In the external region, symmetry conditions are imposed on the side (vertical) boundaries and far-field conditions are imposed on the top boundary. A nominal free-stream Mach number of 0.001 is imposed in the free stream to simulate incompressible flow conditions in the TLNS3D code, which solves compressible flow equations. The code was run in unsteady (URANS) mode until the periodicity was established. The time-mean quantities were obtained by running the code for at least another 15 periods and averaging the flow quantities over these periods. The phase-locked average of flow quantities were assumed to be coincident with their values during the last full time period.
    Keywords: Aerodynamics
    Type: Proceedings of the 2004 Workshop on CFD Validation of Synthetic Jets and Turbulent Separation Control; 1.4.1 - 1.4.5; NASA/CP-2007-214874
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  • 180
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    Time Resolved PIV for Space-Time Correlations in Hot Jets (2007)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Temporally Resolved Particle Image Velocimetry (TR-PIV) is the newest and most exciting tool recently developed to support our continuing efforts to characterize and improve our understanding of the decay of turbulence in jet flows -- a critical element for understanding the acoustic properties of the flow. A new TR-PIV system has been developed at the NASA Glenn Research Center which is capable of acquiring planar PIV image frame pairs at up to 25 kHz. The data reported here were collected at Mach numbers of 0.5 and 0.9 and at temperature ratios of 0.89 and 1.76. The field of view of the TR-PIV system covered 6 nozzle diameters along the lip line of the 50.8 mm diameter jet. The cold flow data at Mach 0.5 were compared with hotwire anemometry measurements in order to validate the new TR-PIV technique. The axial turbulence profiles measured across the shear layer using TR-PIV were thinner than those measured using hotwire anemometry and remained centered along the nozzle lip line. The collected TR-PIV data illustrate the differences in the single point statistical flow properties of cold and hot jet flows. The planar, time-resolved velocity records were then used to compute two-point space-time correlations of the flow at the Mach 0.9 flow condition. The TR-PIV results show that there are differences in the convective velocity and growth rate of the turbulent structures between cold and hot flows at the same Mach number
    Keywords: Aerodynamics
    Type: AIAA Paper 2006-0047 , E-17864 , 45th AIAA Aerospace Sciences Meeting and Exhibit; Jan 08, 2007 - Jan 11, 2007; Reno, NV; United States
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    An Application of Overset Grids to Payload/Fairing Three-Dimensional Internal Flow CFD Analysis (2007)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: The application of overset grids to the computational fluid dynamics analysis of three-dimensional internal flow in the payload/fairing of an expendable launch vehicle is described. In conjunction with the overset grid system, the flowfield in the payload/fairing configuration is obtained with the aid of OVERFLOW Navier-Stokes code. The solution exhibits a highly three dimensional complex flowfield with swirl, separation, and vortices. Some of the computed flow features are compared with the measured Laser-Doppler Velocimetry (LDV) data on a 1/5th scale model of the payload/fairing configuration. The counter-rotating vortex structures and the location of the saddle point predicted by the CFD analysis are in general agreement with the LDV data. Comparisons of the computed (CFD) velocity profiles on horizontal and vertical lines in the LDV measurement plane in the faring nose region show reasonable agreement with the LDV data.
    Keywords: Aerodynamics
    Type: KSC-2007-215
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    The Development of Modal Testing Technology for Wind Turbines: A Historical Perspective (2007)
    In:  Other Sources
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    Publication Date: 2019-07-19
    Description: Wind turbines are very large, flexible structures, with aerodynamic forces on the rotating blades producing periodic forces with frequencies at the harmonics of the rotation frequency. Due to design consideration, these rotational frequencies are comparable to the modal frequencies; thus avoiding resonant conditions is a critical consideration. Consequently, predicting and experimentally validating the modal frequencies of wind turbines has been important to their successful design and operation. Performing modal tests on flexible structures over 120 meters tall is a substantial challenge, which has inspired innovative developments in modal test technology. A further trial to the analyst and experimentalist is that the modal frequencies are dependent on the turbine rotation speed, so testing a parked turbine does not fully validate the analytical predictions. The history and development of this modal testing technology will be reviewed, showing historical tests and techniques, ranging from two-meter to 100-meter turbines for both parked and rotating tests. The NExT (Natural Excitation Technique) was developed in the 1990's, as a predecessor to OMA to overcome these challenges. We will trace the difficulties and successes of wind turbine modal testing over the past twenty-five years from 1982 to the present.
    Keywords: Aerodynamics
    Type: International Operational Modal Analysis Conference; Apr 30, 2007 - May 02, 2007; Copenhagen; Denmark
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    Investigations for Supersonic Transports at Transonic and Supersonic Conditions (2007)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.
    Keywords: Aerodynamics
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  • 184
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    Time-Accurate Computations of Isolated Circular Synthetic Jets in Crossflow (2007)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: Results from unsteady Reynolds-averaged Navier-Stokes computations are described for two different synthetic jet flows issuing into a turbulent boundary layer crossflow through a circular orifice. In one case the jet effect is mostly contained within the boundary layer, while in the other case the jet effect extends beyond the boundary layer edge. Both cases have momentum flux ratios less than 2. Several numerical parameters are investigated, and some lessons learned regarding the CFD methods for computing these types of flow fields are summarized. Results in both cases are compared to experiment.
    Keywords: Aerodynamics
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  • 185
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    F-15B Quiet Spike(TradeMark) Aeroservoelastic Flight-Test Data Analysis (2007)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: System identification is utilized in the aerospace community for development of simulation models for robust control law design. These models are often described as linear, time-invariant processes and assumed to be uniform throughout the flight envelope. Nevertheless, it is well known that the underlying process is inherently nonlinear. Over the past several decades the controls and biomedical communities have made great advances in developing tools for the identification of nonlin ear systems. In this report, we show the application of one such nonlinear system identification technique, structure detection, for the an alysis of Quiet Spike(TradeMark)(Gulfstream Aerospace Corporation, Savannah, Georgia) aeroservoelastic flight-test data. Structure detectio n is concerned with the selection of a subset of candidate terms that best describe the observed output. Structure computation as a tool fo r black-box modeling may be of critical importance for the development of robust, parsimonious models for the flight-test community. The ob jectives of this study are to demonstrate via analysis of Quiet Spike(TradeMark) aeroservoelastic flight-test data for several flight conditions that: linear models are inefficient for modelling aeroservoelast ic data, nonlinear identification provides a parsimonious model description whilst providing a high percent fit for cross-validated data an d the model structure and parameters vary as the flight condition is altered.
    Keywords: Aerodynamics
    Type: NASA/TM-2007-214618 , H-2713
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    Effects of Nose Bluntness on Stability of Hypersonic Boundary Layers over Blunt Cone (2007)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Receptivity and stability of hypersonic boundary layers are numerically investigated for boundary layer flows over a 5-degree straight cone at a free-stream Mach number of 6.0. To compute the shock and the interaction of shock with the instability waves, we solve the Navier-Stokes equations in axisymmetric coordinates. The governing equations are solved using the 5th-order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. After the mean flow field is computed, disturbances are introduced at the upstream end of the computational domain. Generation of instability waves from leading edge region and receptivity of boundary layer to slow acoustic waves are investigated. Computations are performed for a cone with nose radii of 0.001, 0.05 and 0.10 inches that give Reynolds numbers based on the nose radii ranging from 650 to 130,000. The linear stability results showed that the bluntness has a strong stabilizing effect on the stability of axisymmetric boundary layers. The transition Reynolds number for a cone with the nose Reynolds number of 65,000 is increased by a factor of 1.82 compared to that for a sharp cone. The receptivity coefficient for a sharp cone is about 4.23 and it is very small, approx.10(exp -3), for large bluntness.
    Keywords: Aerodynamics
    Type: 37th AIAA Fluid Dynamics Conference and Exhibit; Jun 25, 2007 - Jun 28, 2007; Miami, FL; United States
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    Orion Entry, Descent, and Landing Simulation (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Orion Entry, Descent, and Landing simulation was created over the past two years to serve as the primary Crew Exploration Vehicle guidance, navigation, and control (GN&C) design and analysis tool at the National Aeronautics and Space Administration (NASA). The Advanced NASA Technology Architecture for Exploration Studies (ANTARES) simulation is a six degree-of-freedom tool with a unique design architecture which has a high level of flexibility. This paper describes the decision history and motivations that guided the creation of this simulation tool. The capabilities of the models within ANTARES are presented in detail. Special attention is given to features of the highly flexible GN&C architecture and the details of the implemented GN&C algorithms. ANTARES provides a foundation simulation for the Orion Project that has already been successfully used for requirements analysis, system definition analysis, and preliminary GN&C design analysis. ANTARES will find useful application in engineering analysis, mission operations, crew training, avionics-in-the-loop testing, etc. This paper focuses on the entry simulation aspect of ANTARES, which is part of a bigger simulation package supporting the entire mission profile of the Orion vehicle. The unique aspects of entry GN&C design are covered, including how the simulation is being used for Monte Carlo dispersion analysis and for support of linear stability analysis. Sample simulation output from ANTARES is presented in an appendix.
    Keywords: Aerodynamics
    Type: AIAA GN&C Conference; Aug 20, 2007 - Aug 23, 2007; Hilton Head, SC; United States
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  • 188
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    A Review of Hypersonics Aerodynamics, Aerothermodynamics and Plasmadynamics Activities within NASA's Fundamental Aeronautics Program (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The research program of the aerodynamics, aerothermodynamics and plasmadynamics discipline of NASA's Hypersonic Project is reviewed. Details are provided for each of its three components: 1) development of physics-based models of non-equilibrium chemistry, surface catalytic effects, turbulence, transition and radiation; 2) development of advanced simulation tools to enable increased spatial and time accuracy, increased geometrical complexity, grid adaptation, increased physical-processes complexity, uncertainty quantification and error control; and 3) establishment of experimental databases from ground and flight experiments to develop better understanding of high-speed flows and to provide data to validate and guide the development of simulation tools.
    Keywords: Aerodynamics
    Type: AIAA Paper 2007-4264 , 39th AIAA Thermophysics Conference; Jun 25, 2007 - Jun 28, 2007; Miami, FL; United States
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  • 189
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    Tandem Cylinder Noise Predictions (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In an effort to better understand landing-gear noise sources, we have been examining a simplified configuration that still maintains some of the salient features of landing-gear flow fields. In particular, tandem cylinders have been studied because they model a variety of component level interactions. The present effort is directed at the case of two identical cylinders spatially separated in the streamwise direction by 3.7 diameters. Experimental measurements from the Basic Aerodynamic Research Tunnel (BART) and Quiet Flow Facility (QFF) at NASA Langley Research Center (LaRC) have provided steady surface pressures, detailed off-surface measurements of the flow field using Particle Image Velocimetry (PIV), hot-wire measurements in the wake of the rear cylinder, unsteady surface pressure data, and the radiated noise. The experiments were conducted at a Reynolds number of 166 105 based on the cylinder diameter. A trip was used on the upstream cylinder to insure a fully turbulent shedding process and simulate the effects of a high Reynolds number flow. The parallel computational effort uses the three-dimensional Navier-Stokes solver CFL3D with a hybrid, zonal turbulence model that turns off the turbulence production term everywhere except in a narrow ring surrounding solid surfaces. The current calculations further explore the influence of the grid resolution and spanwise extent on the flow and associated radiated noise. Extensive comparisons with the experimental data are used to assess the ability of the computations to simulate the details of the flow. The results show that the pressure fluctuations on the upstream cylinder, caused by vortex shedding, are smaller than those generated on the downstream cylinder by wake interaction. Consequently, the downstream cylinder dominates the noise radiation, producing an overall directivity pattern that is similar to that of an isolated cylinder. Only calculations based on the full length of the model span were able to capture the complete decay in the spanwise correlation, thereby producing reasonable noise radiation levels.
    Keywords: Aerodynamics
    Type: AIAA Paper-2007-3450 , AIAA/CEAS Aeroacoustics Conference; May 23, 2007 - May 25, 2007; Rome; Italy
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  • 190
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    Pressure Gradient Effects on Hypersonic Cavity Flow Heating (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effect of a pressure gradient on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated in support of the Space Shuttle Return-To-Flight Program. Two blunted-nose test surface geometries were developed, including an expansion plate test surface with nearly constant negative pressure gradient and a flat plate surface with nearly zero pressure gradient. The test surface designs and flow characterizations were performed using two-dimensional laminar computational methods, while the experimental boundary layer state conditions were inferred using the measured heating distributions. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process. Both open-flow and closed-flow cavities were tested on each test surface. The cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary conclusions based on an analysis of only the cavity centerline data indicate that the presence of the pressure gradient did not alter the open cavity heating for laminar-entry/laminar-exit flows, but did raise the average floor heating for closed cavities. The results of these risk-reduction studies will be used to formulate a heating assessment of potential damage scenarios occurring during future Space Shuttle flights.
    Keywords: Aerodynamics
    Type: AIAA 2006-0185 , 44th AIAA Aerospace Sciences Meeting and Exhibit; Jan 09, 2006 - Jan 12, 2006; Reno, NV; United States
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  • 191
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    Validation of the Lockheed Martin Morphing Concept with Wind Tunnel Testing (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Morphing Aircraft Structures (MAS) program is a Defense Advanced Research Projects Agency (DARPA) led effort to develop morphing flight vehicles capable of radical shape change in flight. Two performance parameters of interest are loiter time and dash speed as these define the persistence and responsiveness of an aircraft. The geometrical characteristics that optimize loiter time and dash speed require different geometrical planforms. Therefore, radical shape change, usually involving wing area and sweep, allows vehicle optimization across many flight regimes. The second phase of the MAS program consisted of wind tunnel tests conducted at the NASA Langley Transonic Dynamics Tunnel to demonstrate two morphing concepts and their enabling technologies with large-scale semi-span models. This paper will focus upon one of those wind tunnel tests that utilized a model developed by Lockheed Martin Aeronautics Company (LM). Wind tunnel success criteria were developed by NASA to support the DARPA program objectives. The primary focus of this paper will be the demonstration of the DARPA objectives by systematic evaluation of the wind tunnel model performance relative to the defined success criteria. This paper will also provide a description of the LM model and instrumentation, and document pertinent lessons learned. Finally, as part of the success criteria, aeroelastic characteristics of the LM derived MAS vehicle are also addressed. Evaluation of aeroelastic characteristics is the most detailed criterion investigated in this paper. While no aeroelastic instabilities were encountered as a direct result of the morphing design or components, several interesting and unexpected aeroelastic phenomenon arose during testing.
    Keywords: Aerodynamics
    Type: AIAA 2007-2235 , 48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Apr 23, 2007 - Apr 26, 2007; Waikiki, HI; United States
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  • 192
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    Recent Progress in Aircraft Noise Research (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An overview of the acoustics research at NASA under the Subsonic Fixed Wing project is given. The presentation describes the rationale behind the noise reduction goals of the project in the context of the next generation air transportation system, and the emphasis placed on achieving these goals through a combination of the in-house and collaborative efforts with industry, universities and other government agencies. The presentation also describes the in-house research plan which is focused on the development of advanced noise and flow diagnostic techniques, next generation noise prediction tools, and novel noise reduction techniques that are applicable across a wide range of aircraft.
    Keywords: Aerodynamics
    Type: Fundamental Aeronautics Annual Meeting; Oct 30, 2007 - Nov 01, 2007; New Orleans, LA; United States
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  • 193
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    The Potential Benefits of Advanced Casing Treatment for Noise Attenuation in Utra-High Bypass Ratio Turbofan Engines (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In order to increase stall margin in a high-bypass ratio turbofan engine, an advanced casing treatment was developed that extracted a small amount of flow from the casing behind the fan and injected it back in front of the fan. Several different configurations of this casing treatment were designed by varying the distance of the extraction and injection points, as well as varying the amount of flow. These casing treatments were tested on a 55.9 cm (22 in.) scale model of the Pratt & Whitney Advanced Ducted Propulsor in the NASA Glenn 9 by 15 Low Speed Wind Tunnel. While all of the casing treatment configurations showed the expected increase in stall margin, a few of the designs showed a potential noise benefit for certain engine speeds. This paper will show the casing treatments and the results of the testing as well as propose further research in this area. With better prediction and design techniques, future casing treatment configurations could be developed that may result in an optimized casing treatment that could conceivably reduce the noise further.
    Keywords: Aerodynamics
    Type: NASA/TM-2007-214812 , E-15967 , 35th International Congress and Exposition on Noise Control Engineering (INTER-NOISE 2006); Dec 03, 2006 - Dec 06, 2006; Honolulu, HI; United States
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  • 194
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    Minnowbrook I: 1993 Workshop on End-Stage Boundary Layer Transition (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This volume contains materials presented at the Minnowbrook I-1993 Workshop on End-Stage Boundary Layer Transition, held at the Syracuse University Minnowbrook Conference Center, New York, from August 15 to 18, 1993. This volume was previously published as a Syracuse University report edited by John E. LaGraff. The workshop organizers were John E. LaGraff (Syracuse University), Terry V. Jones (Oxford University), and J. Paul Gostelow (University of Technology, Sydney). The workshop focused on physical understanding of the late stages of transition from laminar to turbulent flows, with the specific goal of contributing to improving engineering design of turbomachinery and wing airfoils. The workshop participants included academic researchers from the United States and abroad, and representatives from the gas-turbine industry and U.S. government laboratories. To improve interaction and discussions among the participants, no formal papers were required. The physical mechanisms discussed were related to natural and bypass transition, wake-induced transition, effects of freestream turbulence, turbulent spots, hairpin vortices, nonlinear instabilities and breakdown, instability wave interactions, intermittency, turbulence, numerical simulation and modeling of transition, heat transfer in boundary-layer transition, transition in separated flows, laminarization, transition in turbomachinery compressors and turbines, hypersonic boundary-layer transition, and other related topics. This volume contains abstracts and copies of the viewgraphs presented, organized according to the workshop sessions. The workshop summary and the plenary discussion transcript clearly outline future research needs.
    Keywords: Aerodynamics
    Type: NASA/CP-2007-214667 , E-15781 , Minnowbrook I: 1993 Workshop on End-Stage Boundary Tayer Transition; Aug 15, 1993 - Aug 18, 1993; Mountain Lake, NY; United States
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  • 195
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    Sonic Boom Modeling Technical Challenge (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This viewgraph presentation reviews the technical challenges in modeling sonic booms. The goal of this program is to develop knowledge, capabilities and technologies to enable overland supersonic flight. The specific objectives of the modeling are: (1) Develop and validate sonic boom propagation model through realistic atmospheres, including effects of turbulence (2) Develop methods enabling prediction of response of and acoustic transmission into structures impacted by sonic booms (3) Develop and validate psychoacoustic model of human response to sonic booms under both indoor and outdoor listening conditions, using simulators.
    Keywords: Aerodynamics
    Type: Fundamental Aeronautics 2007 Annual Meeting; Oct 30, 2007 - Nov 01, 2007; New Orleans, LA; United States
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  • 196
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    Static Extended Trailing Edge for Lift Enhancement: Experimental and Computational Studies (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A static extended trailing edge attached to a NACA0012 airfoil section is studied for achieving lift enhancement at a small drag penalty. It is indicated that the thin extended trailing edge can enhance the lift while the zero-lift drag is not significantly increased. Experiments and calculations are conducted to compare the aerodynamic characteristics of the extended trailing edge with those of Gurney flap and conventional flap. The extended trailing edge, as a simple mechanical device added on a wing without altering the basic configuration, has a good potential to improve the cruise flight efficiency.
    Keywords: Aerodynamics
    Type: 3rd International Symposium on Integrating CFD and Experiments in Aerodynamics; Jun 20, 2007 - Jun 21, 2007; Colorado SPrings, CO; United States
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  • 197
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    Active Control of Flow Separation on a High-Lift System with Slotted Flap at High Reynolds Number (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The NASA Energy Efficient Transport (EET) airfoil was tested at NASA Langley's Low- Turbulence Pressure Tunnel (LTPT) to assess the effectiveness of distributed Active Flow Control (AFC) concepts on a high-lift system at flight scale Reynolds numbers for a medium-sized transport. The test results indicate presence of strong Reynolds number effects on the high-lift system with the AFC operational, implying the importance of flight-scale testing for implementation of such systems during design of future flight vehicles with AFC. This paper describes the wind tunnel test results obtained at the LTPT for the EET high-lift system for various AFC concepts examined on this airfoil.
    Keywords: Aerodynamics
    Type: AIAA Paper 2007-4424 , 25th AIAA Applied Aerodynamics Conference; Jun 25, 2007 - Jun 28, 2007; Miami, FL; United States
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  • 198
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    Development of Aeroservoelastic Analytical Models and Gust Load Alleviation Control Laws of a SensorCraft Wind-Tunnel Model Using Measured Data (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Aeroservoelastic (ASE) analytical models of a SensorCraft wind-tunnel model are generated using measured data. The data was acquired during the ASE wind-tunnel test of the HiLDA (High Lift-to-Drag Active) Wing model, tested in the NASA Langley Transonic Dynamics Tunnel (TDT) in late 2004. Two time-domain system identification techniques are applied to the development of the ASE analytical models: impulse response (IR) method and the Generalized Predictive Control (GPC) method. Using measured control surface inputs (frequency sweeps) and associated sensor responses, the IR method is used to extract corresponding input/output impulse response pairs. These impulse responses are then transformed into state-space models for use in ASE analyses. Similarly, the GPC method transforms measured random control surface inputs and associated sensor responses into an AutoRegressive with eXogenous input (ARX) model. The ARX model is then used to develop the gust load alleviation (GLA) control law. For the IR method, comparison of measured with simulated responses are presented to investigate the accuracy of the ASE analytical models developed. For the GPC method, comparison of simulated open-loop and closed-loop (GLA) time histories are presented.
    Keywords: Aerodynamics
    Type: IFASD 2007: International Forum on Aeroelasticity and Structural Dynamics; Jun 18, 2007 - Jun 20, 2007; Stockholm; Sweden
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  • 199
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    Influence of Initial Vorticity Distribution on Axisymmetric Vortex Breakdown and Reconnection (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An analytical treatment has been developed to study some of the axisymmetric vortex breakdown and reconnection fluid dynamic processes underlying body-vortex interactions that are frequently manifested in rotorcraft and propeller-driven fixed-wing aircraft wakes. In particular, the presence of negative vorticity in the inner core of a vortex filament (one example of which is examined in this paper) subsequent to "cutting" by a solid body has a profound influence on the vortex reconnection, leading to analog flow behavior similar to vortex breakdown phenomena described in the literature. Initial vorticity distributions (three specific examples which are examined) without an inner core of negative vorticity do not exhibit vortex breakdown and instead manifest diffusion-like properties while undergoing vortex reconnection. Though this work focuses on laminar vortical flow, this work is anticipated to provide valuable insight into rotary-wing aerodynamics as well as other types of vortical flow phenomena.
    Keywords: Aerodynamics
    Type: AIAA Aerospace Sciences Meeting; Jan 08, 2007 - Jan 11, 2007; Reno, NV; United States
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  • 200
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    NASA/GE Highly-Loaded Turbine Research Program (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An overview of the NASA/GE Highly-Loaded Turbine Research Program at the NASA Glenn Research Center is presented. The program is sponsored by the Subsonic Fixed Wing Project of the Fundamental Aeronautics Program. The goals of the turbine research program are presented along with their relationship to the higher-level program goals. Two turbine research programs are described; the highly-loaded, single-stage high pressure turbine (HPT) and the highly loaded low pressure turbine (LPT). The HPT program is centered on an extremely high pressure ratio single-stage turbine with an engine stage pressure ratio of 5.5. It was designed with a 33% increase in stage loading. It has shown performance levels 2 points better than current engines operating at the higher work level. Some advantages of the turbine include reduced weight and parts count. Optimization of the blade shape to reduce shock losses is described. The LPT program utilizes a four-stage low pressure turbine with an integral first stage vane/transition duct strut; counter-rotation; low-solidity blading; fully optimized flowpath including vanes, blades, and endwalls; and a fluidically controlled turbine vane frame/exit guide vane. The implementation of the LPT into GE s and NASA s test facilities is described. A description of NASA's Single Spool Turbine Facility that is currently under renovation is given. Facility limits on pressures, temperatures, flow rates, rotational speeds, and power absorption are described. The current renovation status is given.
    Keywords: Aerodynamics
    Type: NASA Fundamental Aeronautics 2007 Annual Meeting; Oct 30, 2007 - Nov 01, 2007; New Orleans, LA; United States
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