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  • 1
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    Universidade Estadual de Maringá. Departamento de Biologia. Programa de Pós-Graduação em Ecologia de Ambientes Aquáticos Continentais.
    Publication Date: 2022-05-13
    Description: The environmental variables are important in the functioning and dynamics of freshwater ecosystems, since changes in physical and chemical conditions of these environments contribute to alterations in the ecological functioning. This research was conducted in four neotropical flood plains: Amazônica, Araguaia, Pantanal e Paraná. The samples of zooplankton were performed in the subsurface region of limnetic lakes each floodplain, in the years 2011 and 2012 (dry and wet). At the same time the environmental variables were measured and obtained water samples for later laboratory analysis. Our goal was to investigate the effect of environmental variables in the abundance of copepods, and the following hypothesis was tested: HI) the abundance of copepods has a great set with environmental variables, and extremes range of environmental variation, the lower the abundance of these organisms, unimodal effect. Local factors (physical and chemical environmental conditions) determined the abundance of copepods, as well as the environmental gradient showed a close relationship in the spatial and temporal distribution of the species, being possible to observe higher abundance at specific levels of environmental variables (toward the great). Studies that focus on tolerance limits of the species and their relationships with environmental conditions has risen in recent years, as the anthropogenic changes in climate and natural environments has caused significant changes in community structure, facilitating the invasion by non-native species and loss species diversity and the consequent change in ecosystem functioning. However, the hypothesis (HI) was partially supported, as the abundance follows a pattern of rise towards the great, but it was not possible to clearly unimodal effect on abundance for many species.
    Description: As variáveis ambientais têm papel relevante no funcionamento e dinâmica de ambientes aquáticos continentais, uma vez que, mudanças nas condições físicas e químicas desses ambientes contribuem para alterações no padrão de funcionamento ecológico. Esta investigação foi realizada em quatro planícies de inundação neotropicais: Amazônica, Araguaia, Pantanal e Paraná. As coletas de organismos zooplanctônicos foram realizadas na subsuperfície da região limnética das lagoas de cada planície de inundação, nos anos de 2011 e 2012 (seca e cheia). Ao mesmo tempo foram medidas as variáveis ambientais e obtidas amostras de água para posterior análise em laboratório. O nosso objetivo foi investigar o efeito das variáveis ambientais na abundância dos copépodes, e foi testada a seguinte hipótese: HI) a abundância de copépodes possui um ótimo estabelecido com as variáveis ambientais, e em extremos de amplitude de variação ambiental, menor será a abundância desses organismos, efeito uninodal. Os fatores locais (condições físicas e químicas do ambiente) determinaram a abundância de copépodes, bem como a variação ambiental apresentou uma estreita relação na distribuição espacial e temporal das espécies, sendo possível observar maiores valores de abundância em níveis específicos das variáveis ambientais (direção ao ótimo). Estudos que focam nos limites de tolerância das espécies e suas relações com as condições ambientais tem ressurgido nos últimos anos, pois as alterações antropogênicas no clima e nos ambientes naturais tem causado expressivas mudanças na estrutura da comunidade, facilitação à invasão por espécies não nativas e perda da diversidade de espécies e a consequente alteração no funcionamento ecossistêmico. Contudo, a hipótese (HI) foi parcialmente corroborada, pois a abundância segue um padrão de ascensão em direção ao ótimo, porém não foi possível estabelecer claramente para muitas espécies um efeito unimodal na distribuição da abundância.
    Description: Masters
    Keywords: Copepoda ; Ciências Ambientais ; Ecology ; Community ; Abundance ; Spatial distribution ; Niche ; Brazil ; Temporal distribution ; Abundância ; Limnologia ; Distribuição espacial ; Nicho ; Copépodes (Crustacea: Copepoda) ; Planícies de inundação ; Comunidades, Ecologia de ; Brasil ; Distribuição temporal ; Limnology
    Repository Name: AquaDocs
    Type: Thesis/Dissertation
    Format: 37pp.
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  • 2
    Publication Date: 2022-05-26
    Description: © The Author(s), 2017. This article is distributed under the terms of the Creative Commons Attribution License. The definitive version was published in Geophysical Research Letters 44 (2017): 209–218, doi:10.1002/2016GL071378.
    Description: Global lake volume estimates are scarce, highly variable, and poorly documented. We developed a rigorous method for estimating global lake depth and volume based on the Hurst coefficient of Earth's surface, which provides a mechanistic connection between lake area and volume. Volume-area scaling based on the Hurst coefficient is accurate and consistent when applied to lake data sets spanning diverse regions. We applied these relationships to a global lake area census to estimate global lake volume and depth. The volume of Earth's lakes is 199,000 km3 (95% confidence interval 196,000–202,000 km3). This volume is in the range of historical estimates (166,000–280,000 km3), but the overall mean depth of 41.8 m (95% CI 41.2–42.4 m) is significantly lower than previous estimates (62–151 m). These results highlight and constrain the relative scarcity of lake waters in the hydrosphere and have implications for the role of lakes in global biogeochemical cycles.
    Description: Knut and Alice Wallenberg Foundation; National Science Foundation Graduate Research Fellowship Program Grant Number: 2388357; National Science Foundation Grant Number: OCE-1315201
    Keywords: Limnology ; Topograhy ; Scaling ; Volume ; Mean depth
    Repository Name: Woods Hole Open Access Server
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  • 3
    Publication Date: 2019-06-15
    Description: Boundary-layer transition in hypersonic flows over a straight cone can be predicted using measured freestream spectra, receptivity, and threshold values for the wall pressure fluctuations at the transition onset points. Simulations are performed for hypersonic boundary-layer flows over a 7-degree half-angle straight cone with varying bluntness at a freestream Mach number of 10. The steady and the unsteady flow fields are obtained by solving the two-dimensional Navier-Stokes equations in axisymmetric coordinates using a 5th-order accurate weighted essentially nonoscillatory (WENO) scheme for space discretization and using a third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The calculated N-factors at the transition onset location increase gradually with increasing unit Reynolds numbers for flow over a sharp cone and remain almost the same for flow over a blunt cone. The receptivity coefficient increases slightly with increasing unit Reynolds numbers. They are on the order of 4 for a sharp cone and are on the order of 1 for a blunt cone. The location of transition onset predicted from the simulation including the freestream spectrum, receptivity, and the linear and the weakly nonlinear evolutions yields a solution close to the measured onset location for the sharp cone. The simulations overpredict transition onset by about twenty percent for the blunt cone.
    Keywords: Aerodynamics
    Type: NF1676L-26446 , AIAA Journal (ISSN 0001-1452) (e-ISSN 1533-385X); 56; 1; 193-208
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  • 4
    Publication Date: 2019-07-26
    Description: Optimal initial conditions for transient growth in a two-dimensional boundary layer flow correspond to stationary, counter-rotating vortices that subsequently develop into streamwise elongated streaks, which are characterized by an alternating pattern of low and high streamwise velocity. For incompressible flows, previous studies have shown that boundary layer modulation due to streaks below a threshold amplitude level can stabilize the Tollmien-Schlichting instability waves, resulting in a delay in the onset of laminar-turbulent transition. In the supersonic regime, the linearly, most-amplified waves become three-dimensional, corresponding to oblique, first-mode waves. This change in the character of dominant instabilities leads to an important change in the transition process, which is now dominated by oblique breakdown via nonlinear interactions between pairs of first-mode waves that propagate at equal but opposite angles with respect to the free stream. Because the oblique breakdown process is characterized by a rapid amplification of stationary streamwise streaks, artificial excitation of such streaks may be expected to promote transition in a supersonic boundary layer. Indeed, suppression of those streaks has been shown to delay the onset of transition in prior literature. Consistent with those findings, the present study shows that optimally growing stationary streaks indeed destabilize the first-mode waves, but only when the spanwise wavelength of the instability waves is equal to or smaller than twice the streak spacing. Transition in a benign disturbance environment typically involves first-mode waves with significantly longer spanwise wavelengths, and hence, these waves are stabilized by the optimal growth streaks. Thus, as long as the amplification factors for the destabilized, short wavelength instability waves remain below the threshold level for transition, a significant net stabilization is achieved, yielding a transition delay that is comparable to the length of the laminar region in the uncontrolled case.
    Keywords: Aerodynamics
    Type: NF1676L-26301 , Journal of Fluid Mechanics (ISSN 0022-1120) (e-ISSN 1469-7645); 831; 524-553
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  • 5
    Publication Date: 2019-06-11
    Description: Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0, 3.5, and 7. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack. Velocity measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty as they relate to the present experiments. Measurement precisions as low as 1 m/s were observed, while the velocity dynamic range was found to be nearly a factor of 500. The spatial resolution of between 1 mm and 5 mm was found to be primarily limited by the FLEET spot size and advection of the flow. Overall measurement uncertainties ranged from 3 to 4 percent.
    Keywords: Aerodynamics
    Type: NF1676L-26518 , AIAA Journal (ISSN 0001-1452) (e-ISSN 1533-385X); 55; 12; 4142-4154
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  • 6
    Publication Date: 2019-07-12
    Description: This report documents a ballistic-range test campaign conducted in 2012 in order to estimate the aerodynamic stability characteristics of two configurations of the Supersonic Flight Dynamics Test (SFDT) vehicle prior to its initial flight in 2014. The SFDT vehicle was a test bed for demonstrating several new aerodynamic decelerator technologies then being developed under the Low-Density Supersonic Decelerator (LDSD) Project. Of particular interest here is the Supersonic Inflatable Aerodynamic Decelerator (SIAD), an inflatable attached torus used to increase the drag surface area of an entry vehicle during the supersonic portion of the entry trajectory. Two model configurations were tested in the ballistic range: one representing the SFDT vehicle prior to deployment of the SIAD, and the other representing the nominal shape with the SIAD inflated. Both models were fabricated from solid metal, and therefore, the effects of the flexibility of the inflatable decelerator were not considered. The test conditions were chosen to match, as close as possible, the Mach number, Reynolds number, and motion dynamics expected for the SFDT vehicle in flight, both with the SIAD stowed and deployed. For SFDT models with the SIAD stowed, 12 shots were performed covering a Mach number range of 3.2 to 3.7. For models representing the deployed SIAD, 37 shots were performed over a Mach number range of 2.0 to 3.8. Pitch oscillation amplitudes covered a range from 0.7 to 20.6 degrees RMS. Portions of this report (data analysis approach, aerodynamic modeling, and resulting aerodynamic coefficients) were originally published as an internal LDSD Project report [1] in 2012. In addition, this report provides a description of the test design approach, the test facility, and experimental procedures. Estimated non-linear aerodynamic coefficients, including pitch damping, for both model configurations are reported, and the shot-by-shot trajectory measurements, plotted in comparison with calculated trajectories based on the derived non-linear aerodynamic coefficients, are provided as appendices. Since the completion of these tests, two full-scale SFDT flights have been successfully conducted: one in June 2014 [2, 3], and one in June 2015 [3].
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219693 , ARC-E-DAA-TN47243
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  • 7
    Publication Date: 2019-07-12
    Description: Models are presented for the aerodynamic coefficients of Supersonic Ringsail and Disk-Gap-Band parachutes as functions of total porosity, Lambda(sub t), Mach number, M, and total angle of attack, Alpha(sub t) (when necessary). The source aerodynamic coefficients data used for creating these models were obtained from a wind tunnel test of subscale parachutes. In this wind tunnel test, subscale parachutes of both parachute types were fabricated from two different fabrics with very different permeabilities. By varying the fabric permeability, while maintaining the parachute geometry constant, it was possible to vary Alpha(sub t). The fabric permeability test data necessary for the calculation of Alpha(sub t) were obtained from samples of the same fabrics used to fabricate the subscale parachutes. Although the models for the aerodynamic coefficients are simple polynomial functions of Alpha(sub t) and M, they are capable of producing good reproductions of the source data. The (Alpha(sub t), M) domains over which these models are applicable are clearly defined. The models are applicable to flight operations on Mars.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219619 , L-20812 , NF1676L-27003
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  • 8
    Publication Date: 2019-07-12
    Description: In July 2017, a three-day Turbulence Modeling Symposium sponsored by the University of Michigan and NASA was held in Ann Arbor, Michigan. This meeting brought together nearly 90 experts from academia, government and industry, with good international participation, to discuss the state of the art in turbulence modeling, emerging ideas, and to wrestle with questions surrounding its future. Emphasis was placed on turbulence modeling in a predictive context in complex problems, rather than on turbulence theory or descriptive modeling. This report summarizes many of the questions, discussions, and conclusions from the symposium, and suggests immediate next steps.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219682 , L-20880 , NF1676L-28239
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  • 9
    Publication Date: 2019-07-12
    Description: A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.
    Keywords: Aerodynamics
    Type: NASA-TM-2017-219596 , L-20784 , NF1676L-26349
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  • 10
    Publication Date: 2019-07-12
    Description: The purpose of this manual is to aid in the design of an aerodynamics test of an earth or planetary entry capsule in a ballistic range. In this manual, much use is made of the results and experience gained in 50 years of ballistic range aerodynamics testing at the NASA Ames Research Center, and in particular, that gained in the last 27 years, while the author was working at NASA Ames. The topics treated herein include: Data to be obtained; flight data needed to design test; Reynolds number and dynamic similarity of flight trajectory and ballistic range test; capabilities of various ballistic ranges; Calculations of swerves due to average and oscillating lift and of drag-induced velocity decreases; Model and sabot design; materials, weights and stresses; Sabot separation; Launches at angle of attack and slapping with paper to produce pitch/yaw oscillations.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219473 , ARC-E-DAA-TN20974
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  • 11
    Publication Date: 2019-07-12
    Description: LM has leveraged our partnership with the Air Force Research Laboratory (AFRL) and NASA on the advanced hybrid wing body (HWB) concept to develop a commercial freighter which addresses the NASA Advanced Air Transport Technology (AATT) Project goals for improved efficiency beyond 2025. The current Air Force Research Laboratory (AFRL) Revolutionary Configurations for Energy Efficiency (RCEE) program established the HWB configuration and technologies needed for military transports to achieve aerodynamic and fuel efficiencies well beyond the commercial industry's most modern designs. This study builds upon that effort to develop a baseline commercial cargo aircraft and two HWB derivative commercial cargo aircraft to quanitify the benefit of the HWB and establish a technology roadmap for further development.
    Keywords: Aerodynamics
    Type: NASA/CR-2017-219653 , NF1676L-26587
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  • 12
    Publication Date: 2019-07-12
    Description: No abstract available
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219696/SUPPL , E-19427 , GRC-E-DAA-TN46228
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  • 13
    Publication Date: 2019-07-12
    Description: An experimental investigation of tip vortices from a NACA0012 airfoil is conducted in a low-speed wind tunnel at a chord Reynolds number (Rc) of 410(exp 4 ). Data for the stationary airfoil at various angles of attack (alpha) are first discussed. Detailed flow-field surveys are done for two cases: alpha = 10deg with attached flow and alpha = 25deg with massive flow separation. Data include mean velocity, streamwise vorticity, and turbulent stresses at various streamwise locations. For all cases, the vortex core is seen to involve a mean velocity deficit. The deficits in these cases trace to the airfoil wake, part of which gets wrapped up by the tip vortex. Comparison with data from the literature suggests that with increasing Rc, the deficit turns into an excess, with the transition occurring in the approximate Rc range of 210(exp 5) to 510(exp 5). Survey results for various shapes of the airfoil wingtip are then presented. The shapes include square and rounded ends and a number of winglet designs. Finally, data under sinusoidal pitching condition, for the airfoil with square ends, are documented. All pitching cases pertain to a mean alpha = 15deg, while the amplitude and frequency are varied. Amplitudes of +/-5deg, +/-10deg, and +/-15deg and reduced frequencies k = 0.08, 0.2, and 0.33 are covered. Digital records of all data and some of the hardware design are made available on a supplemental CD with the electronic version of the paper for those interested in numerical simulation.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219696 , E-19427 , GRC-E-DAA-TN46228
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  • 14
    Publication Date: 2019-07-24
    Description: The advancement of flow measurement techniques continues to extend experimental boundaries and thus significantly contributes to improving our understanding of both basic and applied aerodynamics. This is particularly apparent in the case of particle image velocimetry (PIV), where its application has furthered the existing knowledge in several areas of helicopter rotor aerodynamics. The complex nature of helicopter rotor flows presents unique challenges to experimentalists, including transonic flow, concentrated vortices and dynamic stall. To illustrate the impact of the technological advancements on the way helicopter aerodynamics is studied today, the development of PIV since the early nineties of the last century is reviewed and some recent PIV applications are described. Using examples of main rotor wakes, dynamic stall and flow control investigations, the capabilities of largescale, timeresolved and volumetric PIV are summarized.
    Keywords: Aerodynamics
    Type: NF1676L-24871 , AIAA Journal (ISSN 0001-1452) (e-ISSN 1533-385X); 55; 9; 2859-2874
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  • 15
    Publication Date: 2019-07-20
    Description: A synthesis is presented of recent numerical predictions for the F-16XL aircraft flowfields and aerodynamics. The computational results were all performed with hybrid RANS/LES formulations, with an emphasis on unsteady flows and subsequent aerodynamics, and results from five computational methods are included. The work was focused on one particular low-speed, high angle-of-attack flight test condition, and comparisons against flight-test data are included. This work represents the third coordinated effort using the F-16XL aircraft, and a unique flight-test data set, to advance our knowledge of slender airframe aerodynamics as well as our capability for predicting these aerodynamics with advanced CFD formulations. The prior efforts were identified as Cranked Arrow Wing Aerodynamics Project International, with the acronyms CAWAPI and CAWAPI-2.
    Keywords: Aerodynamics
    Type: NF1676L-24577 , Journal of Aircraft (ISSN 0021-8669) (e-ISSN 1533-3868); 54; 6; 2100-2114
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  • 16
    Publication Date: 2019-07-20
    Description: Precision landing of large payloads on Mars presents a challenge to the Entry, Descent, and Landing (EDL) community. Previous studies indicated that by incorporating the capability for a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) to morph during reentry would result in a more accurate landing footprint by allowing modulation of the lift- to-drag (L/D) vector directly instead of through bank angle control. However, morphing the HIAD shape for trajectory control may expose the HIAD to potential structural loads or aero heating concerns. In this study, the application of an optimal control allocation (OCA) technique was investigated that would to enable the morphing HIAD to maximize trajectory control capabilities while simultaneously keeping the structural loads and aero heating below some thresholds. This concept was demonstrated in a 3 degree-of-freedom (DOF) EDL simulation and provides basis for future research.
    Keywords: Aerodynamics
    Type: NF1676L-27448 , AIAA SciTech Forum 2018; Jan 08, 2018 - Jan 12, 2018; Kissimmee, FL; United States
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  • 17
    Publication Date: 2019-07-13
    Description: Differential drag is a technique for altering the semi-major axis, velocity, and along-track position of a spacecraft in low Earth orbit. It involves varying the spacecrafts cross-sectional area relative to its velocity direction by temporarily changing attitude and solar array angles, thus varying the amount of atmospheric drag on the spacecraft. The technique has recently been proposed and used by at least three satellite systems for initial separation of constellation spacecraft after launch, stationkeeping during the mission, and potentially for conjunction avoidance. Similarly, differential drag has been proposed as a control strategy for rendezvous, removing the need for active propulsion. In theory, some operational missions that lack propulsion capability could use this approach for conjunction avoidance, though options are typically constrained for spacecraft that are already in orbit. Shortly before the spacecraft was decommissioned, an experiment was performed using NASAs EO-1 spacecraft in order to demonstrate differential drag on an operational spacecraft in orbit, and discover some of the effects differential drag might manifest. EO-1 was not designed to maintain off-nominal orientations for long periods, and as a result the team experienced unanticipated challenges during the experiment. This paper will discuss operations limitations identified before the experiment, as well as those discovered during the experiment. The effective displacement that resulted from increasing the drag area for 39 hours will be compared to predictions as well as the expected position if the spacecraft maintained nominal operations. A hypothetical scenario will also be examined, studying the relative risks of maintaining an operational spacecraft bus in order to maintain the near-maximum drag area orientation and hasten reentry.
    Keywords: Aerodynamics
    Type: GSFC-E-DAA-TN47408 , IAA Conference on Space Situational Awareness (ICSSA); Nov 13, 2017 - Nov 15, 2017; Orlando, FL; United States
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  • 18
    Publication Date: 2019-07-19
    Description: The Adaptable Deployable Entry and Placement Technology (ADEPT) project will be conducting the first flight test of ADEPT, titled Sounding Rocket One (SR-1), in just two months. The need for this flight test stems from the fact that ADEPT's supersonic dynamic stability has not yet been characterized. The SR-1 flight test will provide critical data describing the flight mechanics of ADEPT in ballistic flight. These data will feed decision making on future ADEPT mission designs. This presentation will describe the SR-1 scientific data products, possible flight test outcomes, and the implications of those outcomes on future ADEPT development. In addition, this presentation will describe free-flight ground testing performed in advance of the flight test. A subsonic flight dynamics test conducted at the Vertical Spin Tunnel located at NASA Langley Research Center provided subsonic flight dynamics data at high and low altitudes for multiple center of mass (CoM) locations. A ballistic range test at the Hypervelocity Free Flight Aerodynamics Facility (HFFAF) located at NASA Ames Research Center provided supersonic flight dynamics data at low supersonic Mach numbers. Execution and outcomes of these tests will be discussed. Finally, a hypothesized trajectory estimate for the SR-1 flight will be presented.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN39602 , International Planetary Probe Workshop; Jun 12, 2017 - Jun 16, 2017; The Hague; Netherlands
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  • 19
    Publication Date: 2019-07-13
    Description: Low-Reynolds number testing was conducted at the 7 ft. x 10 ft. Walter H. Beech Memorial Wind Tunnel at Wichita State University to study the aerodynamic effects of ice shapes on a swept wing. A total of 17 ice shape configurations of varying geometric detail were tested. Simplified versions of an ice shape may help improve current ice accretion simulation methods and therefore aircraft design, certification, and testing. For each configuration, surface pressure, force balance, and fluorescent mini-tuft data were collected and for a selected subset of configurations oil-flow visualization and wake survey data were collected. A comparison of two ice shape geometries and two configurations with simplified geometric detail for each ice shape geometry is presented in this paper.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN42638 , AIAA Atmospheric and Space Environments Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 20
    Publication Date: 2019-07-13
    Description: Fully automated meshing for Reynolds-Averaged Navier-Stokes Simulations, Mesh generation for complex geometry continues to be the biggest bottleneck in the RANS simulation process; Fully automated Cartesian methods routinely used for inviscid simulations about arbitrarily complex geometry; These methods lack of an obvious & robust way to achieve near wall anisotropy; Goal: Extend these methods for RANS simulation without sacrificing automation, at an affordable cost; Note: Nothing here is limited to Cartesian methods, and much becomes simpler in a body-fitted setting.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN39522 , Advanced Modeling & Simulation (AMS) Seminar Series; Feb 23, 2017; Moffett Field, CA; United States
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  • 21
    Publication Date: 2019-07-13
    Description: This paper investigates the effect of nonlinear large deflection bending on the aerodynamic performance of a high aspect ratio flexible wing. A set of nonlinear static aeroelastic equations are derived for the large bending deflection of a high aspect ratio wing structure. An analysis is conducted to compare the nonlinear bending theory with the linear bending theory. The results show that the nonlinear bending theory is length-preserving whereas the linear bending theory causes a non-physical effect of lengthening the wing structure under the no axial load condition. A modified lifting line theory is developed to compute the lift and drag coefficients of a wing structure undergoing a large bending deflection. The lift and drag coefficients are more accurately estimated by the nonlinear bending theory due to its length-preserving property. The nonlinear bending theory yields lower lift and span efficiency than the linear bending theory. A coupled aerodynamic-nonlinear finite element model is developed to implement the nonlinear bending theory for a Common Research Model (CRM) flexible wing wind tunnel model to be tested in the University of Washington Aeronautical Laboratory (UWAL). The structural stiffness of the model is designed to give about 10% wing tip deflection which is large enough that could cause the nonlinear deflection effect to become significant. The computational results show that the nonlinear bending theory yields slightly less lift than the linear bending theory for this wind tunnel model. As a result, the linear bending theory is deemed adequate for the CRM wind tunnel model.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN42885 , AIAA Aviation Forum; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 22
    Publication Date: 2019-06-13
    Description: The Cranked-Arrow Wing Aerodynamics Project International investigation is continued with the FUN3D and USM3D flow solvers to fuse flight test, wind-tunnel test, and simulation of swept-wing aerodynamic features. Simulations of a low-speed, high-angle-of-attack condition are compared: detached-eddy simulation, modified delayed detached-eddy simulation, and the SpalartAllmaras Reynolds-averaged NavierStokes model. Isosurfaces of Q criterion show the development of coherent primary and secondary vortices on the upper surface of the wing that spiral, burst, and commingle. Mean detached-eddy simulation and modified delayed detached-eddy simulation pressures better predict the flight-test measurements than SpalartAllmaras model predictions, especially on the outer-wing section. The USM3D simulations predicted many sharp tones in volume point pressure spectra with low broadband noise, and the FUN3D simulations predicted more broadband noise with weaker tones. Spectra of the volume points near the outer-wing leading edge were primarily broadband for both codes. Time-averaged forces are very similar between FUN3D simulations and between USM3D simulations, but FUN3D predicts slightly higher lift and lower drag than USM3D. There is more variation in the pitching moment predictions. Spectra of the unsteady forces and moment are mostly broadband for FUN3D and tonal for USM3D simulations.
    Keywords: Aerodynamics
    Type: NF1676L-30239 , Journal of Aircraft (ISSN 0021-8669) (e-ISSN 1533-3868); 54; 6; 2027-2049
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  • 23
    Publication Date: 2019-08-24
    Description: An aircraft includes a propulsor supported within an aft portion of the fuselage. A thrust reverser is supported proximate the propulsor for redirecting thrust forward to slow the aircraft upon landing. A tail extending from the aft portion of the fuselage is angled forward away from the aft portion and out of the discharge of airflow from the thrust reverser.
    Keywords: Aerodynamics
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  • 24
    Publication Date: 2019-08-13
    Description: Computational fluid dynamics is now considered to be an indispensable tool for the design and development of scramjet engine components. Unfortunately, the quantification of uncertainties is rarely addressed with anything other than sensitivity studies, so the degree of confidence associated with the numerical results remains exclusively with the subject matter expert that generated them. This practice must be replaced with a formal uncertainty quantification process for computational fluid dynamics to play an expanded role in the system design, development, and flight certification process. Given the limitations of current hypersonic ground test facilities, this expanded role is believed to be a requirement by some in the hypersonics community if scramjet engines are to be given serious consideration as a viable propulsion system. The present effort describes a simple, relatively low cost, nonintrusive approach to uncertainty quantification that includes the basic ingredients required to handle both aleatoric (random) and epistemic (lack of knowledge) sources of uncertainty. The nonintrusive nature of the approach allows the computational fluid dynamicist to perform the uncertainty quantification with the flow solver treated as a "black box". Moreover, a large fraction of the process can be automated, allowing the uncertainty assessment to be readily adapted into the engineering design and development workflow. In the present work, the approach is applied to a model scramjet isolator problem where the desire is to validate turbulence closure models in the presence of uncertainty. In this context, the relevant uncertainty sources are determined and accounted for to allow the analyst to delineate turbulence model-form errors from other sources of uncertainty associated with the simulation of the facility flow.
    Keywords: Aerodynamics
    Type: NF1676L-27196 , JANNAF Joint Subcommittee Meeting; Dec 04, 2017 - Dec 08, 2017; Newport News, VA; United States
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  • 25
    Publication Date: 2019-08-13
    Description: This talk will provide an overview of investments in the Entry Systems Modeling project, along with some context of where the effort sits in the overall Space Technology EDL Portfolio. Technical highlights, particularly with referent to work on Ablation Modeling, will be given. Future directions will be discussed.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN46281 , Ablation Workshop; Aug 30, 2017 - Aug 31, 2017; Bozeman, MT; United States
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  • 26
    Publication Date: 2019-08-10
    Description: PADRI: A common platform for validation of aircraft drag reduction technologies; Generic strut-braced wing configuration; Slightly swept wing for low cruise Mach number (0.72); Simplified geometry without engines, empennage or flap-track fairings; Significant wave-drag and flow separation at strut-wing intersection; Focus of this workshop is to redesign the junction
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN49604 , Platform for Aircraft Drag Reduction Innovation (PADRI 2017); Nov 29, 2017 - Dec 01, 2017; Barcelona; Spain|European Community on Computational Methods in Applied Sciences (ECCOMAS) Advanced Course (EAC); Nov 29, 2017 - Dec 01, 2017; Barcelona; Spain
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  • 27
    Publication Date: 2019-08-08
    Description: The FAST-MAC circulation control model was modified to test an array of unsteady sweeping-jet actuators at realistic flight Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center. Two types of sweeping jet actuators were fabricated using rapid prototype techniques, and directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60, and a transonic cruise configuration with a 0 flap deflection. For the 30 flap high-lift configuration, the sweeping jets achieved comparable lift performance in the separation control regime, while reducing the mass flow by 54% as compared to steady blowing. However, the sweeping jets were not effective for the 60 flap. For the transonic cruise configuration, the sweeping jets reduced the drag by 3.3% at an off design condition. The drag reduction for the design lift coefficient for the sweeping jets provided only half the drag reduction shown for the steady blowing case (6.5%), but accomplished this with a 74% reduction in mass flow.
    Keywords: Aerodynamics
    Type: NF1676L-27684
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  • 28
    Publication Date: 2019-07-13
    Description: This paper demonstrates a technique for locating the optimal control surface layout of an aeroservoelastic Common Research Model wingbox, in the context of maneuver load alleviation and active utter suppression. The combinatorial actuator layout design is solved using ideas borrowed from topology optimization, where the effectiveness of a given control surface is tied to a layout design variable, which varies from zero (the actuator is removed) to one (the actuator is retained). These layout design variables are optimized concurrently with a large number of structural wingbox sizing variables and control surface actuation variables, in order to minimize the sum of structural weight and actuator weight. Results are presented that demonstrate interdependencies between structural sizing patterns and optimal control surface layouts, for both static and dynamic aeroelastic physics.
    Keywords: Aerodynamics
    Type: NF1676L-24456 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 29
    Publication Date: 2019-07-13
    Description: The computational fluid dynamics (CFD) prediction workshops sponsored by the AIAA have created invaluable opportunities in which to discuss the predictive capabilities of CFD in areas in which it has struggled, e.g., cruise drag, high-lift, and sonic boom pre diction. While there are many factors that contribute to disagreement between simulated and experimental results, such as modeling or discretization error, quantifying the errors contained in a simulation is important for those who make decisions based on the computational results. The linearized error transport equations (ETE) combined with a truncation error estimate is a method to quantify one source of errors. The ETE are implemented with a complex-step method to provide an exact linearization with minimal source code modifications to CFD and multidisciplinary analysis methods. The equivalency of adjoint and linearized ETE functional error correction is demonstrated. Uniformly refined grids from a series of AIAA prediction workshops demonstrate the utility of ETE for multidisciplinary analysis with a connection between estimated discretization error and (resolved or under-resolved) flow features.
    Keywords: Aerodynamics
    Type: NF1676L-24480 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 30
    Publication Date: 2019-07-13
    Description: This paper reports the findings from a study that applies wall-resolved large-eddy simulation to investigate flow separation over the NASA wall-mounted hump geometry. Despite its conceptually simple flow configuration, this benchmark problem has proven to be a challenging test case for various turbulence simulation methods that have attempted to predict flow separation arising from the adverse pressure gradient on the aft region of the hump. The momentum-thickness Reynolds number of the incoming boundary layer has a value that is near the upper limit achieved by recent direct numerical simulation and large-eddy simulation of incompressible turbulent boundary layers. The high Reynolds number of the problem necessitates a significant number of grid points for wall-resolved calculations. The present simulations show a significant improvement in the separation-bubble length prediction compared to Reynolds-Averaged Navier-Stokes calculations. The current simulations also provide good overall prediction of the skin-friction distribution, including the relaminarization observed over the front portion of the hump due to the strong favorable pressure gradient. We discuss a number of problems that were encountered during the course of this work and present possible solutions. A systematic study regarding the effect of domain span, subgrid-scale model, tunnel back pressure, upstream boundary layer conditions and grid refinement is performed. The predicted separation-bubble length is found to be sensitive to the span of the domain. Despite the large number of grid points used in the simulations, some differences between the predictions and experimental observations still exist (particularly for Reynolds stresses) in the case of the wide-span simulation, suggesting that additional grid resolution may be required.
    Keywords: Aerodynamics
    Type: NF1676L-24481 , AIAA SciTech Forum and Exposition 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 31
    Publication Date: 2019-07-13
    Description: This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing (BSCW) configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results show the effects of the temporal and spatial resolution, the coupling scheme between the flow and the structural solvers, and the initial excitation conditions on the numerical flutter onset. Depending on the free stream condition and the angle of attack, the above parameters do affect the flutter onset. Two conditions are analyzed: Mach 0.74 with angle of attack 0 and Mach 0.85 with angle of attack 5. The results are presented in the form of the damping values computed from the wing pitch angle response as a function of the dynamic pressure or in the form of dynamic pressure as a function of the Mach number.
    Keywords: Aerodynamics
    Type: NF1676L-24500 , AIAA SciTech 2017; Jan 09, 2017 - Jan 17, 2017; Grapevine, TX; United States
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  • 32
    Publication Date: 2019-07-13
    Description: A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.
    Keywords: Aerodynamics
    Type: NF1676L-24587 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 33
    Publication Date: 2019-07-13
    Description: The following details recent efforts undertaken at the NASA Ames Unitary Plan wind tunnels to design and deploy an advanced, production-level infrared (IR) flow visualization data system. Highly sensitive IR cameras, coupled with in-line image processing, have enabled the visualization of wind tunnel model surface flow features as they develop in real-time. Boundary layer transition, shock impingement, junction flow, vortex dynamics, and buffet are routinely observed in both transonic and supersonic flow regimes all without the need of dedicated ramps in test section total temperature. Successful measurements have been performed on wing-body sting mounted test articles, semi-span floor mounted aircraft models, and sting mounted launch vehicle configurations. The unique requirements of imaging in production wind tunnel testing has led to advancements in the deployment of advanced IR cameras in a harsh test environment, robust data acquisition storage and workflow, real-time image processing algorithms, and evaluation of optimal surface treatments. The addition of a multi-camera IR flow visualization data system to the Ames UPWT has demonstrated itself to be a valuable analyses tool in the study of new and old aircraft/launch vehicle aerodynamics and has provided new insight for the evaluation of computational techniques.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN35006 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 34
    Publication Date: 2019-07-16
    Description: Seal whiskers have been found to produce unique wake flow structures that minimize self-induced vibration and reduce drag. The cause of these wake features are due to the peculiar three-dimensional morphology of the whisker surface. The whisker can be defined as an elliptical cross section with variation in the major and minor axis of the ellipse along the length of the whisker as well as rotation of the elliptical plane with respect to the whisker axis, angle of incidence. While the effects of several dominant parameters of the whisker morphology have been studied, the effect of the rotation of the elliptical plane has not been well understood. This paper will examine the influence of the angle of incidence on the wake flow structure through series of water channel studies. Several models of whisker-like geometries will be tested which isolate the rotation angle as the only variation between models. The wake structure behind each seal whisker model will be measured through particle image velocimetry (PIV). The different models wake structures will be compared identifying the effect of angle of incidence on the wake structure. Angle of incidence was found to influence the wake structure through reorganization of velocity patterns, reduction of recovery length and modification of magnitude of Tu. These results help provide a more complete understanding of the seal whisker geometry relationship to wake structure and can provide insight into design practices for application of whisker geometery to various engineering problems.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN52934 , AIAA Aerospace Sciences Meeting 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 35
    Publication Date: 2019-07-13
    Description: BOS for flight test focusing on aircraft operations and piloting.
    Keywords: Aerodynamics
    Type: AFRC-E-DAA-TN43978 , SETP European Symposium; Jul 11, 2017 - Jul 14, 2017; Bristol; United Kingdom
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  • 36
    Publication Date: 2019-07-13
    Description: Contributions of the unstructured Reynolds-averaged Navier-Stokes code, FUN3D, to the 3rd AIAA Propulsion Aerodynamics Workshop are described for the diffusing IFCPT S-Duct. Using workshop-supplied grids, results for the baseline S-Duct, baseline S-Duct with Aerodynamic Interface Plane (AIP) rake hardware, and baseline S-Duct with flow control devices are compared with experimental data and results computed with output-based, off-body grid adaptation in FUN3D. Due to the absence of influential geometry components, total pressure recovery is overpredicted on the baseline S-Duct and S-Duct with flow control vanes when compared to experimental values. An estimate for the exact value of total pressure recovery is derived for these cases given an infinitely refined mesh. When results from output-based mesh adaptation are compared with those computed on workshop-supplied grids, a considerable improvement in predicting total pressure recovery is observed. By including more representative geometry, output-based mesh adaptation compares very favorably with experimental data in terms of predicting the total pressure recovery cost-function; whereas, results computed using the workshop-supplied grids are underpredicted.
    Keywords: Aerodynamics
    Type: NF1676L-27446 , 2017 AIAA Propulsion and Energy Forum and Exposition; Jul 10, 2017 - Jul 12, 2017; Atlanta, GA; United States
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  • 37
    Publication Date: 2019-07-13
    Description: This paper studies the whirl flutter stability of the NASA experimental electric propulsion aircraft designated the X-57 Maxwell. whirl flutter stability is studied at two flight conditions: sea level at 2700 RPM to represent take-off and landing and 8000 feet at 2250 RPM to represent cruise. Two multibody dynamics analyses are used: CAMRAD II and Dymore. The CAMRAD II model is a semi-span X-57 model with a modal representation for the wing/pylon system. The Dymore model is a semi-span wing with a propeller composed of beam elements for the wing/pylon system that airloads can be applied to. The two multibody dynamics analyses were verified by comparing structural properties between each other and the NASTRAN analysis. For whirl flutter, three design revisions of the wing and pylon mount system are studied. The predicted frequencies and damping ratio of the wing modes show good agreements between the two analyses. Dymore tended to predict a slightly lower damping ratio as velocity increased for all three dynamic modes presented. Whirl flutter for the semi-span model was not present up to 500 knots for the latest design, well above the operating range of the X-57.
    Keywords: Aerodynamics
    Type: NF1676L-25815 , AIAA Aviation 2017 Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 38
    Publication Date: 2019-07-13
    Description: A computational study of a distributed electric propulsion wing with a 40deg flap deflection has been completed using FUN3D. Two lift-augmentation power conditions were compared with the power-off configuration on the high-lift wing (40deg flap) at a 73 mph freestream flow and for a range of angles of attack from -5 degrees to 14 degrees. The computational study also included investigating the benefit of corotating versus counter-rotating propeller spin direction to powered-lift performance. The results indicate a large benefit in lift coefficient, over the entire range of angle of attack studied, by using corotating propellers that all spin counter to the wingtip vortex. For the landing condition, 73 mph, the unpowered 40deg flap configuration achieved a maximum lift coefficient of 2.3. With high-lift blowing the maximum lift coefficient increased to 5.61. Therefore, the lift augmentation is a factor of 2.4. Taking advantage of the fullspan lift augmentation at similar performance means that a wing powered with the distributed electric propulsion system requires only 42 percent of the wing area of the unpowered wing. This technology will allow wings to be 'cruise optimized', meaning that they will be able to fly closer to maximum lift over drag conditions at the design cruise speed of the aircraft.
    Keywords: Aerodynamics
    Type: NF1676L-25692 , AIAA Aviation 2017 Forum; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 39
    Publication Date: 2019-07-13
    Description: Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on the iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind-tunnel testing and computational flow simulations were carried out for an 8.9%-scale semispan wing based upon the Common Research Model airplane configuration. The wind-tunnel testing was conducted at the Wichita State University 7 ft x 10 ft Beech wind tunnel from Reynolds numbers of 0.810(exp 6) to 2.410(exp 6) and corresponding Mach numbers of 0.09 to 0.27. This paper presents the results of initial studies investigating the model mounting configuration, clean-wing aerodynamics and effects of artificial ice roughness. Four different model mounting configurations were considered and a circular splitter plate combined with a streamlined shroud was selected as the baseline geometry for the remainder of the experiments and computational simulations. A detailed study of the clean-wing aerodynamics and stall characteristics was made. In all cases, the flow over the outboard sections of the wing separated as the wing stalled with the inboard sections near the root maintaining attached flow. Computational flow simulations were carried out with the ONERA elsA software that solves the compressible, three-dimensional RANS equations. The computations were carried out in either fully turbulent mode or with natural transition. Better agreement between the experimental and computational results was obtained when considering computations with free transition compared to turbulent solutions. These results indicate that experimental evolution of the clean wing performance coefficients were due to the effect of three-dimensional transition location and that this must be taken into account for future data analysis. This research also confirmed that artificial ice roughness created with rapid-prototype manufacturing methods can generate aerodynamic performance effects comparable to grit roughness of equivalent size when proper care is exercised in design and installation. The conclusions of this combined experimental and computational study contributed directly to the successful implementation of follow-on test campaigns with numerous artificial ice-shape configurations for this 8.9% scale model.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN42235 , AIAA Atmospheric and Space Environments Conference 2017; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 40
    Publication Date: 2019-07-13
    Description: The Orion Multi-purpose Crew Vehicle (MPCV) Orion spacecraft will return humans from beyond earth's orbit, including Mars and will be required to land 20,000 pounds of mass safely in the ocean. The parachute system nominally lands under 3 main parachutes, but the system is designed to be fault tolerant and land under 2 main parachutes. During several of the parachute development tests, it was observed that a pendulum, or swinging, motion could develop while the Crew Module (CM) was descending under two parachutes. This pendulum effect had not been previously predicted by modeling. Landing impact analysis showed that the landing loads would double in some places across the spacecraft. The CM structural design limits would be exceeded upon landing if this pendulum motion were to occur. The Orion descent and landing team was faced with potentially millions of dollars in structural modifications and a severe mass increase. A multidisciplinary team was formed to determine root cause, model the pendulum motion, study alternate canopy planforms and assess alternate operational vehicle controls & operations providing mitigation options resulting in a reliability level deemed safe for human spaceflight. The problem and solution is a balance of risk to a known solution versus a chance to improve the landing performance for the next human-rated spacecraft.
    Keywords: Aerodynamics
    Type: JSC-CN-39169 , Aerodynamic Decelerator Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 41
    Publication Date: 2019-07-13
    Description: The Hyperloop concept is proposed as a faster, cheaper alternative to high-speed rail and traditional short-haul aircraft. It consists of a passenger pod traveling through a tube under light vacuum while being propelled and levitated by a combination of permanent and electro-magnets. The concept addresses NASA's research thrusts for growth in demand, sustainability, and technology convergence for high-speed transport. Hyperloop is a radical departure from other advanced aviation concepts, however it remains an aeronautics concept that tackles the same strategic goals of low-carbon propulsion and ultra-effcient vehicles. System feasibility was investigated by building a multidisciplinary vehicle sizing model that takes into account aerodynamic, thermodynamic, structures, electromagnetic, weight, and mission analyses. The sizing process emphasized the strong coupling between the two largest systems: the tube and the passenger pod. The model was then exercised to examine Hyperloop from a technical and cost perspective. The structural sizing analysis of the travel tube demonstrates potential for signi cant capital cost reductions by considering an underwater route. Examination of varying passenger capacity indicates that the system can be operated with a wide range of passenger loads without significant change in operating expenses. Lastly, a high-level sizing study simulated variations in tube area, pressure, pod speed, and passenger capacity showing that there is a tube pressure that minimizes operating energy usage. The value of this optimal tube pressure is highly sensitive to numerous design details. These combined estimates of energy consumption, passenger throughput, and mission analyses all support Hyperloop as a faster and cheaper alternative to short-haul flights. The tools and expertise used to quantify these results also demonstrate how traditional aerospace design methods can be leveraged to handle the complex and coupled design process. Much of the technology development required for the Hyperloop is shared with next-generation aircraft. Furthermore, the substantial public interest and active commercial development make it an ideal candidate as an aircraft technology driver and test bed.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN37945 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 42
    Publication Date: 2019-07-13
    Description: Discrepancies between the model and balance coordinate systems lead to biases in the aerodynamic measurements during wind-tunnel testing. The reference coordinate system relative to the calibration coordinate system at which the forces and moments are resolved is crucial to the overall accuracy of force measurements. This paper discusses sources of discrepancies and estimates of coordinate system rotation and translation due to machining and assembly differences. A methodology for numerically estimating the coordinate system biases will be discussed and developed. Two case studies are presented using this methodology to estimate the model alignment. Examples span from angle measurement system shifts on the calibration system to discrepancies in actual wind-tunnel data. The results from these case-studies will help aerodynamic researchers and force balance engineers to better the understand and identify potential differences in calibration systems due to coordinate system rotation and translation.
    Keywords: Aerodynamics
    Type: NF1676L-25870 , AIAA Aviation Technology, Integration, and Operations Conference (AVIATION 2017); Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 43
    Publication Date: 2019-07-13
    Description: A third wind tunnel test of the FAST-MAC circulation control semi-span model was completed in the National Transonic Facility at the NASA Langley Research Center where the model was configured for transonic testing of the cruise configuration with 0deg flap detection to determine the potential for transonic drag reduction with the circulation control blowing. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged ap. Recent upgrades to transonic semi-span flow control testing at the NTF have demonstrated an improvement to overall data repeatability, particularly for the drag measurement, that allows for increased confidence in the data results. The static thrust generated by the blowing slot was removed from the wind-on data using force and moment balance data from wind-o thrust tares. This paper discusses the impact of the trailing-edge blowing to the transonic aerodynamics of the FAST-MAC model in the cruise configuration, where at flight Reynolds numbers, the thrust-removed corrected data showed that an overall drag reduction and increased aerodynamic efficiency was realized as a consequence of the blowing.
    Keywords: Aerodynamics
    Type: NF1676L-25806 , AIAA Aviation 2017 Forum; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 44
    Publication Date: 2019-07-13
    Description: There is a drive to devise acoustic treatments with reduced skin-friction and therefore fuel-burn penalty for engine nacelles on commercial airplanes. The studies have been experimental, and the effects on skin-friction are deduced from measurements of the pressure drop along a duct. We conduct a detailed CFD analysis of the installation, for two purposes. The first is to predict the effects of the finite size of the rig, including its near-square cross-section and the moderate length of the treated patch; this introduces transient and blockage effects, which have not been included so far in the analysis. In addition, the flow is compressible, so that even with homogeneous surface conditions, it is not homogeneous in the streamwise direction. The second purpose is to extract an effective sand-grain roughness size for a particular liner, which in turn can be used in a CFD analysis of the aircraft, leading to actual predictions of the effect of acoustic treatments on fuel burn in service. The study is entirely based on classical turbulence models, with an appropriate modification for effective roughness effects, rather than directly modeling the liners.
    Keywords: Aerodynamics
    Type: NF1676L-25719 , AIAA Aviation 2017 Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 45
    Publication Date: 2019-07-13
    Description: This work considers an aeroelastic wingbox model seeded with run-out blade stiffeners along the skins. Topology optimization is conducted within the shell webs of the stiffeners, in order to add cutouts and holes for mass reduction. This optimization is done with a global-local approach in order to moderate the computational cost: aeroelastic loads are computed at the wing-level, but the topology and sizing optimization is conducted at the panel-level. Each panel is optimized separately under stress, buckling, and adjacency constraints, and periodically reassembled to update the trimmed aeroelastic loads. The resulting topology is baselined against a design with standard full-depth solid stiffener blades, and found to weigh 7.43% less.
    Keywords: Aerodynamics
    Type: NF1676L-25663 , AIAA Aviation Technology, Integration, and Operations Conference (AVIATION 2017); Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 46
    Publication Date: 2019-07-13
    Description: Direct numerical simulations (DNS) of flow over an NACA-0012 airfoil are performed at a low and a moderate Reynolds numbers of Re(sub c)=50 times10(exp 3) and 1times 10(exp 6). The angles of attack are 5 and 15 degrees at the low and the moderate Reynolds number cases respectively. The three-dimensional unsteady compressible Navier-Stokes equations are solved using higher order compact schemes. The flow field in the low Reynolds number case consists of a long separation bubble near the leading-edge region and an attached boundary layer on the aft part of the airfoil. The shear layer that formed in the separated region persisted up to the end of the airfoil. The roles of the turbulent diffusion, advection, and dissipation terms in the turbulent kinetic-energy balance equation change as the boundary layer evolves over the airfoil. In the higher Reynolds number case, the leading-edge separation bubble is very small in length and in height. A fully developed turbulent boundary layer is observed in a short distance downstream of the reattachment point. The boundary layer velocity near the wall gradually decreases along the airfoil. Eventually, the boundary layer separates near the trailing edge. The Reynolds stresses peak in the outer part of the boundary layer and the maximum amplitude also gradually increases along the chord.
    Keywords: Aerodynamics
    Type: NF1676L-25674 , AIAA Aviation Technology, Integration, and Operations Conference (AVIATION 2017); Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 47
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    In:  CASI
    Publication Date: 2019-07-13
    Description: NASA is in the process of qualifying the mid-size Super Pressure Balloon (SPB) to provide constant density altitude flight for science investigations at polar and mid-latitudes. The status of the development of the 18.8 million cubic foot SPB capable of carrying one-tone of science to 110,000 feet, will be given. In addition, the operating considerations such as launch sites, flight safety considerations, and recovery will be discussed.
    Keywords: Aerodynamics
    Type: GSFC-E-DAA-TN42724 , Scientific Ballooning Technologies Workshop 2017; May 16, 2017 - May 18, 2017; Minneapolis, MN; United States
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  • 48
    Publication Date: 2019-07-13
    Description: A 24 foot diameter 3-stage axial compressor powered by variable-speed induction motors provides the airflow in the closed-return 11- by 11-Foot Transonic Wind Tunnel (11-Foot TWT) Facility at NASA Ames Research Center at Moffett Field, California. The facility is part of the Unitary Plan Wind Tunnel, which was completed in 1955. Since then, upgrades made to the 11-Foot TWT such as flow conditioning devices and instrumentation have increased blockage and pressure loss in the tunnel, somewhat reducing the peak Mach number capability of the test section. Due to erosion effects on the existing aluminum alloy rotor blades, fabrication of new steel rotor blades is planned. This presents an opportunity to increase the Mach number capability of the tunnel by redesigning the compressor for increased pressure ratio. Challenging design constraints exist for any proposed design, demanding the use of the existing driveline, rotor disks, stator vanes, and hub and casing flow paths, so as to minimize cost and installation time. The current effort was undertaken to characterize the performance of the existing compressor design using available design tools and computational fluid dynamics (CFD) codes and subsequently recommend a new compressor design to achieve higher pressure ratio, which directly correlates with increased test section Mach number. The constant cross-sectional area of the compressor leads to highly diffusion factors, which presents a challenge in simulating the existing design. The CFD code APNASA was used to simulate the aerodynamic performance of the existing compressor. The simulations were compared to performance predictions from the HT0300 turbomachinery design and analysis code, and to compressor performance data taken during a 1997 facility test. It was found that the CFD simulations were sensitive to endwall leakages associated with stator buttons, and to a lesser degree, under-stator-platform flow recirculation at the hub. When stator button leakages were modeled, pumping capability increased by over 20 of pressure rise at design point due to a large reduction in aerodynamic blockage at the hub. Incorporating the stator button leakages was crucial to matching test data. Under-stator-platform flow recirculation was thought to be large due to a lack of seals. The effect of this recirculation was assessed with APNASA simulations recirculating 0.5, 1, and 2 of inlet flow about stators 1 and 2, modeled as axisymmetric mass flux boundary conditions on the hub before and after the vanes. The injection of flow ahead of the stators tended to re-energize the boundary layer and reduce hub separations, resulting in about 3 increased stall margin per 1 of inlet flow recirculated. In order to assess the value of the flow recirculation, a mixing plane simulation of the compressor which gridded the under-stator cavities was generated using the ADPAC CFD code. This simulation indicated that about 0.65 of the inlet flow is recirculated around each shrouded stator. This collective information was applied during the redesign of the compressor. A potential design was identified using HT0300 which improved overall pressure ratio by removing pre-swirl into rotor 1, replacing existing NASA 65 series rotors with double circular arc sections, and re-staggering rotors and the existing stators. The performance of the new design predicted by APNASA and HT0300 is compared to the existing design.
    Keywords: Aerodynamics
    Type: GT2017-65139 , GRC-E-DAA-TN40518 , Turbomachinery Technical Conference & Exposition; Jun 26, 2017 - Jun 30, 2017; Charlotte, NC; United States
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  • 49
    Publication Date: 2019-07-13
    Description: A preliminary natural laminar flow (NLF) design method that has been developed and applied to supersonic and transonic wings with moderate-to-high leading-edge sweeps at flight Reynolds numbers is further extended and evaluated in this paper. The modular design approach uses a knowledge-based design module linked with different flow solvers and boundary layer stability analysis methods to provide a multifidelity capability for NLF analysis and design. An assessment of the effects of different options for stability analysis is included using pressures and geometry from an NLF wing designed for the Common Research Model (CRM). Several extensions to the design module are described, including multiple new approaches to design for controlling attachment line contamination and transition. Finally, a modification to the NLF design algorithm that allows independent control of Tollmien-Schlichting (TS) and cross flow (CF) modes is proposed. A preliminary evaluation of the TS-only option applied to the design of an NLF nacelle for the CRM is performed that includes the use of a low-fidelity stability analysis directly in the design module.
    Keywords: Aerodynamics
    Type: NF1676L-25627 , AIAA Aviation 2017 Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 50
    Publication Date: 2019-07-13
    Description: A natural laminar flow (NLF) wind tunnel model has been designed and analyzed for a wind tunnel test in the National Transonic Facility (NTF) at the NASA Langley Research Center. The NLF design method is built into the CDISC design module and uses a Navier-Stokes flow solver, a boundary layer profile solver, and stability analysis and transition prediction software. The NLF design method alters the pressure distribution to support laminar flow on the upper surface of wings with high sweep and flight Reynolds numbers. The method addresses transition due to attachment line contamination/transition, Gortler vortices, and crossflow and Tollmien-Schlichting modal instabilities. The design method is applied to the wing of the Common Research Model (CRM) at transonic flight conditions. Computational analysis predicts significant extents of laminar flow on the wing upper surface, which results in drag savings. A 5.2 percent scale semispan model of the CRM NLF wing will be built and tested in the NTF. This test will aim to validate the NLF design method, as well as characterize the laminar flow testing capabilities in the wind tunnel facility.
    Keywords: Aerodynamics
    Type: AIAA Paper 2017-3058 , NF1676L-25543 , AIAA Aviation 2017; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 51
    Publication Date: 2019-07-13
    Description: A two-week test campaign was conducted in the National Full-Scale Aerodynamics Complex 80 x 120-ft Wind Tunnel in support of Orion parachute pendulum mitigation activities. The test gathered static aerodynamic data using an instrumented, 3-tether system attached to the parachute vent in combination with an instrumented parachute riser. Dynamic data was also gathered by releasing the tether system and measuring canopy performance using photogrammetry. Several canopy configurations were tested and compared against the current Orion parachute design to understand changes in drag performance and aerodynamic stability. These configurations included canopies with varying levels and locations of geometric porosity as well as sails with increased levels of fullness. In total, 37 runs were completed for a total of 392 data points. Immediately after the end of the testing campaign a down-select decision was made based on preliminary data to support follow-on sub-scale air drop testing. A summary of a more rigorous analysis of the test data is also presented.
    Keywords: Aerodynamics
    Type: JSC-CN-39271 , AIAA Aerodynamics Decelerator Systems Technology Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 52
    Publication Date: 2019-07-13
    Description: Capsule Parachute Assembly System (CPAS) flight tests regularly stage a helicopter to observe inflation of 116 ft D o ringsail Main parachutes. These side views can be used to generate 3-D models of inflating canopies to estimate enclosed volume. Assuming a surface of revolution is inadequate because reefed canopies in a cluster are elongated due to mutual aerodynamic interference. A method was developed to combine the side views with upward looking HD video to account for non-circular cross sections. Approximating the cross sections as elliptical greatly improves accuracy. But since that correction requires manually tracing projected outlines, the actual irregular shapes can be used to generate high fidelity models. Compensation is also made for apparent tilt angle. Validation was accomplished by comparing perimeter and projected area with known line lengths and/or high quality photogrammetry.
    Keywords: Aerodynamics
    Type: JSC-CN-38970 , AIAA Aerodynamic Decelerator Systems Technology Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 53
    Publication Date: 2019-07-13
    Description: Air Data Systems (FADS) are becoming more prevalent on re-entry vehicles, as evi- denced by the Mars Science Laboratory and the Orion Multipurpose Crew Vehicle. A FADS consists of flush-mounted pressure transducers located at various locations on the fore-body of a flight vehicle or the heat shield of a re-entry capsule. A pressure model converts the pressure readings into useful air data quantities. Two algorithms for converting pressure readings to air data have become predominant- the iterative Least Squares State Estimator (LSSE) and the Triples Algorithm. What follows herein is a new algorithm that takes advantage of the best features of both the Triples Algorithm and the LSSE. This approach employs the potential flow model and strategic differencing of the Triples Algorithm to obtain the defective flight angles; however, the requirements on port placement are far less restrictive, allowing for configurations that are considered optimal for a FADS.
    Keywords: Aerodynamics
    Type: JSC-CN-38856 , AIAA Aviation forum; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 54
    Publication Date: 2019-07-13
    Description: Results from the Sixth AIAA CFD Drag Prediction Workshop Common Research Model Cases 2 to 5 are presented. As with past workshops, numerical calculations are performed using industry-relevant geometry, methodology, and test cases. Cases 2 to 5 focused on force/moment and pressure predictions for the NASA Common Research Model wing-body and wing-body-nacelle-pylon configurations, including Case 2 - a grid refinement study and nacelle-pylon drag increment prediction study; Case 3 - an angle-of-attack buffet study; Case 4 - an optional wing-body grid adaption study; and Case 5 - an optional wing-body coupled aero-structural simulation. The Common Research Model geometry differed from previous workshops in that it was deformed to the appropriate static aeroelastic twist and deflection at each specified angle-of-attack. The grid refinement study used a common set of overset and unstructured grids, as well as user created Multiblock structured, unstructured, and Cartesian based grids. For the supplied common grids, six levels of refinement were created resulting in grids ranging from 7x10(exp 6) to 208x10(exp 6) cells. This study (Case 2) showed further reduced scatter from previous workshops, and very good prediction of the nacelle-pylon drag increment. Case 3 studied buffet onset at M=0.85 using the Medium grid (20 to 40x10(exp 6) nodes) from the above described sequence. The prescribed alpha sweep used finely spaced intervals through the zone where wing separation was expected to begin. Although the use of the prescribed aeroelastic twist and deflection at each angle-of-attack greatly improved the wing pressure distribution agreement with test data, many solutions still exhibited premature flow separation. The remaining solutions exhibited a significant spread of lift and pitching moment at each angle-of-attack, much of which can be attributed to excessive aft pressure loading and shock location variation. Four Case 4 grid adaption solutions were submitted. Starting with grids less than 2x10(exp 6) grid points, two solutions showed a rapid convergence to an acceptable solution. Four Case 5 coupled aerostructural solutions were submitted. Both showed good agreement with experimental data. Results from this workshop highlight the continuing need for CFD improvement, particularly for conditions with significant flow separation. These comparisons also suggest the need for improved experimental diagnostics to guide future CFD development.
    Keywords: Aerodynamics
    Type: NF1676L-26060 , 2017 AIAA SciTech; Jan 09, 2017 - Jan 13, 2017; Dallas, TX; United States
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  • 55
    Publication Date: 2019-07-13
    Description: We demonstrate a novel algorithm for computing the sensitivity of statistics in chaotic flow simulations to parameter perturbations. The algorithm is non-intrusive but requires exposing an interface. Based on the principle of shadowing in dynamical systems, this algorithm is designed to reduce the effect of the sampling error in computing sensitivity of statistics in chaotic simulations. We compare the effectiveness of this method to that of the conventional finite difference method.
    Keywords: Aerodynamics
    Type: NF1676L-26102 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 56
    Publication Date: 2019-07-13
    Description: A NATO symposium held in 2008 identified many promising sensitivity analysis and un-certainty quantification technologies, but the maturity and suitability of these methods for realistic applications was not known. The STO Task Group AVT-191 was established to evaluate the maturity and suitability of various sensitivity analysis and uncertainty quantification methods for application to realistic problems of interest to NATO. The program ran from 2011 to 2015, and the work was organized into four discipline-centric teams: external aerodynamics, internal aerodynamics, aeroelasticity, and hydrodynamics. This paper presents an overview of the AVT-191 program content.
    Keywords: Aerodynamics
    Type: NF1676L-24607 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 57
    Publication Date: 2019-07-13
    Description: A NATO symposium held in Greece in 2008 identified many promising sensitivity analysis and uncertainty quantification technologies, but the maturity and suitability of these methods for realistic applications was not clear. The NATO Science and Technology Organization, Task Group AVT-191 was established to evaluate the maturity and suitability of various sensitivity analysis and uncertainty quantification methods for application to realistic vehicle development problems. The program ran from 2011 to 2015, and the work was organized into four discipline-centric teams: external aerodynamics, internal aerodynamics, aeroelasticity, and hydrodynamics. This paper summarizes findings and lessons learned from the task group.
    Keywords: Aerodynamics
    Type: NF1676L-24614 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapvine, TX; United States
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  • 58
    Publication Date: 2019-07-13
    Description: The ability to re-enter the atmosphere at a desired location is important for spacecraft containing components that may survive re-entry. While impact point targeting has traditionally been initiated through impulsive burns with chemical thrusters on large vehicles such as the Space Shuttle, and the Soyuz and Apollo capsules, many small spacecraft do not host thrusters and require an alternative means of impact point targeting to ensure that falling debris do not cause harm to persons or property. This paper discusses the use of solely aerodynamic drag force to perform this targeting. It is shown that by deploying and retracting a drag device to vary the ballistic coefficient of the spacecraft, any desired longitude and latitude on the ground can be targeted provided that the maneuvering begins early enough and the latitude is less than the inclination of the orbit. An analytical solution based on perturbations from a numerically propagated trajectory is developed to map the initial state and ballistic coefficient profile of a spacecraft to its impact point. This allows the ballistic coefficient profile necessary to reach a given target point to be rapidly calculated, making it feasible to generate the guidance for the decay trajectory onboard the spacecraft. The ability to target an impact point using aerodynamic drag will enhance the capabilities of small spacecraft and will enable larger space vehicles containing thrusters to save fuel by more effectively leveraging the available aerodynamic drag.
    Keywords: Aerodynamics
    Type: KSC-E-DAA-TN36411 , AIAA Science and Technology Forum (SciTech 2017); Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 59
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    In:  CASI
    Publication Date: 2019-08-26
    Description: An aircraft includes a fuselage including a propulsion system supported within an aft portion. A thrust reverser is mounted proximate to the propulsion system for directing thrust in a direction to slow the aircraft. The thrust reverser directs thrust at an angle relative to a vertical plane to reduce interference on control surfaces and reduce generation of underbody lift.
    Keywords: Aerodynamics
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  • 60
    Publication Date: 2019-07-12
    Description: This chapter describes a new intelligent platform for learning optimal designs of morphing wings based on Variable Camber Continuous Trailing Edge Flaps (VCCTEF) in conjunction with a leading edge flap called the Variable Camber Krueger (VCK). The new platform consists of a Computational Fluid Dynamics (CFD) methodology coupled with a semi-supervised learning methodology. The CFD component of the intelligent platform comprises of a full Navier-Stokes solution capability (NASA OVERFLOW solver with Spalart-Allmaras turbulence model) that computes flow over a tri-element inboard NASA Generic Transport Model (GTM) wing section. Various VCCTEF/VCK settings and configurations were considered to explore optimal design for high-lift flight during take-off and landing. To determine globally optimal design of such a system, an extremely large set of CFD simulations is needed. This is not feasible to achieve in practice. To alleviate this problem, a recourse was taken to a semi-supervised learning (SSL) methodology, which is based on manifold regularization techniques. A reasonable space of CFD solutions was populated and then the SSL methodology was used to fit this manifold in its entirety, including the gaps in the manifold where there were no CFD solutions available. The SSL methodology in conjunction with an elastodynamic solver (FiDDLE) was demonstrated in an earlier study involving structural health monitoring. These CFD-SSL methodologies define the new intelligent platform that forms the basis for our search for optimal design of wings. Although the present platform can be used in various other design and operational problems in engineering, this chapter focuses on the high-lift study of the VCK-VCCTEF system. Top few candidate design configurations were identified by solving the CFD problem in a small subset of the design space. The SSL component was trained on the design space, and was then used in a predictive mode to populate a selected set of test points outside of the given design space. The new design test space thus populated was evaluated by using the CFD component by determining the error between the SSL predictions and the true (CFD) solutions, which was found to be small. This demonstrates the proposed CFD-SSL methodologies for isolating the best design of the VCK-VCCTEF system, and it holds promise for quantitatively identifying best designs of flight systems, in general.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN37043
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  • 61
    Publication Date: 2019-07-12
    Description: This manual describes the installation and execution of FUN3D version 13.2, including optional dependent packages. FUN3D is a suite of computational fluid dynamics simulation and design tools that uses mixed-element unstructured grids in a large number of formats, including structured multiblock and overset grid systems. A discretely-exact adjoint solver enables efficient gradient-based design and grid adaptation to reduce estimated discretization error. FUN3D is available with and without a reacting, real-gas capability. This generic gas option is available only for those persons that qualify for its beta release status.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219661 , L-20872 , NF1676L-28058
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  • 62
    Publication Date: 2019-07-12
    Description: This paper identifies the unsteady aerodynamic forces and moments for a typical section contained in the NACA Report No. 496, "General Theory of Aerodynamic Instability and the Mechanism of Flutter," by Theodore Theodorsen. These quantities are named Theodorsen's aerodynamic forces (TAFs). The TAFs are compared to the generalized aerodynamic forces (GAFs) for a very high aspect ratio wing (AR = 20) at zero Mach number computed by the doublet lattice method. Agreement between TAFs and GAFs is very-good-to-excellent. The paper also reveals that simple proportionality relationships that are known to exist between the real parts of some GAFs and the imaginary parts of others also hold for the real and imaginary parts of the corresponding TAFs.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219667 , L-20873 , NF1676L-28080
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  • 63
    Publication Date: 2019-07-12
    Description: The near wake of a flat plate is investigated via direct numerical simulations (DNS). Many earlier experimental investigations have used thin plates with sharp trailing edges and turbulent boundary layers to create the wake. This results in large theta divided by D (sub TE) values (theta is the boundary layer momentum thickness towards the end of the plate and D (sub TE) is the trailing edge thickness). In the present study the emphasis is on relatively thick plates with circular trailing edges (CTE) resulting in theta divided by D values less than one (D is the plate thickness and the diameter of the CTE), and vigorous vortex shedding. The Reynolds numbers based on the plate length and D are 1.255 x 10 (sup 6) and 10,000, respectively. Two cases are computed; one with turbulent boundary layers on both the upper and lower surfaces of the plate (statistically the same, symmetric wake, Case TT) and, a second with turbulent and laminar boundary layers on the upper and lower surfaces, respectively (asymmetric case, Case TL). The data and understanding obtained is of considerable engineering interest, particularly in turbomachinery where the pressure side of an airfoil can remain laminar or transitional because of a favorable pressure gradient and the suction side is turbulent. Shed-vortex structure and phase-averaged velocity statistics obtained in the two cases are compared here. The upper negative shed vortices in Case TL (turbulent separating boundary layer) are weaker than the lower positive ones (laminar separating boundary layer) at inception (a factor 1.27 weaker in terms of peak phase-averaged spanwise vorticity at first appearance of a peak). The upper vortices weaken rapidly as they travel downstream. A second feature of interest in Case TL is a considerable increase in the peak phase-averaged, streamwise normal intensity (random component) with increasing streamwise distance (x divided by D) that occurs nears the positive vortex cores. This behavior is observed for a few diameters in the near wake. This is counter to Case TT where the peak value essentially decreases with increasing x divided by D. Both these effects are examined in detail and the important contributors are identified.
    Keywords: Aerodynamics
    Type: NASA/TM-2017-219522 , ARC-E-DAA-TN43149
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  • 64
    Publication Date: 2019-07-13
    Description: The Mars Atmosphere and Volatile EvolutioN (MAVEN) spacecraft is a NASA orbiter designed to explore the Mars upper atmosphere, typically from 140 to 160 km altitude. In addition to the nominal science mission, MAVEN has performed several Deep Dip campaigns in which the orbit's closest point of approach, also called periapsis, was lowered to an altitude range of 115 to 135 km. MAVEN accelerometer data were used during mission operations to estimate atmospheric parameters such as density, scale height, along-track gradients, and wave structures. Density and scale height estimates were compared against those obtained from the Mars Global Reference Atmospheric Model and used to aid the MAVEN navigation team in planning maneuvers to raise and lower periapsis during Deep Dip operations. This paper describes the processes used to reconstruct atmosphere parameters from accelerometers data and presents the results of their comparison to model and navigation-derived values.
    Keywords: Aerodynamics
    Type: AAS Paper 17-273 , NF1676L-25544 , AAS/AIAA Space Flight Mechanics Meeting; Feb 05, 2017 - Feb 09, 2017; San Antonio, TX; United States
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  • 65
    Publication Date: 2019-07-13
    Description: Airframe-propulsion integration concepts that use boundary layer ingestion have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA's Starc-ABL aircraft configuration, which offers the potential for 12% mission fuel burn reduction by using a turbo-electric propulsion system with an aft-mounted electrically driven boundary layer ingestion propulsor. This large potential for improved performance motivates a more detailed study of the boundary layer ingestion propulsor design, but to date, analyses of boundary layer ingestion have used uncoupled methods. These methods account for only aerodynamic effects on the propulsion system or propulsion system effects on the aerodynamics, but not both simultaneously. This work presents a new approach for building fully coupled propulsive-aerodynamic models of boundary layer ingestion propulsion systems. A 1D thermodynamic cycle analysis is coupled to a RANS simulation to model the Starc-ABL aft propulsor at a cruise condition and the effects variation in propulsor design on performance are examined. The results indicates that both propulsion and aerodynamic effects contribute equally toward the overall performance and that the fully coupled model yields substantially different results compared to uncoupled. The most significant finding is that boundary layer ingestion, while offering substantial fuel burn savings, introduces throttle dependent aerodynamics effects that need to be accounted for. This work represents a first step toward the multidisciplinary design optimization of boundary layer ingestion propulsion systems.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN37844 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 66
    Publication Date: 2019-07-13
    Description: This paper originates from the joint efforts of an aeroelastic study team in the Applied Vehicle Technology Panel from NATO Science and Technology Organization, with the Task Group number AVT-191, titled "Application of Sensitivity Analysis and Uncertainty Quantification to Military Vehicle Design." We present aeroelastic uncertainty quantification studies using the SemiSpan Supersonic Transport wind tunnel model at the NASA Langley Research Center. The aeroelastic study team decided treat both structural and aerodynamic input parameters as uncertain and represent them as samples drawn from statistical distributions, propagating them through aeroelastic analysis frameworks. Uncertainty quantification processes require many function evaluations to asses the impact of variations in numerous parameters on the vehicle characteristics, rapidly increasing the computational time requirement relative to that required to assess a system deterministically. The increased computational time is particularly prohibitive if high-fidelity analyses are employed. As a remedy, the Istanbul Technical University team employed an Euler solver in an aeroelastic analysis framework, and implemented reduced order modeling with Polynomial Chaos Expansion and Proper Orthogonal Decomposition to perform the uncertainty propagation. The NASA team chose to reduce the prohibitive computational time by employing linear solution processes. The NASA team also focused on determining input sample distributions.
    Keywords: Aerodynamics
    Type: NF1676L-24684 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 67
    Publication Date: 2019-07-13
    Description: This presentation describes aerodynamics and performance X-plane airworthiness guidelines and best practices.
    Keywords: Aerodynamics
    Type: AFRC-E-DAA-TN43168 , AIAA Aviation 2017; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 68
    Publication Date: 2019-07-13
    Description: The topic of laminar flow effects on hover performance is introduced with respect to some historical efforts where laminar flow was either measured or attempted. An analysis method is outlined using combined blade element, momentum method coupled to an airfoil analysis method, which includes the full e(sup N) transition model. The analysis results compared well with the measured hover performance including the measured location of transition on both the upper and lower blade surfaces. The analysis method is then used to understand the upper limits of hover efficiency as a function of disk loading. The impact of laminar flow is higher at low disk loading, but significant improvement in terms of power loading appears possible even up to high disk loading approaching 20 ps f. A optimum planform design equation is derived for cases of zero profile drag and finite drag levels. These results are intended to be a guide for design studies and as a benchmark to compare higher fidelity analysis results. The details of the analysis method are given to enable other researchers to use the same approach for comparison to other approaches.
    Keywords: Aerodynamics
    Type: NF1676L-25565 , AHS International Annual Forum and Technology Display; May 09, 2017 - May 11, 2017; Fort Worth, TX; United States
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  • 69
    Publication Date: 2019-07-13
    Description: The current simulation for the Capsule Parachute Assembly System (CPAS) lacks fidelity in representing added mass for the 116 ft Do ringsail Main parachute. The availability of 3-D models of inflating Main canopies allowed for better estimation the enclosed air volume as a function of time. This was combined with trajectory state information to estimate the components making up measured axial loads. A proof-of-concept for an alternate simulation algorithm was developed based on enclosed volume as the primary independent variable rather than drag area growth. Databases of volume growth and parachute drag area vs. volume were developed for several flight tests. Other state information was read directly from test data, rather than numerically propagated. The resulting simulated peak loads were close in timing and magnitude to the measured loads data. However, results are very sensitive to data curve fitting and may not be suitable for Monte Carlo simulations. It was assumed that apparent mass was either negligible or a small fraction of enclosed mass, with little difference in results.
    Keywords: Aerodynamics
    Type: JSC-CN-38969 , AIAA Aerodynamic Decelerator Systems Conference; Jun 05, 2017 - Jun 09, 2017; Denver, CO; United States
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  • 70
    Publication Date: 2019-07-13
    Description: The Common Research Model wing-body configuration is investigated with the k-kL-MEAH2015 turbulence model implemented in FUN3D. This includes results presented at the Sixth Drag Prediction Workshop and additional results generated after the workshop with a nonlinear Quadratic Constitutive Relation (QCR) variant of the same turbulence model. The workshop provided grids are used, and a uniform grid refinement study is performed at the design condition. A large variation between results with and without a reconstruction limiter is exhibited on "medium" grid sizes, indicating that the medium grid size is too coarse for drawing conclusions in comparison with experiment. This variation is reduced with grid refinement. At a fixed angle of attack near design conditions, the QCR variant yielded decreased lift and drag compared with the linear eddy-viscosity model by an amount that was approximately constant with grid refinement. The k-kL-MEAH2015 turbulence model produced wing root junction flow behavior consistent with wind tunnel observations.
    Keywords: Aerodynamics
    Type: NF1676L-25300 , 2017 AIAA SciTech; Jan 09, 2017 - Jan 13, 2017; Dallas, TX; United States
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  • 71
    Publication Date: 2019-07-13
    Description: An experiment involving a Mach-scaled, 11:08 f t: diameter rotor was performed in hover during the summer of 2016 at NASA Langley Research Center. The experiment investigated the hover performance as a function of the laminar to turbulent transition state of the boundary layer, including both natural and fixed transition cases. The boundary layer transition locations were measured on both the upper and lower aerodynamic surfaces simultaneously. The measurements were enabled by recent advances in infrared sensor sensitivity and stability. The infrared thermography measurement technique was enhanced by a paintable blade surface heater, as well as a new high-sensitivity long wave infrared camera. The measured transition locations showed extensive amounts, x=c〉0:90, of laminar flow on the lower surface at moderate to high thrust (CT=s 〉 0:068) for the full blade radius. The upper surface showed large amounts, x=c 〉 0:50, of laminar flow at the blade tip for low thrust (CT=s 〈 0:045). The objective of this paper is to provide an experimental data set for comparisons to newly developed and implemented rotor boundary layer transition models in CFD and rotor design tools. The data is expected to be used as part of the AIAA Rotorcraft SimulationWorking Group
    Keywords: Aerodynamics
    Type: NF1676L-26006 , AIAA SciTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapvine, TX; United States
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  • 72
    Publication Date: 2019-07-13
    Description: A graphical framework is used for statistical analysis of the results from an extensive N- version test of a collection of Reynolds-averaged Navier-Stokes computational uid dynam- ics codes. The solutions were obtained by code developers and users from North America, Europe, Asia, and South America using both common and custom grid sequencees as well as multiple turbulence models for the June 2016 6th AIAA CFD Drag Prediction Workshop sponsored by the AIAA Applied Aerodynamics Technical Committee. The aerodynamic con guration for this workshop was the Common Research Model subsonic transport wing- body previously used for both the 4th and 5th Drag Prediction Workshops. This work continues the statistical analysis begun in the earlier workshops and compares the results from the grid convergence study of the most recent workshop with previous workshops.
    Keywords: Aerodynamics
    Type: NF1676L-24932 , 2017 AIAA SciTech; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 73
    Publication Date: 2019-07-13
    Description: In support of NASA's Commercial Supersonics Technology (CST) project, a test was conducted in the 9-by-7 ft. supersonic section of the NASA Ames Unitary Plan Wind Tunnel (UPWT). The tests were designed to study the interaction of shocks with a supersonic jet characteristic of those that may occur on a commercial supersonic aircraft. Multiple shock generating geometries were tested to examine the interaction dynamics as they pertain to sonic boom mitigation. An integral part of the analyses of these interactions are the interpretation of the data generated from the retroreflective Background Oriented Schlieren (RBOS) imaging technique employed for this test. The regularization- based optical flow methodology used to generate these data is described. Sample results are compared to those using normalized cross-correlation. The reduced noise, additional feature detail, and fewer false artifacts provided by the optical flow technique produced clearer time-averaged images, allowing for better interpretation of the underlying flow phenomena. These images, coupled with pressure signatures in the near field, are used to provide an overview of the detailed interaction flowfields.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN38216 , AIAA Sci-Tech; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 74
    Publication Date: 2019-07-13
    Description: Typically, a limited number of dynamic pressure sensors is employed to determine the unsteady aerodynamic forces on large, slender aerospace structures. This paper describes a robust calculation procedure based on frequency-specific correlation lengths and validation from an experiment conducted on a flat panel coated with fast response pressure-sensitive paint. The first part of the paper describes the procedure used to analyze the pressure sensitive paint images and a calibration method using dynamic pressure transducers. Excellent comparison in spectra, coherence, and phase, measured via pressure-sensitive paint and dynamic pressure sensors, validates the pressure-sensitive paint data. The second part of the paper describes the buffet validation process, the first step of which was to use pressure histories from all pixels to determine the true force fluctuations. In the next step, only a selected number of pixels was chosen as virtual sensors, and a correlation-length-based buffet calculation procedure was applied to determine modeled force fluctuations. By progressively decreasing the number of virtual sensors, it was observed that the present calculation procedure was able to make a close estimate of the true unsteady forces only from eight sensors. It is believed that the present work provides the first validation of a buffet calculation procedure.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN56170 , Journal of Aircraft (ISSN 0021-8669) (e-ISSN 1533-3868); 54; 5; 1791-1801
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  • 75
    Publication Date: 2019-07-13
    Description: Time-Resolved shadowgraph and infrared (IR) imaging were performed to investigate off-body and on-body flow features of a generic, 'hammer-head' launch vehicle geometry previously tested by Coe and Nute (1962). The measurements discussed here were one part of a large range of wind tunnel test techniques that included steady-state pressure sensitive paint (PSP), dynamic PSP, unsteady surface pressures, and unsteady force measurements. Image data was captured over a Mach number range of 0.6 less than or equal to M less than or equal to 1.2 at a Reynolds number of 3 million per foot. Both shadowgraph and IR imagery were captured in conjunction with unsteady pressures and forces and correlated with IRIG-B timing. High-speed shadowgraph imagery was used to identify wake structure and reattachment behind the payload fairing of the vehicle. Various data processing strategies were employed and ultimately these results correlated well with the location and magnitude of unsteady surface pressure measurements. Two research grade IR cameras were positioned to image boundary layer transition at the vehicle nose and flow reattachment behind the payload fairing. The poor emissivity of the model surface treatment (fast PSP) proved to be challenging for the infrared measurement. Reference image subtraction and contrast limited adaptive histogram equalization (CLAHE) were used to analyze this dataset. Ultimately turbulent boundary layer transition was observed and located forward of the trip dot line at the model sphere-cone junction. Flow reattachment location was identified behind the payload fairing in both steady and unsteady thermal data. As demonstrated in this effort, recent advances in high-speed and thermal imaging technology have modernized classical techniques providing a new viewpoint for the modern researcher
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN35004 , ScitTech 2017; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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  • 76
    Publication Date: 2019-07-13
    Description: The blade crossing event of a coaxial counter-rotating rotor is a potential source of noise and impulsive blade loads. Blade crossings occur many times during each rotor revolution. In previous research by the authors, this phenomenon was analyzed by simulating two airfoils passing each other at specified speeds and vertical separation distances, using the compressible Navier-Stokes solver OVERFLOW. The simulations explored mutual aerodynamic interactions associated with thickness, circulation, and compressibility effects. Results revealed the complex nature of the aerodynamic impulses generated by upperlower airfoil interactions. In this paper, the coaxial rotor system is simulated using two trains of airfoils, vertically offset, and traveling in opposite directions. The simulation represents multiple blade crossings in a rotor revolution by specifying horizontal distances between each airfoil in the train based on the circumferential distance between blade tips. The shed vorticity from prior crossing events will affect each pair of upperlower airfoils. The aerodynamic loads on the airfoil and flow field characteristics are computed before, at, and after each airfoil crossing. Results from the multiple-airfoil simulation show noticeable changes in the airfoil aerodynamics by introducing additional fluctuation in the aerodynamic time history.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN45472 , SAE International Journal of Aerospace; 10; 2; 68-76|SAE AeroTech Congress & Exhibition; Sep 26, 2017 - Sep 28, 2017; Fort Worth, TX; United States
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  • 77
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN38582 , AIAA Aerospace Sciences Meeting; Jan 09, 2017 - Jan 13, 2017; Grapevine, TX; United States
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