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  • 1
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    Understanding earthquakes: The key role of radar images (2013)
    Elsevier Science Limited
    Details
    Publication Date: 2017-04-03
    Description: The investigation of the fault rupture underlying earthquakes greatly improved thanks to the spread of radar images. Following pioneer applications in the eighties, Interferometry from Synthetic Aperture Radar (InSAR) gained a prominent role in geodesy. Its capability to measure millimetric deformations for wide areas and the increased data availability from the early nineties, made InSAR a diffused and accepted analysis tool in tectonics, though several factors contribute to reduce the data quality. With the introduction of analytical or numerical modeling, InSAR maps are used to infer the source of an earthquake by means of data inversion. Newly developed algorithms, known as InSAR time-series, allowed to further improve the data accuracy and completeness, strengthening the InSAR contribution even in the study of the inter- and post-seismic phase. In this work we describe the rationale at the base of the whole processing, showing its application to the New Zealand 2010-2011 seismic sequence.
    Description: Published
    Description: 178–181
    Description: 1.10. TTC - Telerilevamento
    Description: JCR Journal
    Description: restricted
    Keywords: InSAR ; Fault modeling ; Tectonics ; 04. Solid Earth::04.02. Exploration geophysics::04.02.07. Instruments and techniques
    Repository Name: Istituto Nazionale di Geofisica e Vulcanologia (INGV)
    Type: article
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  • 2
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    A ~125° post-early Serravallian counterclockwise rotation of the Gorgoglione Formation (Southern Apennines, Italy): New constraints for the formation of the Calabrian Arc (2013)
    Elsevier Science Limited
    Details
    Publication Date: 2017-04-04
    Description: The Southern Apennines, Calabro-Peloritane block, and Sicilian Maghrebides form a ~700 km long orogenic bend, known as Calabrian Arc (Cifelli et al., 2007). The bending of this orogenic system was realized progressively through opposite-sense rotation of the two limbs, counterclockwise (CCW) in the Southern Apennines and clockwise (CW) in the Sicilian Maghrebides, synchronous to the Miocene-to-Present opening of the Tyrrhenian Sea. Despite the wealth of paleomagnetic data from the Southern Apennines, the main Miocene rotational phase still remains poorly constrained in time and, more importantly, data from the most internal paleogeographic domains of the belt are completely lacking. The Gorgoglione Formation, a middle Miocene piggy-back deposit of the Southern Apennines, unconformably resting over the internal Sicilide Unit, offers the unique opportunity to document the deformation pattern of the most internal units, and reconstruct the incipient tectonic phases leading to the formation of the Calabrian Arc. New paleomagnetic and biostratigraphic data from the Gorgoglione Fm. reveal a post-early Serravallian ~125° CCW rotation with respect to stable Africa. Such a large rotation, affecting the Gorgoglione Fm. (and consequently the underneath allochthonous Sicilide nappe) exceeds by ~45° the maximum mean CCW rotation previously reported for the Southern Apennines. We propose that the additional ~45° CCW rotation measured in the Sicilide Unit is the result of an earlier, late Miocene phase of deformation related to the onset of the Tyrrhenian Sea opening and affecting the most internal paleogeographic domains of the Southern Apennines. Our reconstructed tectonic scenario confirms and emphasizes the central role of the Ionian slab in the geodynamic evolution of the central Mediterranean.
    Description: Published
    Description: 24-37
    Description: 2.2. Laboratorio di paleomagnetismo
    Description: JCR Journal
    Description: restricted
    Keywords: Southern Apennines ; Gorgoglione Formation ; Paleomagnetism ; Tectonics ; Calabrian Arc ; Biostratigraphy ; 04. Solid Earth::04.04. Geology::04.04.08. Sediments: dating, processes, transport ; 04. Solid Earth::04.05. Geomagnetism::04.05.06. Paleomagnetism
    Repository Name: Istituto Nazionale di Geofisica e Vulcanologia (INGV)
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  • 3
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    Kapitel Wald (2013)
    In:  ???
    Details
    Publication Date: 2013
    Description: Simulationen mithilfe des Models 4C zu möglichen Auswirkungen der Klimaänderungen des RCP 8.5 Klimaszenariums auf Wälder in Deutschland Kiefer Fichte Eiche Buche KATASTER-BESCHREIBUNG: Auswirkungen des Klimawandels (Temperatur, Niederschlag, CO2-Gehalt der Atmosphäre) auf die Wälder KATASTER-DETAIL: Delta T (Frühjahr) + und Delta Nied (Frühjahr) -, dann Produktivität der Wälder -; Delta C02 + um 25 - 30 %, dann Produktion der Wälder + um 9 - 20%; Delta T + (an nicht wasserlimitierten Standorten), dann Produktivität der Wälder +; Delta CO2+, dann Wassernutzungseffizienz der Wälder +; Delta T (Sommer) +, dann Waldbrandgefahr +; Delta T (Sommer) + und Delta Nied (Sommer) - (= WaBi -), dann Trockenstress der Wälder + um bis zu 9% und dann Produktivität der Wälder -; Delta T (Sommer) + und Delta Nied (Sommer) -, dann Populationsdichte Kiefern-Großschädlinge +;
    Keywords: Deutschland ; 20. und 21. Jahrhundert ; Boden ; Buche ; Eiche ; Fichte ; Forst ; Kiefer ; Klima ; Niederschlag ; Pflanzenschädling ; Phänologie ; Sturmschaden ; Temperatur ; Trockenheit ; Verdunstung ; Waldbrand ; Waldwachstum ; Wassermangel ; Wind ; Grundwasser ; Modell
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  • 4
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    Analysis of uncertainties in the hydrological response of a model-based climate change impact assessment in a subcatchment of the Spree River, Germany (2013)
    In:  Hydrological Processes, early view: 10.1002/hyp.9933.
    Details
    Publication Date: 2013
    Description: Ergebnisse aus Berechnungen mit den hydrologischen Modellen HBV-light und WaSiM-ETH angetrieben von den regionalen Klimamodellen (RCM) STAR (+2K, 100 Realisierungen, statistisch), WettReg2010 (A1B, 10 Realisierungen, statistisch), CCLM (A1B, 2 Realisierungen, dynamisch) und REMO (A1B, 1 Realisierung, dynamisch). Untersucht wurden zum einen die Unsicherheiten bezüglich des Einflusses der RCMs und hydrologischen Modelle auf die Endergebnisse sowie die zukünftige Abflussentwicklung in einem Teileinzugsgebiet der Oberen Spree KATASTER-BESCHREIBUNG: Bei der Betrachtung von langjährigen Abflussmittelwerten trägt die Wahl des RCMs die größte Unsicherheit zur Gesamtbandbreite des Modellensembles bei. Die Ergebnisse der hydrologischen Modelle unterscheiden sich nur geringfügig. Die Vulnerabilität des Einzugsgebiets bezüglich klimatischer Änderungen bleibt unbeantwortet, da zum Teil gegensätzlich Entwicklung der tatsächlichen Verdunstung und des Abflusses simuliert werden aufgrund der unterschiedlichen Niederschlagsentwicklungen der RCM KATASTER-DETAIL: Delta Temp: Alle RCMs (STAR (+2.4 °C), WettReg (+1.9 °C), CCLM (+1.6 °C), REMO (+1.2 °C)) Temperaturanstieg bis 2060 (dynamische RCMs schwächeren Anstieg im Vergleich zu den statistischen RCMs) Delta ETP: Alle RCMs Anstieg der potenziellen Verdunstung (ETP, Ansatz: Penman-Monteith). Statistische Modelle deutlich stärkeren Anstieg aufgrund geringerer Luftfeuchte und stärkeren Anstieg der Globalstrahlung (STAR: +135 mm/a, WettReg: +171 mm/a) im Vergleich zu den dynamischen RCMs (REMO: +12 mm/a, CCLM: +40 mm/a); Delta Nied: dynamische RCMs berechnen im Mittel leichten Anstieg (REMO: +38 mm/a, CCLM: +1 mm/a), statistische RCMs im Mittel erheblichen Rückgang (STAR: 120 mm/a, WettReg: 105 mm/a); Delta Tatsächliche Verdunstung: dynamische RCMs leichten Anstieg (REMO: zwischen +7-+18 mm/a, CCLM: zwischen +4-+10 mm/a, je nach hydrologischem Modell), statistische RCMs leichten Rückgang (STAR: zwischen -12 - -22 mm/a, WettReg: zwischen -2 - -10 mm/a, je nach hydrologischem Modell)
    Keywords: Lausitz, Nordostdeutschland ; 2031-2060 ; Klima ; Niederschlag ; Temperatur ; Verdunstung ; Globalstrahlung ; Sonnenscheindauer ; Niedrigwasser ; Abfluss ; Grundwasser ; Modell
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  • 5
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    Structural Dynamics Modeling of HIRENASD in Support of the Aeroelastic Prediction Workshop (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: An Aeroelastic Prediction Workshop (AePW) was held in April 2012 using three aeroelasticity case study wind tunnel tests for assessing the capabilities of various codes in making aeroelasticity predictions. One of these case studies was known as the HIRENASD model that was tested in the European Transonic Wind Tunnel (ETW). This paper summarizes the development of a standardized enhanced analytical HIRENASD structural model for use in the AePW effort. The modifications to the HIRENASD finite element model were validated by comparing modal frequencies, evaluating modal assurance criteria, comparing leading edge, trailing edge and twist of the wing with experiment and by performing steady and unsteady CFD analyses for one of the test conditions on the same grid, and identical processing of results.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-1801 , NF1676L-15290 , 54th AIAA/ASME/ASCE/AHS/ASC, Structures, Structural Dynamics, and Materials Conference; 8-1` Apr. 2013; Boston, MA; United States
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  • 6
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    DRE-Enhanced Swept-Wing Natural Laminar Flow at High Reynolds Numbers (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: Nonlinear parabolized stability equations and secondary instability analyses are used to provide a computational assessment of the potential use of the discrete roughness elements (DRE) technology for extending swept-wing natural laminar flow at chord Reynolds numbers relevant to transport aircraft. Computations performed for the boundary layer on a natural laminar flow airfoil with a leading-edge sweep angle of 34.6deg, free-stream Mach number of 0.75 and chord Reynolds numbers of 17 x 10(exp 6), 24 x 10(exp 6) and 30 x 10(exp 6) suggest that DRE could delay laminar-turbulent transition by about 20% when transition is caused by stationary crossflow disturbances. Computations show that the introduction of small wavelength stationary crossflow disturbances (i.e., DRE) also suppresses the growth of most amplified traveling crossflow disturbances.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0412 , NF1676L-14855 , 51st AIAA Aerospace Sciences Meeting; 7-110 Jan. 2013; Grapevine, TX; United States
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  • 7
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    Aerodynamic Classification of Swept-Wing Ice Accretion (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current stateof- the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice and spanwise-ridge ice. In the case of horn ice it is shown that a further subclassification of "nominally 3D" or "highly 3D" horn ice may be necessary. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-216556 , E-18753 , GRC-E-DAA-TN9727 , AIAA Atmospheric and Space Enviroments Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 8
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    Swept-Wing Ice Accretion Characterization and Aerodynamics (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA, FAA, ONERA, the University of Illinois and Boeing have embarked on a significant, collaborative research effort to address the technical challenges associated with icing on large-scale, three-dimensional swept wings. The overall goal is to improve the fidelity of experimental and computational simulation methods for swept-wing ice accretion formation and resulting aerodynamic effect. A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic experiments and computational simulations. As the baseline, full-scale, swept-wing-reference geometry, this research will utilize the 65% scale Common Research Model configuration. Ice-accretion testing will be conducted in the NASA Icing Research Tunnel for three hybrid swept-wing models representing the 20%, 64% and 83% semispan stations of the baseline-reference wing. Three-dimensional measurement techniques are being developed and validated to document the experimental ice-accretion geometries. Artificial ice shapes of varying geometric fidelity will be developed for aerodynamic testing over a large Reynolds number range in the ONERA F1 pressurized wind tunnel and in a smaller-scale atmospheric wind tunnel. Concurrent research will be conducted to explore and further develop the use of computational simulation tools for ice accretion and aerodynamics on swept wings. The combined results of this research effort will result in an improved understanding of the ice formation and aerodynamic effects on swept wings. The purpose of this paper is to describe this research effort in more detail and report on the current results and status to date. 1
    Keywords: Aerodynamics
    Type: NASA/TM-2013-216555 , AIAA Paper 2013-28 , E-18752 , GRC-E-DAA-TN9726 , AIAA Atmospheric and Space Enviroment Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 9
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    A Non-Cut Cell Immersed Boundary Method for Use in Icing Simulations (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes a computational fluid dynamic method used for modelling changes in aircraft geometry due to icing. While an aircraft undergoes icing, the accumulated ice results in a geometric alteration of the aerodynamic surfaces. In computational simulations for icing, it is necessary that the corresponding geometric change is taken into consideration. The method used, herein, for the representation of the geometric change due to icing is a non-cut cell Immersed Boundary Method (IBM). Computational cells that are in a body fitted grid of a clean aerodynamic geometry that are inside a predicted ice formation are identified. An IBM is then used to change these cells from being active computational cells to having properties of viscous solid bodies. This method has been implemented in the NASA developed node centered, finite volume computational fluid dynamics code, FUN3D. The presented capability is tested for two-dimensional airfoils including a clean airfoil, an iced airfoil, and an airfoil in harmonic pitching motion about its quarter chord. For these simulations velocity contours, pressure distributions, coefficients of lift, coefficients of drag, and coefficients of pitching moment about the airfoil's quarter chord are computed and used for comparison against experimental results, a higher order panel method code with viscous effects, XFOIL, and the results from FUN3D's original solution process. The results of the IBM simulations show that the accuracy of the IBM compares satisfactorily with the experimental results, XFOIL results, and the results from FUN3D's original solution process.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217881 , E-18692 , AIAA Paper 2012-1204 , AIAA Aerospace Sciences Meeting; Jan 09, 2012 - Jan 12, 2012; Nashville, TN; United States
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  • 10
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    Testing of the Trim Tab Parametric Model in NASA Langley's Unitary Plan Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In support of NASA's Entry, Descent, and Landing technology development efforts, testing of Langley's Trim Tab Parametric Models was conducted in Test Section 2 of NASA Langley's Unitary Plan Wind Tunnel. The objectives of these tests were to generate quantitative aerodynamic data and qualitative surface pressure data for experimental and computational validation and aerodynamic database development. Six component force-and-moment data were measured on 38 unique, blunt body trim tab configurations at Mach numbers of 2.5, 3.5, and 4.5, angles of attack from -4deg to +20deg, and angles of sideslip from 0deg to +8deg. Configuration parameters investigated in this study were forebody shape, tab area, tab cant angle, and tab aspect ratio. Pressure Sensitive Paint was used to provide qualitative surface pressure mapping for a subset of these flow and configuration variables. Over the range of parameters tested, the effects of varying tab area and tab cant angle were found to be much more significant than varying tab aspect ratio relative to key aerodynamic performance requirements. Qualitative surface pressure data supported the integrated aerodynamic data and provided information to aid in future analyses of localized phenomena for trim tab configurations.
    Keywords: Aerodynamics
    Type: NF1676L-15681 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 11
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    Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This is the presentation related to the paper of the same name describing Reynolds Averaged Navier Stokes (RANS) computational Fluid Dynamics (CFD) analysis of low speed stall aerodynamics of a swept wing with a laminar flow wing glove.
    Keywords: Aerodynamics
    Type: DFRC-E-DAA-TN9982 , DFRC-E-DAA-TN10986 , Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 12
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    Laminar, Transitional, and Turbulent Heating on Mid Lift-to-Drag Ratio Entry Vehicles (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The boundary-layer transition characteristics and convective aeroheating levels on mid lift-to-drag ratio entry vehicle configurations have been studied through wind-tunnel testing. Several configurations were investigated, including elliptically blunted cylinders with both circular and elliptically flattened cross sections, biconic geometries based on launch vehicle dual-use shrouds, and parametrically optimized analytic geometries. Vehicles of this class have been proposed for high-mass Mars missions, such as sample return and crewed exploration, for which the conventional sphere-cone entry-vehicle geometries of previous Mars missions are insufficient. Testing was conducted at Mach 6 over a range of Reynolds numbers sufficient to generate laminar, transitional, and turbulent flow. Transition onset locations, both straight-line and cross-flow, and heating rates were obtained through global phosphor thermography. Supporting computations were performed to obtain heating rates for comparison with the data. Laminar data and predictions agreed to well within the experimental uncertainty. Fully turbulent data and predictions also agreed well. However, in transitional flow regions, greater differences were observed.
    Keywords: Aerodynamics
    Type: NF1676L-22328 , Journal of Spacecraft and Rockets; 50; 5; 937-959|AIAA Fluid Dynamics Conference and Exhibit; Jun 25, 2012 - Jun 28, 2012; New Orleans, LA; United States
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  • 13
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    Enhancements to the FAST-MAC Circulation Control Model and Recent High-Reynolds Number Testing in the National Transonic Facility (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were 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 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.
    Keywords: Aerodynamics
    Type: NF1676L-15676 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 14
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    Long Penetration Mode Counterflowing Jets for Supersonic Slender Configurations - A Numerical Study (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A novel approach of using counterflowing jets positioned strategically on the aircraft and exploiting its long penetration mode (LPM) of interaction towards sonic-boom mitigation forms the motivation for this study. Given that most previous studies on the counterflowing LPM jet have all been on blunt bodies and at high supersonic or hypersonic flow conditions, exploring the feasibility to obtain a LPM jet issuing from a slender body against low supersonic freestream conditions is the main focus of this study. Computational fluid dynamics computations of axisymmetric models (cone-cylinder and quartic geometry), of relevance to NASA's High Speed project, are carried out using the space-time conservation element solution element viscous flow solver with unstructured meshes. A systematic parametric study is conducted to determine the optimum combination of counterflowing jet size, mass flow rate, and nozzle geometry for obtaining LPM jets. Details from these computations will be used to assess the potential of the LPM counterflowing supersonic jet as a means of active flow control for enabling supersonic flight over land and to establish the knowledge base for possible future implementation of such technologies.
    Keywords: Aerodynamics
    Type: NF1676L-15643 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 15
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    Toward Verification of USM3D Extensions for Mixed Element Grids (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The unstructured tetrahedral grid cell-centered finite volume flow solver USM3D has been recently extended to handle mixed element grids composed of hexahedral, prismatic, pyramidal, and tetrahedral cells. Presently, two turbulence models, namely, baseline Spalart-Allmaras (SA) and Menter Shear Stress Transport (SST), support mixed element grids. This paper provides an overview of the various numerical discretization options available in the newly enhanced USM3D. Using the SA model, the flow solver extensions are verified on three two-dimensional test cases available on the Turbulence Modeling Resource website at the NASA Langley Research Center. The test cases are zero pressure gradient flat plate, planar shear, and bump-inchannel. The effect of cell topologies on the flow solution is also investigated using the planar shear case. Finally, the assessment of various cell and face gradient options is performed on the zero pressure gradient flat plate case.
    Keywords: Aerodynamics
    Type: NF1676L-15586 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 16
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    Aeroacoustic Simulation of Nose Landing Gear on Adaptive Unstructured Grids With FUN3D (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Numerical simulations have been performed for a partially-dressed, cavity-closed nose landing gear configuration that was tested in NASA Langley s closed-wall Basic Aerodynamic Research Tunnel (BART) and in the University of Florida's open-jet acoustic facility known as the UFAFF. The unstructured-grid flow solver FUN3D, developed at NASA Langley Research center, is used to compute the unsteady flow field for this configuration. Starting with a coarse grid, a series of successively finer grids were generated using the adaptive gridding methodology available in the FUN3D code. A hybrid Reynolds-averaged Navier-Stokes/large eddy simulation (RANS/LES) turbulence model is used for these computations. Time-averaged and instantaneous solutions obtained on these grids are compared with the measured data. In general, the correlation with the experimental data improves with grid refinement. A similar trend is observed for sound pressure levels obtained by using these CFD solutions as input to a FfowcsWilliams-Hawkings noise propagation code to compute the farfield noise levels. In general, the numerical solutions obtained on adapted grids compare well with the hand-tuned enriched fine grid solutions and experimental data. In addition, the grid adaption strategy discussed here simplifies the grid generation process, and results in improved computational efficiency of CFD simulations.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-2071 , NF1676L-15503 , AIAA/CEAS Aeroacoustics Conference; May 27, 2013 - May 29, 2013; Berlin; Germany
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  • 17
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    Design and Execution of the Hypersonic Inflatable Aerodynamic Decelerator Large-Article Wind Tunnel Experiment (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The testing of 3- and 6-meter diameter Hypersonic Inflatable Aerodynamic Decelerator (HIAD) test articles was completed in the National Full-Scale Aerodynamics Complex 40 ft x 80 ft Wind Tunnel test section. Both models were stacked tori, constructed as 60 degree half-angle sphere cones. The 3-meter HIAD was tested in two configurations. The first 3-meter configuration utilized an instrumented flexible aerodynamic skin covering the inflatable aeroshell surface, while the second configuration employed a flight-like flexible thermal protection system. The 6-meter HIAD was tested in two structural configurations (with and without an aft-mounted stiffening torus near the shoulder), both utilizing an instrumented aerodynamic skin.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN9525 , 10th International Planetary Probe Workshop; Jun 17, 2013 - Jun 21, 2013; San Jose, CA; United States
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  • 18
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    Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft s swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First AIAA CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.
    Keywords: Aerodynamics
    Type: DFRC-E-DAA-TN9632 , 31st Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 19
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    FUN3D Airload Predictions for the Full-Scale UH-60A Airloads Rotor in a Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An unsteady Reynolds-Averaged Navier-Stokes solver for unstructured grids, FUN3D, is used to compute the rotor performance and airloads of the UH-60A Airloads Rotor in the National Full-Scale Aerodynamic Complex (NFAC) 40- by 80-foot Wind Tunnel. The flow solver is loosely coupled to a rotorcraft comprehensive code, CAMRAD-II, to account for trim and aeroelastic deflections. Computations are made for the 1-g level flight speed-sweep test conditions with the airloads rotor installed on the NFAC Large Rotor Test Apparatus (LRTA) and in the 40- by 80-ft wind tunnel to determine the influence of the test stand and wind-tunnel walls on the rotor performance and airloads. Detailed comparisons are made between the results of the CFD/CSD simulations and the wind tunnel measurements. The computed trends in solidity-weighted propulsive force and power coefficient match the experimental trends over the range of advance ratios and are comparable to previously published results. Rotor performance and sectional airloads show little sensitivity to the modeling of the wind-tunnel walls, which indicates that the rotor shaft-angle correction adequately compensates for the wall influence up to an advance ratio of 0.37. Sensitivity of the rotor performance and sectional airloads to the modeling of the rotor with the LRTA body/hub increases with advance ratio. The inclusion of the LRTA in the simulation slightly improves the comparison of rotor propulsive force between the computation and wind tunnel data but does not resolve the difference in the rotor power predictions at mu = 0.37. Despite a more precise knowledge of the rotor trim loads and flight condition, the level of comparison between the computed and measured sectional airloads/pressures at an advance ratio of 0.37 is comparable to the results previously published for the high-speed flight test condition.
    Keywords: Aerodynamics
    Type: NF1676L-15521 , 69th Annual American Helicopter Society Forum and Technology Display; May 21, 2013 - May 23, 2013; Phoenix, AZ; United States
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    Sweeping Jet Actuators - A New Design Tool for High Lift Generation (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Active Flow Control (AFC) experiments performed at the Caltech Lucas Wind Tunnel on a generic airplane vertical tail model proved the effectiveness of sweeping jets in improving the control authority of a rudder. The results indicated that a momentum coefficient (C(sub u)) of approximately 2% increased the side force in excess of 50% at the maximum conventional rudder deflection angle in the absence of yaw. However, sparsely distributed actuators providing a collective C(sub u) approx. = 0.1% were able to increase the side force in excess of 20%. This result is achieved by reducing the spanwise flow along the swept back rudder and its success is attributed to the large sweep back angle of the vertical tail. This current effort was sponsored by the NASA Environmentally Responsible Aviation (ERA) project.
    Keywords: Aerodynamics
    Type: NF1676L-16485 , Innovative Control Effectors for Military Vehicles (AVT-215); May 20, 2013 - May 22, 2013; Stockholm; Sweden
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    Wind Tunnel Measurements of the Wake of a Full-Scale UH-60A Rotor in Forward Flight (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A full-scale UH-60A rotor was tested in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel in May 2010. The test was designed to acquire a suite of measurements to validate state-of-the-art modeling tools. Measurements include blade airloads (from a single pressure-instrumented blade), blade structural loads (strain gages), rotor performance (rotor balance and torque measurements), blade deformation (stereo-photogrammetry), and rotor wake measurements (Particle Image Velocimetry (PIV) and Retro-reflective Backward Oriented Schlieren (RBOS)). During the test, PIV measurements of flow field velocities were acquired in a stationary cross-flow plane located on the advancing side of the rotor disk at approximately 90 deg rotor azimuth. At each test condition, blade position relative to the measurement plane was varied. The region of interest (ROI) was 4-ft high by 14-ft wide and covered the outer half of the blade radius. Although PIV measurements were acquired in only one plane, much information can be gleaned by studying the rotor wake trajectory in this plane, especially when such measurements are augmented by blade airloads and RBOS data. This paper will provide a comparison between PIV and RBOS measurements of tip vortex position and vortex filament orientation for multiple rotor test conditions. Blade displacement measurements over the complete rotor disk will also be presented documenting blade-to-blade differences in tip-path-plane and providing additional information for correlation with PIV and RBOS measurements of tip vortex location. In addition, PIV measurements of tip vortex core diameter and strength will be presented. Vortex strength will be compared with measurements of maximum bound circulation on the rotor blade determined from pressure distributions obtained from 235 pressure sensors distributed over 9 radial stations.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN6205 , 69th American Helicopter Society Annual Forum; May 21, 2013 - May 23, 2013; Phoenix, AZ
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    Measurements of Tip Vortices from a Full-Scale UH-60A Rotor by Retro- Reflective Background Oriented Schlieren and Stereo Photogrammetry (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Positions of vortices shed by a full-scale UH-60A rotor in forward flight were measured during a test in the National Full- Scale Aerodynamics Complex at NASA Ames Research Center. Vortices in a region near the tip of the advancing blade were visualized from two directions by Retro-Reflective Background-Oriented Schlieren (RBOS). Correspondence of points on the vortex in the RBOS images from both cameras was established using epipolar geometry. The object-space coordinates of the vortices were then calculated from the image-plane coordinates using stereo photogrammetry. One vortex from the tip of the blade that had most recently passed was visible in most of the data. The visibility of the vortices was greatest at high thrust and low advance ratios. At these favorable conditions, vortices from the most recent passages of all four blades were detected. The vortex positions were in good agreement with PIV data for a case where PIV measurements were also made. RBOS and photogrammetry provided measurements of the angle at which each vortex passed through the PIV plane.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN8772 , 69th American Helicopter Society Annual Forum; May 21, 2013 - May 23, 2013; Phoenix, AZ
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    The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Mission Applications Study (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The objective of the HIAD Mission Applications Study is to quantify the benefits of HIAD infusion to the concept of operations of high priority exploration missions. Results of the study will identify the range of mission concepts ideally suited to HIADs and provide mission-pull to associated technology development programs while further advancing operational concepts associated with HIAD technology. A summary of Year 1 modeling and analysis results is presented covering missions focusing on Earth and Mars-based applications. Recommended HIAD scales are presented for near term and future mission opportunities and the associated environments (heating and structural loads) are described.
    Keywords: Aerodynamics
    Type: NF1676L-16189 , 22nd AIAA Aerodynamic Decelerator Systems Technology Conference; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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    Effect of Boundary-Layer Bleed Hole Inclination Angle and Scaling on Flow Coefficient Behavior (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Phase II data results of the Fundamental Inlet Bleed Experiments study at NASA Glenn Research Center are presented which include flow coefficient behavior for 21 bleed hole configurations. The bleed configurations are all round holes with hole diameters ranging from 0.795 to 6.35 mm, hole inclination angles from 20deg to 90deg, and thickness-to-diameter ratios from 0.25 to 2.0. All configurations were tested at a unit Reynolds number of 2.46 10(exp 7)/m and at discrete local Mach numbers of 1.33, 1.62, 1.98, 2.46, and 2.92. Interactions between the design parameters of hole diameter, hole inclination angle, and thickness-to-diameter as well as the interactions between the flow parameters of pressure ratio and Mach number upon the flow coefficient are examined, and a preliminary statistical model is proposed. An existing correlation is also examined with respect to this data.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217843 , E-18617 , AIAA Paper 2013-0424 , 51st Aerospace Science Conference; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Flow Simulation of N3-X Hybrid Wing-Body Configuration (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: System studies show that a N3-X hybrid wing-body aircraft with a turboelectric distributed propulsion system using a mail-slot inlet/nozzle nacelle can meet the environmental and performance goals for N+3 generation transports (three generations beyond the current air transport technology level) set by NASA s Subsonic Fixed Wing Project. In this study, a Navier-Stokes flow simulation of N3-X on hybrid unstructured meshes was conducted, including the mail-slot propulsor. The geometry of the mail-slot propulsor was generated by a CAD (Computer-Aided Design)-free shape parameterization. A body force approach was used for a more realistic and efficient simulation of the turning and loss effects of the fan blades and the inlet-fan interactions. Flow simulation results of the N3-X demonstrates the validity of the present approach.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0221 , E-18576 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Reconstruction of Orion Engineering Development Unit (EDU) Parachute Inflation Loads (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The process of reconstructing inflation loads of Capsule Parachute Assembly System (CPAS) has been updated as the program transitioned to testing Engineering Development Unit (EDU) hardware. The equations used to reduce the test data have been re-derived based on the same physical assumptions made by simulations. Due to instrumentation challenges, individual parachute loads are determined from complementary accelerometer and load cell measurements. Cluster inflations are now simulated by modeling each parachute individually to better represent different inflation times and non-synchronous disreefing. The reconstruction procedure is tailored to either infinite mass or finite mass events based on measurable characteristics from the test data. Inflation parameters are determined from an automated optimization routine to reduce subjectivity. Infinite mass inflation parameters have been re-defined to avoid unrealistic interactions in Monte Carlo simulations. Sample cases demonstrate how best-fit inflation parameters are used to generate simulated drag areas and loads which favorably agree with test data.
    Keywords: Aerodynamics
    Type: JSC-CN-28034 , 22nd AIAA Aerodynamic Decelerator Systems Technology Conference; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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    An Extension of the Time-Spectral Method to Overset Solvers (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Relative motion in the Cartesian or overset framework causes certain spatial nodes to move in and out of the physical domain as they are dynamically blanked by moving solid bodies. This poses a problem for the conventional Time-Spectral approach, which expands the solution at every spatial node into a Fourier series spanning the period of motion. The proposed extension to the Time-Spectral method treats unblanked nodes in the conventional manner but expands the solution at dynamically blanked nodes in a basis of barycentric rational polynomials spanning partitions of contiguously defined temporal intervals. Rational polynomials avoid Runge's phenomenon on the equidistant time samples of these sub-periodic intervals. Fourier- and rational polynomial-based differentiation operators are used in tandem to provide a consistent hybrid Time-Spectral overset scheme capable of handling relative motion. The hybrid scheme is tested with a linear model problem and implemented within NASA's OVERFLOW Reynolds-averaged Navier- Stokes (RANS) solver. The hybrid Time-Spectral solver is then applied to inviscid and turbulent RANS cases of plunging and pitching airfoils and compared to time-accurate and experimental data. A limiter was applied in the turbulent case to avoid undershoots in the undamped turbulent eddy viscosity while maintaining accuracy. The hybrid scheme matches the performance of the conventional Time-Spectral method and converges to the time-accurate results with increased temporal resolution.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN7536 , AIAA Aerospace Sciences Meeting; Feb 07, 2013 - Feb 10, 2013; Grapevine, TX; United States
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    Summary of Data from the Fifth AIAA CFD Drag Prediction Workshop (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Results from the Fifth AIAA CFD Drag Prediction Workshop (DPW-V) are presented. As with past workshops, numerical calculations are performed using industry-relevant geometry, methodology, and test cases. This workshop focused on force/moment predictions for the NASA Common Research Model wing-body configuration, including a grid refinement study and an optional buffet study. The grid refinement study used a common grid sequence derived from a multiblock topology structured grid. Six levels of refinement were created resulting in grids ranging from 0.64x10(exp 6) to 138x10(exp 6) hexahedra - a much larger range than is typically seen. The grids were then transformed into structured overset and hexahedral, prismatic, tetrahedral, and hybrid unstructured formats all using the same basic cloud of points. This unique collection of grids was designed to isolate the effects of grid type and solution algorithm by using identical point distributions. This study showed reduced scatter and standard deviation from previous workshops. The second test case studied buffet onset at M=0.85 using the Medium grid (5.1x106 nodes) from the above described sequence. The prescribed alpha sweep used finely spaced intervals through the zone where wing separation was expected to begin. Some solutions exhibited a large side of body separation bubble that was not observed in the wind tunnel results. An optional third case used three sets of geometry, grids, and conditions from the Turbulence Model Resource website prepared by the Turbulence Model Benchmarking Working Group. These simple cases were intended to help identify potential differences in turbulence model implementation. Although a few outliers and issues affecting consistency were identified, the majority of participants produced consistent results.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0046 , NF1676L-15853 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Statistical Analysis of CFD Solutions From the Fifth AIAA Drag Prediction Workshop (2013)
    In:  CASI
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    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 fluid dynamics codes. The solutions were obtained by code developers and users from North America, Europe, Asia, and South America using a common grid sequence and multiple turbulence models for the June 2012 fifth Drag Prediction Workshop sponsored by the AIAA Applied Aerodynamics Technical Committee. The aerodynamic configuration for this workshop was the Common Research Model subsonic transport wing-body previously used for the 4th Drag Prediction Workshop. 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: AIAA Paper 2013-0047 , NF1676L-15131 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Overview of the Aeroelastic Prediction Workshop (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The Aeroelastic Prediction Workshop brought together an international community of computational fluid dynamicists as a step in defining the state of the art in computational aeroelasticity. This workshop's technical focus was prediction of unsteady pressure distributions resulting from forced motion, benchmarking the results first using unforced system data. The most challenging aspects of the physics were identified as capturing oscillatory shock behavior, dynamic shock-induced separated flow and tunnel wall boundary layer influences. The majority of the participants used unsteady Reynolds-averaged Navier Stokes codes. These codes were exercised at transonic Mach numbers for three configurations and comparisons were made with existing experimental data. Substantial variations were observed among the computational solutions as well as differences relative to the experimental data. Contributing issues to these differences include wall effects and wall modeling, non-standardized convergence criteria, inclusion of static aeroelastic deflection, methodology for oscillatory solutions, post-processing methods. Contributing issues pertaining principally to the experimental data sets include the position of the model relative to the tunnel wall, splitter plate size, wind tunnel expansion slot configuration, spacing and location of pressure instrumentation, and data processing methods.
    Keywords: Aerodynamics
    Type: NF1676L-15678 , AIAA Paper 2013-0783 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    CFL3D, FUN3d, and NSU3D Contributions to the Fifth Drag Prediction Workshop (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Results presented at the Fifth Drag Prediction Workshop using CFL3D, FUN3D, and NSU3D are described. These are calculations on the workshop provided grids and drag adapted grids. The NSU3D results have been updated to reflect an improvement to skin friction calculation on skewed grids. FUN3D results generated after the workshop are included for custom participant generated grids and a grid from a previous workshop. Uniform grid refinement at the design condition shows a tight grouping in calculated drag, where the variation in the pressure component of drag is larger than the skin friction component. At this design condition, A fine-grid drag value was predicted with a smaller drag adjoint adapted grid via tetrahedral adaption to a metric and mixed-element subdivision. The buffet study produced larger variation than the design case, which is attributed to large differences in the predicted side-of-body separation extent. Various modeling and discretization approaches had a strong impact on predicted side-of-body separation. This large wing root separation bubble was not observed in wind tunnel tests indicating that more work is necessary in modeling wing root juncture flows to predict experiments.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0050 , NF1676L-15475 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Analysis of Highly-Resolved Simulations of 2-D Humps Toward Improvement of Second-Moment Closures (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Fully resolved simulation data of flow separation over 2-D humps has been used to analyze the modeling terms in second-moment closures of the Reynolds-averaged Navier- Stokes equations. Existing models for the pressure-strain and dissipation terms have been analyzed using a priori calculations. All pressure-strain models are incorrect in the high-strain region near separation, although a better match is observed downstream, well into the separated-flow region. Near-wall inhomogeneity causes pressure-strain models to predict incorrect signs for the normal components close to the wall. In a posteriori computations, full Reynolds stress and explicit algebraic Reynolds stress models predict the separation point with varying degrees of success. However, as with one- and two-equation models, the separation bubble size is invariably over-predicted.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0684 , NF1676L-14868 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Assessments of a Turbulence Model Based on Menter's Modification to Rotta's Two-Equation Model (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The main objective of this paper is to construct a turbulence model with a more reliable second equation simulating length scale. In the present paper, we assess the length scale equation based on Menter s modification to Rotta s two-equation model. Rotta shows that a reliable second equation can be formed in an exact transport equation from the turbulent length scale L and kinetic energy. Rotta s equation is well suited for a term-by-term modeling and shows some interesting features compared to other approaches. The most important difference is that the formulation leads to a natural inclusion of higher order velocity derivatives into the source terms of the scale equation, which has the potential to enhance the capability of Reynolds-averaged Navier-Stokes (RANS) to simulate unsteady flows. The model is implemented in the PAB3D solver with complete formulation, usage methodology, and validation examples to demonstrate its capabilities. The detailed studies include grid convergence. Near-wall and shear flows cases are documented and compared with experimental and Large Eddy Simulation (LES) data. The results from this formulation are as good or better than the well-known SST turbulence model and much better than k-epsilon results. Overall, the study provides useful insights into the model capability in predicting attached and separated flows.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0341 , NF1676L-14859 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    FUN3D Analyses in Support of the First Aeroelastic Prediction Workshop (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: This paper presents the computational aeroelastic results generated in support of the first Aeroelastic Prediction Workshop for the Benchmark Supercritical Wing (BSCW) and the HIgh REynolds Number AeroStructural Dynamics (HIRENASD) configurations and compares them to the experimental data. The computational results are obtained using FUN3D, an unstructured grid Reynolds-averaged Navier-Stokes solver developed at NASA Langley Research Center. The analysis results for both configurations include aerodynamic coefficients and surface pressures obtained for steady-state or static aeroelastic equilibrium (BSCW and HIRENASD, respectively) and for unsteady flow due to a pitching wing (BSCW) or modally-excited wing (HIRENASD). Frequency response functions of the pressure coefficients with respect to displacement are computed and compared with the experimental data. For the BSCW, the shock location is computed aft of the experimentally-located shock position. The pressure distribution upstream of this shock is in excellent agreement with the experimental data, but the pressure downstream of the shock in the separated flow region does not match as well. For HIRENASD, very good agreement between the numerical results and the experimental data is observed at the mid-span wing locations.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0785 , NF1676L-14862 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Characterization of Aerodynamic Interactions with the Mars Science Laboratory Reaction Control System Using Computation and Experiment (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: On August 5 , 2012, the Mars Science Laboratory (MSL) entry capsule successfully entered Mars' atmosphere and landed the Curiosity rover in Gale Crater. The capsule used a reaction control system (RCS) consisting of four pairs of hydrazine thrusters to fly a guided entry. The RCS provided bank control to fly along a flight path commanded by an onboard computer and also damped unwanted rates due to atmospheric disturbances and any dynamic instabilities of the capsule. A preliminary assessment of the MSL's flight data from entry showed that the capsule flew much as predicted. This paper will describe how the MSL aerodynamics team used engineering analyses, computational codes and wind tunnel testing in concert to develop the RCS system and certify it for flight. Over the course of MSL's development, the RCS configuration underwent a number of design iterations to accommodate mechanical constraints, aeroheating concerns and excessive aero/RCS interactions. A brief overview of the MSL RCS configuration design evolution is provided. Then, a brief description is presented of how the computational predictions of RCS jet interactions were validated. The primary work to certify that the RCS interactions were acceptable for flight was centered on validating computational predictions at hypersonic speeds. A comparison of computational fluid dynamics (CFD) predictions to wind tunnel force and moment data gathered in the NASA Langley 31-Inch Mach 10 Tunnel was the lynch pin to validating the CFD codes used to predict aero/RCS interactions. Using the CFD predictions and experimental data, an interaction model was developed for Monte Carlo analyses using 6-degree-of-freedom trajectory simulation. The interaction model used in the flight simulation is presented.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0971 , NF1676L-14844 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Dynamic Modeling Accuracy Dependence on Errors in Sensor Measurements, Mass Properties, and Aircraft Geometry (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A nonlinear simulation of the NASA Generic Transport Model was used to investigate the effects of errors in sensor measurements, mass properties, and aircraft geometry on the accuracy of dynamic models identified from flight data. Measurements from a typical system identification maneuver were systematically and progressively deteriorated and then used to estimate stability and control derivatives within a Monte Carlo analysis. Based on the results, recommendations were provided for maximum allowable errors in sensor measurements, mass properties, and aircraft geometry to achieve desired levels of dynamic modeling accuracy. Results using other flight conditions, parameter estimation methods, and a full-scale F-16 nonlinear aircraft simulation were compared with these recommendations.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0949 , NF1676L-14841 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Trip-Induced Transition Measurements in a Hypersonic Boundary Layer Using Molecular Tagging Velocimetry (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Measurements of mean streamwise velocity, fluctuating streamwise velocity, and instantaneous streamwise velocity profiles in a hypersonic boundary layer were obtained over a 10-degree half-angle wedge model. A laser-induced fluorescence-based molecular tagging velocimetry technique was used to make the measurements. The nominal edge Mach number was 4.2. Velocity profiles were measured both in an untripped boundary layer and in the wake of a 4-mm diameter cylindrical tripping element centered 75.4 mm downstream of the sharp leading edge. Three different trip heights were investigated: k = 0.53 mm, k = 1.0 mm and k = 2.0 mm. The laminar boundary layer thickness at the position of the measurements was approximately 1 mm, though the exact thickness was dependent on Reynolds number and wall temperature. All of the measurements were made starting from a streamwise location approximately 18 mm downstream of the tripping element. This measurement region continued approximately 30 mm in the streamwise direction. Additionally, measurements were made at several spanwise locations. An analysis of flow features show how the magnitude, spatial location, and spatial growth of streamwise velocity instabilities are affected by parameters such as the ratio of trip height to boundary layer thickness and roughness Reynolds number. The fluctuating component of streamwise velocity measured along the centerline of the model increased from approximately 75 m/s with no trip to +/-225 m/s with a 0.53-mm trip, and to +/-240 m/s with a 1-mm trip, while holding the freestream Reynolds number constant. These measurements were performed in the 31-inch Mach 10 Air Tunnel at the NASA Langley Research Center.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0042 , NF1676L-14839 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub t) approx. =. 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Re(sub t) approx. = 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of omega(delta)/U(sub infinity) approx. = 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of "eddy Mach wave" radiation.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0532 , NF1676L-14821 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Development of Advanced High Lift Leading Edge Technology for Laminar Flow Wings (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes the Advanced High Lift Leading Edge (AHLLE) task performed by Northrop Grumman Systems Corporation, Aerospace Systems (NGAS) for the NASA Subsonic Fixed Wing project in an effort to develop enabling high-lift technology for laminar flow wings. Based on a known laminar cruise airfoil that incorporated an NGAS-developed integrated slot design, this effort involved using Computational Fluid Dynamics (CFD) analysis and quality function deployment (QFD) analysis on several leading edge concepts, and subsequently down-selected to two blown leading-edge concepts for testing. A 7-foot-span AHLLE airfoil model was designed and fabricated at NGAS and then tested at the NGAS 7 x 10 Low Speed Wind Tunnel in Hawthorne, CA. The model configurations tested included: baseline, deflected trailing edge, blown deflected trailing edge, blown leading edge, morphed leading edge, and blown/morphed leading edge. A successful demonstration of high lift leading edge technology was achieved, and the target goals for improved lift were exceeded by 30% with a maximum section lift coefficient (Cl) of 5.2. Maximum incremental section lift coefficients ( Cl) of 3.5 and 3.1 were achieved for a blown drooped (morphed) leading edge concept and a non-drooped leading edge blowing concept, respectively. The most effective AHLLE design yielded an estimated 94% lift improvement over the conventional high lift Krueger flap configurations while providing laminar flow capability on the cruise configuration.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0211 , NF1676L-15885 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Wake Instabilities Behind Discrete Roughness Elements in High Speed Boundary Layers (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Computations are performed to study the flow past an isolated, spanwise symmetric roughness element in zero pressure gradient boundary layers at Mach 3.5 and 5.9, with an emphasis on roughness heights of less than 55 percent of the local boundary layer thickness. The Mach 5.9 cases include flow conditions that are relevant to both ground facility experiments and high altitude flight ("cold wall" case). Regardless of the Mach number, the mean flow distortion due to the roughness element is characterized by long-lived streamwise streaks in the roughness wake, which can support instability modes that did not exist in the absence of the roughness element. The higher Mach number cases reveal a variety of instability mode shapes with velocity fluctuations concentrated in different localized regions of high base flow shear. The high shear regions vary from the top of a mushroom shaped structure characterizing the centerline streak to regions that are concentrated on the sides of the mushroom. Unlike the Mach 3.5 case with nearly same values of scaled roughness height k/delta and roughness height Reynolds number Re(sub kk), the odd wake modes in both Mach 5.9 cases are significantly more unstable than the even modes of instability. Additional computations for a Mach 3.5 boundary layer indicate that the presence of a roughness element can also enhance the amplification of first mode instabilities incident from upstream. Interactions between multiple roughness elements aligned along the flow direction are also explored.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0081 , NF1676L-14814 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Trajectory Reconstruction and Uncertainty Analysis Using Mars Science Laboratory Pre-Flight Scale Model Aeroballistic Testing (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: As part of the Mars Science Laboratory (MSL) trajectory reconstruction effort at NASA Langley Research Center, free-flight aeroballistic experiments of instrumented MSL scale models was conducted at Aberdeen Proving Ground in Maryland. The models carried an inertial measurement unit (IMU) and a flush air data system (FADS) similar to the MSL Entry Atmospheric Data System (MEADS) that provided data types similar to those from the MSL entry. Multiple sources of redundant data were available, including tracking radar and on-board magnetometers. These experimental data enabled the testing and validation of the various tools and methodologies that will be used for MSL trajectory reconstruction. The aerodynamic parameters Mach number, angle of attack, and sideslip angle were estimated using minimum variance with a priori to combine the pressure data and pre-flight computational fluid dynamics (CFD) data. Both linear and non-linear pressure model terms were also estimated for each pressure transducer as a measure of the errors introduced by CFD and transducer calibration. Parameter uncertainties were estimated using a "consider parameters" approach.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-1132 , NF1676L-14815 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Elastically Shaped Wing Optimization and Aircraft Concept for Improved Cruise Efficiency (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents the findings of a study conducted tn 2010 by the NASA Innovation Fund Award project entitled "Elastically Shaped Future Air Vehicle Concept". The study presents three themes in support of meeting national and global aviation challenges of reducing fuel burn for present and future aviation systems. The first theme addresses the drag reduction goal through innovative vehicle configurations via non-planar wing optimization. Two wing candidate concepts have been identified from the wing optimization: a drooped wing shape and an inflected wing shape. The drooped wing shape is a truly biologically inspired wing concept that mimics a seagull wing and could achieve about 5% to 6% drag reduction, which is aerodynamically significant. From a practical perspective, this concept would require new radical changes to the current aircraft development capabilities for new vehicles with futuristic-looking wings such as this concept. The inflected wing concepts could achieve between 3% to 4% drag reduction. While the drag reduction benefit may be less, the inflected-wing concept could have a near-term impact since this concept could be developed within the current aircraft development capabilities. The second theme addresses the drag reduction goal through a new concept of elastic wing shaping control. By aeroelastically tailoring the wing shape with active control to maintain optimal aerodynamics, a significant drag reduction benefit could be realized. A significant reduction in fuel burn for long-range cruise from elastic wing shaping control could be realized. To realize the potential of the elastic wing shaping control concept, the third theme emerges that addresses the drag reduction goal through a new aerodynamic control effector called a variable camber continuous trailing edge flap. Conventional aerodynamic control surfaces are discrete independent surfaces that cause geometric discontinuities at the trailing edge region. These discontinuities promote vorticities which result in drag rises as well as noise sources. The variable camber trailing edge flap concept could provide a substantial drag reduction benefit over a conventional discrete flap system. Aerodynamic simulations show a drag reduction of over 50% could be achieved with the flap concept over a conventional discrete flap system.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN7097 , AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Airloads Correlation of the UH-60A Rotor Inside the 40- by 80-Foot Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The presented research validates the capability of a loosely-coupled computational fluid dynamics (CFD) and comprehensive rotorcraft analysis (CRA) code to calculate the flowfield around a rotor and test stand mounted inside a wind tunnel. The CFD/CRA predictions for the full-scale UH-60A Airloads Rotor inside the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames Research Center are compared with the latest measured airloads and performance data. The studied conditions include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed up to and including stall. For the speed sweep, wind tunnel modeling becomes important at advance ratios greater than 0.37 and test stand modeling becomes increasingly important as the advance ratio increases. For the thrust sweep, both the wind tunnel and test stand modeling become important as the rotor approaches stall. Despite the beneficial effects of modeling the wind tunnel and test stand, the new models do not completely resolve the current airload discrepancies between prediction and experiment.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN7153 , ARC-E-DAA-TN7085 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    A Baseline Load Schedule for the Manual Calibration of a Force Balance (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A baseline load schedule for the manual calibration of a force balance was developed that takes current capabilities at the NASA Ames Balance Calibration Laboratory into account. The load schedule consists of 18 load series with a total of 194 data points. It was designed to satisfy six requirements: (i) positive and negative loadings should be applied for each load component; (ii) at least three loadings should be applied between 0 % and 100 % load capacity; (iii) normal and side force loadings should be applied at the forward gage location, the aft gage location, and the balance moment center; (iv) the balance should be used in UP and DOWN orientation to get axial force loadings; (v) the constant normal and side force approaches should be used to get the rolling moment loadings; (vi) rolling moment loadings should be obtained for 0, 90, 180, and 270 degrees balance orientation. Three different approaches are also reviewed that may be used to independently estimate the natural zeros of the balance. These three approaches provide gage output differences that may be used to estimate the weight of both the metric and non-metric part of the balance. Manual calibration data of NASA s MK29A balance and machine calibration data of NASA s MC60D balance are used to illustrate and evaluate different aspects of the proposed baseline load schedule design.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN6376 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Dallas, TX; United States
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    Hidden Connections between Regression Models of Strain-Gage Balance Calibration Data (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Hidden connections between regression models of wind tunnel strain-gage balance calibration data are investigated. These connections become visible whenever balance calibration data is supplied in its design format and both the Iterative and Non-Iterative Method are used to process the data. First, it is shown how the regression coefficients of the fitted balance loads of a force balance can be approximated by using the corresponding regression coefficients of the fitted strain-gage outputs. Then, data from the manual calibration of the Ames MK40 six-component force balance is chosen to illustrate how estimates of the regression coefficients of the fitted balance loads can be obtained from the regression coefficients of the fitted strain-gage outputs. The study illustrates that load predictions obtained by applying the Iterative or the Non-Iterative Method originate from two related regression solutions of the balance calibration data as long as balance loads are given in the design format of the balance, gage outputs behave highly linear, strict statistical quality metrics are used to assess regression models of the data, and regression model term combinations of the fitted loads and gage outputs can be obtained by a simple variable exchange.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN6184 , 51st AIAA Aerospace Sciences Meeting; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.8310(exp 5) to 0.8510(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-218070 , E-18746 , AIAA Paper 2012-3879 , Joint Propulsion Conference and Exhibit; Jul 29, 2012 - Aug 01, 2012; Atlanta, GA; United States
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    An Overview of Modifications Applied to a Turbulence Response Analysis Method for Flexible Aircraft Configurations (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A software program and associated methodology to study gust loading on aircraft exists for a classification of geometrically simplified flexible configurations. This program consists of a simple aircraft response model with two rigid and three flexible symmetric degrees of freedom and allows for the calculation of various airplane responses due to a discrete one-minus-cosine gust as well as continuous turbulence. Simplifications, assumptions, and opportunities for potential improvements pertaining to the existing software program are first identified, then a revised version of the original software tool is developed with improved methodology to include more complex geometries, additional excitation cases, and output data so as to provide a more useful and accurate tool for gust load analysis. Revisions are made in the categories of aircraft geometry, computation of aerodynamic forces and moments, and implementation of horizontal tail mode shapes. In order to improve the original software program to enhance usefulness, a wing control surface and horizontal tail control surface is added, an extended application of the discrete one-minus-cosine gust input is employed, a supplemental continuous turbulence spectrum is implemented, and a capability to animate the total vehicle deformation response to gust inputs in included. These revisions and enhancements are implemented and an analysis of the results is used to validate the modifications.
    Keywords: Aerodynamics
    Type: NF1676L-16284 , AIAA Region I-MA Student Conference; Apr 05, 2013 - Apr 06, 2013; College Park, MD; United States
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    Hybrid Large-Eddy/Reynolds-Averaged Simulation of a Supersonic Cavity Using VULCAN (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Simulations of a supersonic recessed-cavity flow are performed using a hybrid large-eddy/Reynolds-averaged simulation approach utilizing an inflow turbulence recycling procedure and hybridized inviscid flux scheme. Calorically perfect air enters a three-dimensional domain at a free stream Mach number of 2.92. Simulations are performed to assess grid sensitivity of the solution, efficacy of the turbulence recycling, and the effect of the shock sensor used with the hybridized inviscid flux scheme. Analysis of the turbulent boundary layer upstream of the rearward-facing step for each case indicates excellent agreement with theoretical predictions. Mean velocity and pressure results are compared to Reynolds-averaged simulations and experimental data for each case and indicate good agreement on the finest grid. Simulations are repeated on a coarsened grid, and results indicate strong grid density sensitivity. Simulations are performed with and without inflow turbulence recycling on the coarse grid to isolate the effect of the recycling procedure, which is demonstrably critical to capturing the relevant shear layer dynamics. Shock sensor formulations of Ducros and Larsson are found to predict mean flow statistics equally well.
    Keywords: Aerodynamics
    Type: NF1676L-15505 , 2013 IEEE Aerospace Conference; Mar 02, 2013 - Mar 09, 2013; Big Sky, MT; United States
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    Recalibration of the Shear Stress Transport Model to Improve Calculation of Shock Separated Flows (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Menter Shear Stress Transport (SST) k . turbulence model is one of the most widely used two-equation Reynolds-averaged Navier-Stokes turbulence models for aerodynamic analyses. The model extends Menter s baseline (BSL) model to include a limiter that prevents the calculated turbulent shear stress from exceeding a prescribed fraction of the turbulent kinetic energy via a proportionality constant, a1, set to 0.31. Compared to other turbulence models, the SST model yields superior predictions of mild adverse pressure gradient flows including those with small separations. In shock - boundary layer interaction regions, the SST model produces separations that are too large while the BSL model is on the other extreme, predicting separations that are too small. In this paper, changing a1 to a value near 0.355 is shown to significantly improve predictions of shock separated flows. Several cases are examined computationally and experimental data is also considered to justify raising the value of a1 used for shock separated flows.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217851 , AIAA Paper 2013-0685 , E-18639 , 51st AIAA Aerospace Science Meeting and Exhibit; Jan 07, 2013 - Jan 10, 2013; Dallas, TX; United States
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    Application of Statistically Derived CPAS Parachute Parameters (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Capsule Parachute Assembly System (CPAS) Analysis Team is responsible for determining parachute inflation parameters and dispersions that are ultimately used in verifying system requirements. A model memo is internally released semi-annually documenting parachute inflation and other key parameters reconstructed from flight test data. Dispersion probability distributions published in previous versions of the model memo were uniform because insufficient data were available for determination of statistical based distributions. Uniform distributions do not accurately represent the expected distributions since extreme parameter values are just as likely to occur as the nominal value. CPAS has taken incremental steps to move away from uniform distributions. Model Memo version 9 (MMv9) made the first use of non-uniform dispersions, but only for the reefing cutter timing, for which a large number of sample was available. In order to maximize the utility of the available flight test data, clusters of parachutes were reconstructed individually starting with Model Memo version 10. This allowed for statistical assessment for steady-state drag area (CDS) and parachute inflation parameters such as the canopy fill distance (n), profile shape exponent (expopen), over-inflation factor (C(sub k)), and ramp-down time (t(sub k)) distributions. Built-in MATLAB distributions were applied to the histograms, and parameters such as scale (sigma) and location (mu) were output. Engineering judgment was used to determine the "best fit" distribution based on the test data. Results include normal, log normal, and uniform (where available data remains insufficient) fits of nominal and failure (loss of parachute and skipped stage) cases for all CPAS parachutes. This paper discusses the uniform methodology that was previously used, the process and result of the statistical assessment, how the dispersions were incorporated into Monte Carlo analyses, and the application of the distributions in trajectory benchmark testing assessments with parachute inflation parameters, drag area, and reefing cutter timing used by CPAS.
    Keywords: Aerodynamics
    Type: JSC-CN-28160 , AIAA Aerodynamic Deccelerator System Technology; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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    Exhaust Nozzle Plume and Shock Wave Interaction (2013)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Fundamental research for sonic boom reduction is needed to quantify the interaction of shock waves generated from the aircraft wing or tail surfaces with the exhaust plume. Both the nozzle exhaust plume shape and the tail shock shape may be affected by an interaction that may alter the vehicle sonic boom signature. The plume and shock interaction was studied using Computational Fluid Dynamics simulation on two types of convergent-divergent nozzles and a simple wedge shock generator. The nozzle plume effects on the lower wedge compression region are evaluated for two- and three-dimensional nozzle plumes. Results show that the compression from the wedge deflects the nozzle plume and shocks form on the deflected lower plume boundary. The sonic boom pressure signature of the wedge is modified by the presence of the plume, and the computational predictions show significant (8 to 15 percent) changes in shock amplitude.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217838 , AIAA Paper 2013-0012 , 51st Aerospace Science Conference; Jan 07, 2013 - Jan 10, 2013; Grapevine, TX; United States
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    Damping Effects of Drogue Parachutes on Orion Crew Module Dynamics (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Currently, simulation predictions of the Orion Crew Module (CM) dynamics with drogue parachutes deployed are under-predicting the amount of damping as seen in free-flight tests. The Apollo Legacy Chute Damping model has been resurrected and applied to the Orion system. The legacy model has been applied to predict CM damping under drogue parachutes for both Vertical Spin Tunnel free flights and the Pad Abort-1 flight test. Comparisons between the legacy Apollo prediction method and test data are favorable. A key hypothesis in the Apollo legacy drogue damping analysis is that the drogue parachutes' net load vector aligns with the CM drogue attachment point velocity vector. This assumption seems reasonable and produces good results, but has never been quantitatively proven. The wake of the CM influences the drogue parachutes, which makes performance predictions of the parachutes difficult. Many of these effects are not currently modeled in the simulations. A forced oscillation test of the CM with parachutes was conducted in the NASA LaRC 20-Ft Vertical Spin Tunnel (VST) to gather additional data to validate and refine the Apollo legacy drogue model. A second loads balance was added to the original Orion VST model to measure the drogue parachute loads independently of the CM. The objective of the test was to identify the contribution of the drogues to CM damping and provide additional information to quantify wake effects and the interactions between the CM and parachutes. The drogue parachute force vector was shown to be highly dependent on the CM wake characteristics. Based on these wind tunnel test data, the Apollo Legacy Chute Damping model was determined to be a sufficient approximation of the parachute dynamics in relationship to the CM dynamics for preliminary entry vehicle system design. More wake effects should be included to better model the system. These results are being used to improve simulation model fidelity of CM flight with drogues deployed, which has been identified by the project as key to a successful Orion Critical Design Review.
    Keywords: Aerodynamics
    Type: NF1676L-16912
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    Aerodynamic Classification of Swept-Wing Ice Accretion (2013)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current state-of-the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice, and spanwise-ridge ice. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN5969 , NASA/TM-2013-216381 , DOT/FAA/TC-13/21 , E-18556
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    High-Order Entropy Stable Finite Difference Schemes for Nonlinear Conservation Laws: Finite Domains (2013)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: Developing stable and robust high-order finite difference schemes requires mathematical formalism and appropriate methods of analysis. In this work, nonlinear entropy stability is used to derive provably stable high-order finite difference methods with formal boundary closures for conservation laws. Particular emphasis is placed on the entropy stability of the compressible Navier-Stokes equations. A newly derived entropy stable weighted essentially non-oscillatory finite difference method is used to simulate problems with shocks and a conservative, entropy stable, narrow-stencil finite difference approach is used to approximate viscous terms.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217971 , L-20223 , NF16767L-15999
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    Wind-US Code Contributions to the First AIAA Shock Boundary Layer Interaction Prediction Workshop (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: This report discusses the computations of a set of shock wave/turbulent boundary layer interaction (SWTBLI) test cases using the Wind-US code, as part of the 2010 American Institute of Aeronautics and Astronautics (AIAA) shock/boundary layer interaction workshop. The experiments involve supersonic flows in wind tunnels with a shock generator that directs an oblique shock wave toward the boundary layer along one of the walls of the wind tunnel. The Wind-US calculations utilized structured grid computations performed in Reynolds-averaged Navier-Stokes mode. Four turbulence models were investigated: the Spalart-Allmaras one-equation model, the Menter Baseline and Shear Stress Transport k-omega two-equation models, and an explicit algebraic stress k-omega formulation. Effects of grid resolution and upwinding scheme were also considered. The results from the CFD calculations are compared to particle image velocimetry (PIV) data from the experiments. As expected, turbulence model effects dominated the accuracy of the solutions with upwinding scheme selection indicating minimal effects.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217837 , E-18611
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    Analysis of the Effects of Streamwise Lift Distribution on Sonic Boom Signature (2013)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: Investigation of sonic boom has been one of the major areas of study in aeronautics due to the benefits a low-boom aircraft has in both civilian and military applications. This work conducts a numerical analysis of the effects of streamwise lift distribution on the shock coalescence characteristics. A simple wing-canard-stabilator body model is used in the numerical simulation. The streamwise lift distribution is varied by fixing the canard at a deflection angle while trimming the aircraft with the wing and the stabilator at the desired lift coefficient. The lift and the pitching moment coefficients are computed using the Missile DATCOM v. 707. The flow field around the wing-canard- stabilator body model is resolved using the OVERFLOW-2 flow solver. Overset/ chimera grid topology is used to simplify the grid generation of various configurations representing different streamwise lift distributions. The numerical simulations are performed without viscosity unless it is required for numerical stability. All configurations are simulated at Mach 1.4, angle-of-attack of 1.50, lift coefficient of 0.05, and pitching moment coefficient of approximately 0. Four streamwise lift distribution configurations were tested.
    Keywords: Aerodynamics
    Type: DRC-009-025 , NASA Tech Briefs, January 2013; 27
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    Dry wind tunnel system (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: This invention is a ground flutter testing system without a wind tunnel, called Dry Wind Tunnel (DWT) System. The DWT system consists of a Ground Vibration Test (GVT) hardware system, a multiple input multiple output (MIMO) force controller software, and a real-time unsteady aerodynamic force generation software, that is developed from an aerodynamic reduced order model (ROM). The ground flutter test using the DWT System operates on a real structural model, therefore no scaled-down structural model, which is required by the conventional wind tunnel flutter test, is involved. Furthermore, the impact of the structural nonlinearities on the aeroelastic stability can be included automatically. Moreover, the aeroservoelastic characteristics of the aircraft can be easily measured by simply including the flight control system in-the-loop. In addition, the unsteady aerodynamics generated computationally is interference-free from the wind tunnel walls. Finally, the DWT System can be conveniently and inexpensively carried out as a post GVT test with the same hardware, only with some possible rearrangement of the shakers and the inclusion of additional sensors.
    Keywords: Aerodynamics
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    Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail Interactions on a Slender Wing at Subsonic and Transonic Speeds (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed. 15.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217982 , L-19987 , NF1676L-12153
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    Experimental Investigation of Aeroelastic Deformation of Slender Wings at Supersonic Speeds Using a Video Model Deformation Measurement Technique (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A video-based photogrammetric model deformation system was established as a dedicated optical measurement technique at supersonic speeds in the NASA Langley Research Center Unitary Plan Wind Tunnel. This system was used to measure the wing twist due to aerodynamic loads of two supersonic commercial transport airplane models with identical outer mold lines but different aeroelastic properties. One model featured wings with deflectable leading- and trailing-edge flaps and internal channels to accommodate static pressure tube instrumentation. The wings of the second model were of single-piece construction without flaps or internal channels. The testing was performed at Mach numbers from 1.6 to 2.7, unit Reynolds numbers of 1.0 million to 5.0 million, and angles of attack from -4 degrees to +10 degrees. The video model deformation system quantified the wing aeroelastic response to changes in the Mach number, Reynolds number concurrent with dynamic pressure, and angle of attack and effectively captured the differences in the wing twist characteristics between the two test articles.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217977 , L-20228 , NF1676L-16083
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    Low noise wing slat system with rigid cove-filled slat (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Concepts and technologies described herein provide for a low noise aircraft wing slat system. According to one aspect of the disclosure provided herein, a cove-filled wing slat is used in conjunction with a moveable panel rotatably attached to the wing slat to provide a high lift system. The moveable panel rotates upward against the rear surface of the slat during deployment of the slat, and rotates downward to bridge a gap width between the stowed slat and the lower wing surface, completing the continuous outer mold line shape of the wing, when the cove-filled slat is retracted to the stowed position.
    Keywords: Aerodynamics
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    Wake Measurement Probe (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A probe includes a stem having a tip that measures a wake produced by an object moving through a fluid. The probe includes temperature and pressure sensors co-located in the tip.
    Keywords: Aerodynamics
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    Experimental Measurements of Sonic Boom Signatures Using a Continuous Data Acquisition Technique (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A wind tunnel investigation was conducted in the Langley Unitary Plan Wind Tunnel to determine the effectiveness of a technique to measure aircraft sonic boom signatures using a single conical survey probe while continuously moving the model past the probe. Sonic boom signatures were obtained using both move-pause and continuous data acquisition methods for comparison. The test was conducted using a generic business jet model at a constant angle of attack and a single model-to-survey-probe separation distance. The sonic boom signatures were obtained at a Mach number of 2.0 and a unit Reynolds number of 2 million per foot. The test results showed that it is possible to obtain sonic boom signatures while continuously moving the model and that the time required to acquire the signature is at least 10 times faster than the move-pause method. Data plots are presented with a discussion of the results. No tabulated data or flow visualization photographs are included.
    Keywords: Aerodynamics
    Type: NASA/TP-2013-218035 , L-20226 , NF1676L-16049
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    Effectiveness of a Wedge Probe to Measure Sonic Boom Signatures in a Supersonic Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A wind tunnel investigation was conducted in the Langley Unitary Plan Wind Tunnel (UPWT) to determine the effectiveness of a wedge probe to measure sonic boom pressure signatures compared to a slender conical probe. A generic business jet model at a constant angle of attack and at a single model to probe separation distance was used to generate a sonic boom signature. Pressure signature data were acquired with both the wedge probe and a slender conical probe for comparison. The test was conducted at a Mach number of 2.0 and a free-stream unit Reynolds number of 2 million per foot. The results showed that the wedge probe was not effective in measuring the sonic boom pressure signature of the aircraft model in the supersonic wind tunnel. Data plots and a discussion of the results are presented. No tabulated data or flow visualization photographs are included.
    Keywords: Aerodynamics
    Type: NASA/TP-2013-218036 , L-20239 , NF1676L-16253
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    Analysis of a Split-Plot Experimental Design Applied to a Low-Speed Wind Tunnel Investigation (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A procedure to analyze a split-plot experimental design featuring two input factors, two levels of randomization, and two error structures in a low-speed wind tunnel investigation of a small-scale model of a fighter airplane configuration is described in this report. Standard commercially-available statistical software was used to analyze the test results obtained in a randomization-restricted environment often encountered in wind tunnel testing. The input factors were differential horizontal stabilizer incidence and the angle of attack. The response variables were the aerodynamic coefficients of lift, drag, and pitching moment. Using split-plot terminology, the whole plot, or difficult-to-change, factor was the differential horizontal stabilizer incidence, and the subplot, or easy-to-change, factor was the angle of attack. The whole plot and subplot factors were both tested at three levels. Degrees of freedom for the whole plot error were provided by replication in the form of three blocks, or replicates, which were intended to simulate three consecutive days of wind tunnel facility operation. The analysis was conducted in three stages, which yielded the estimated mean squares, multiple regression function coefficients, and corresponding tests of significance for all individual terms at the whole plot and subplot levels for the three aerodynamic response variables. The estimated regression functions included main effects and two-factor interaction for the lift coefficient, main effects, two-factor interaction, and quadratic effects for the drag coefficient, and only main effects for the pitching moment coefficient.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-218013 , NF1676L-12426 , L-20005
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    Display Provides Pilots with Real-Time Sonic-Boom Information (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Supersonic aircraft generate shock waves that move outward and extend to the ground. As a cone of pressurized air spreads across the landscape along the flight path, it creates a continuous sonic boom along the flight track. Several factors can influence sonic booms: weight, size, and shape of the aircraft; its altitude and flight path; and weather and atmospheric conditions. This technology allows pilots to control the impact of sonic booms. A software system displays the location and intensity of shock waves caused by supersonic aircraft. This technology can be integrated into cockpits or flight control rooms to help pilots minimize sonic boom impact in populated areas. The system processes vehicle and flight parameters as well as data regarding current atmospheric conditions. The display provides real-time information regarding sonic boom location and intensity, enabling pilots to make the necessary flight adjustments to control the timing and location of sonic booms. This technology can be used on current-generation supersonic aircraft, which generate loud sonic booms, as well as future- generation, low-boom aircraft, anticipated to be quiet enough for populated areas.
    Keywords: Aerodynamics
    Type: DRC-008-001 , NASA Tech Briefs, March 2013; 6
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    Entropy Stable Spectral Collocation Schemes for the Navier-Stokes Equations: Discontinuous Interfaces (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Nonlinear entropy stability and a summation-by-parts framework are used to derive provably stable, polynomial-based spectral collocation methods of arbitrary order. The new methods are closely related to discontinuous Galerkin spectral collocation methods commonly known as DGFEM, but exhibit a more general entropy stability property. Although the new schemes are applicable to a broad class of linear and nonlinear conservation laws, emphasis herein is placed on the entropy stability of the compressible Navier-Stokes equations.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-218039 , L-20317 , NF1676L-17321
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    Stability, Transient Response, Control, and Safety of a High-Power Electric Grid for Turboelectric Propulsion of Aircraft (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: This document contains the deliverables for the NASA Research and Technology for Aerospace Propulsion Systems (RTAPS) regarding the stability, transient response, control, and safety study for a high power cryogenic turboelectric distributed propulsion (TeDP) system. The objective of this research effort is to enumerate, characterize, and evaluate the critical issues facing the development of the N3-X concept aircraft. This includes the proposal of electrical grid architecture concepts and an evaluation of any needs for energy storage.
    Keywords: Aerodynamics
    Type: NASA/CR-2013-217865 , E-18658
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    Recalibration of the Shear Stress Transport Model to Improve Calculation of Shock Separated Flows (2013)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: No abstract available
    Keywords: Aerodynamics
    Type: E-664421 , Annual Shock Wave/Boundary Layer Interaction Workshop (SWBLI); Apr 24, 2013 - Apr 25, 2013; Dayton, OH; United States
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    Effect of Hybrid Flow Control on a Normal Shock Boundary-Layer Interaction (2013)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: No abstract available
    Keywords: Aerodynamics
    Type: E-664420 , Shock Wave/Boundary-Layer Interraction Workshop (SWBLI); Apr 24, 2013 - Apr 25, 2013; Dayton, OH; United States
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    Silent and Efficient Supersonic Bi-Directional Flying Wing (2013)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: The supersonic bi-directional (SBiDir) flying wing (FW) concept has a great potential to achieve low sonic boom with high supersonic aerodynamic performance due to removal of performance conflict between high speed and low speed by rotating goo in flight. This NIAC Phase 1 research has achieved three objectives: 2) prove the concept based on simulation that it can achieve very low boom with smooth Sine wave ground over-pressure signature and excellent aerodynamic efficiency; 3) conduct trade study to correlate the geometric parameters with sonic boom and aerodynamic performance for further automated design optimization in Phase II. The design methodology developed in Phase I includes three parts: 1) an advanced geometry model, which can vary airfoil meanline angle distribution to control the expansion and shock waves on the airplane surface to mitigate sonic boom and improve aerodynamic efficiency. 2) a validated CFD procedure to resolve near field flow with accurate shock strength. The sonic boom propagation from near field to far field ground is simulated by NASA NF Boom code. The surface friction drag prediction is based on fiat plate correlation adopted by Seebass and supported by the experimental study of Winter and Smith, which is on the conservative side and is more reliable than CFD RANS simulation. 3) a mission analysis tool based on Corke's model that provides design requirements and constraints of supersonic airplanes for range, payload, volume, size, weight, etc. The design mission target is a supersonic transport with cruise Mach number 1.6, 100 passengers, and 4000nm range. The trade study has several very important findings: 1) The far field ground sonic boom signature is directly related to the smoothness of the flow on the airplane surface. The meanline angle distribution is a very effective control methodology to mitigate surface shock and expansion wave strength, and mitigating compression wave coalescing by achieving smooth loading distribution chord-wise. Compared with a linear meanline angle distribution, a design using nonlinear and non-monotonic meanline angle distribution is able to reduce the sonic boom ground loudness by over 20dBP1. The design achieves sonic boom ground loudness less than 70dBP1 and aerodynamic dynamic efficiency 1/D of 8.4. 2) Decreasing sweep angle within the Mach cone will increase 1/D as well as sonic boom. A design with variable sweep from 84 at the very leading edge to 68 at the tip achieves an extraordinarily high 1/D of 10.4 at Mach number 1.6 due to the low wave drag. If no sonic boom constraint is attached, SBiDir-FW concept still has a lot of room to increase the 1/D. 3) The round leading edge and trailing edge under high sweep angle are beneficial to improve aerodynamic performance, sonic boom, and to increase volume of the airplane. 4) Subsonic performance is benefited greatly from the high slenderness of supersonic configuration after rotating goo. A design with excellent supersonic aspect ratio of 0.44, 1/D of 8.g, gives an extraordinary subsonic aspect ration of 10 and 1/D of 1g.7. Two configurations are designed in details to install internal seats, landing gears, and engine installation to demonstrate the feasibility of SBiDir-FW configuration to accommodate all the required volume for realistic airplane. Here we emphasize that the qualitative findings in Phase I are very encouraging, more important than the quantitative results. Qualitative findings give the understanding of physics and provide the path to achieve the ultimate high performance design. The promising quantitative results achieved in Phase I need to be confirmed by wind tunnel testing in Phase II and ultimately proved by flight test. The other important step forward will be made to study the rotation transition from both CFD unsteady simulation and wind tunnel testing.
    Keywords: Aerodynamics
    Type: HQ-E-DAA-TN63203
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    High-Order Entropy Stable Formulations for Computational Fluid Dynamics (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A systematic approach is presented for developing entropy stable (SS) formulations of any order for the Navier-Stokes equations. These SS formulations discretely conserve mass, momentum, energy and satisfy a mathematical entropy inequality. They are valid for smooth as well as discontinuous flows provided sufficient dissipation is added at shocks and discontinuities. Entropy stable formulations exist for all diagonal norm, summation-by-parts (SBP) operators, including all centered finite-difference operators, Legendre collocation finite-element operators, and certain finite-volume operators. Examples are presented using various entropy stable formulations that demonstrate the current state-of-the-art of these schemes.
    Keywords: Aerodynamics
    Type: NF1676L-15651 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Direct Numerical Simulation of Transition in a Swept-Wing Boundary Layer (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Direct numerical simulation (DNS) is performed to examine laminar to turbulent transition due to high-frequency secondary instability of stationary crossflow vortices in a subsonic swept-wing boundary layer for a realistic natural-laminar-flow airfoil configuration. The secondary instability is introduced via inflow forcing derived from a two-dimensional, partial-differential-equation based eigenvalue computation; and the mode selected for forcing corresponds to the most amplified secondary instability mode which, in this case, derives a majority of its growth from energy production mechanisms associated with the wall-normal shear of the stationary basic state. Both the growth of the secondary instability wave and the resulting onset of laminar-turbulent transition are captured within the DNS computations. The growth of the secondary instability wave in the DNS solution compares well with linear secondary instability theory when the amplitude is small; the linear growth is followed by a region of reduced growth resulting from nonlinear effects before an explosive onset of laminar breakdown to turbulence. The peak fluctuations are concentrated near the boundary layer edge during the initial stage of transition, but rapidly propagates towards the surface during the process of laminar breakdown. Both time-averaged statistics and flow visualization based on the DNS reveal a sawtooth transition pattern that is analogous to previously documented surface flow visualizations of transition due to stationary crossflow instability. The memory of the stationary crossflow vortex is found to persist through the transition zone and well beyond the location of the maximum skin friction.
    Keywords: Aerodynamics
    Type: NF1676L-15634 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Developing an Accurate CFD Based Gust Model for the Truss Braced Wing Aircraft (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The increased flexibility of long endurance aircraft having high aspect ratio wings necessitates attention to gust response and perhaps the incorporation of gust load alleviation. The design of civil transport aircraft with a strut or truss-braced high aspect ratio wing furthermore requires gust response analysis in the transonic cruise range. This requirement motivates the use of high fidelity nonlinear computational fluid dynamics (CFD) for gust response analysis. This paper presents the development of a CFD based gust model for the truss braced wing aircraft. A sharp-edged gust provides the gust system identification. The result of the system identification is several thousand time steps of instantaneous pressure coefficients over the entire vehicle. This data is filtered and downsampled to provide the snapshot data set from which a reduced order model is developed. A stochastic singular value decomposition algorithm is used to obtain a proper orthogonal decomposition (POD). The POD model is combined with a convolution integral to predict the time varying pressure coefficient distribution due to a novel gust profile. Finally the unsteady surface pressure response of the truss braced wing vehicle to a one-minus-cosine gust, simulated using the reduced order model, is compared with the full CFD.
    Keywords: Aerodynamics
    Type: NF1676L-15527 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Testing Small CPAS Parachutes Using HIVAS (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The High Velocity Airflow System (HIVAS) facility at the Naval Air Warfare Center (NAWC) at China Lake was successfully used as an alternative to flight test to determine parachute drag performance of two small Capsule Parachute Assembly System (CPAS) canopies. A similar parachute with known performance was also tested as a control. Realtime computations of drag coefficient were unrealistically low. This is because HIVAS produces a non-uniform flow which rapidly decays from a high central core flow. Additional calibration runs were performed to characterize this flow assuming radial symmetry from the centerline. The flow field was used to post-process effective flow velocities at each throttle setting and parachute diameter using the definition of the momentum flux factor. Because one parachute had significant oscillations, additional calculations were required to estimate the projected flow at off-axis angles. The resulting drag data from HIVAS compared favorably to previously estimated parachute performance based on scaled data from analogous CPAS parachutes. The data will improve drag area distributions in the next version of the CPAS Model Memo.
    Keywords: Aerodynamics
    Type: JSC-CN-28250 , 22nd AIAA Aerodynamic Delerator Systems; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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    Flutter Analysis of the Thermal Protection Layer on the NASA HIAD (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A combination of classical plate theory and a supersonic aerodynamic model is used to study the aeroelastic flutter behavior of a proposed thermal protection system (TPS) for the NASA HIAD. The analysis pertains to the rectangular configurations currently being tested in a NASA wind-tunnel facility, and may explain why oscillations of the articles could be observed. An analysis using a linear flat plate model indicated that flutter was possible well within the supersonic flow regime of the wind tunnel tests. A more complex nonlinear analysis of the TPS, taking into account any material curvature present due to the restraint system or substructure, indicated that significantly greater aerodynamic forcing is required for the onset of flutter. Chaotic and periodic limit cycle oscillations (LCOs) of the TPS are possible depending on how the curvature is imposed. When the pressure from the base substructure on the bottom of the TPS is used as the source of curvature, the flutter boundary increases rapidly and chaotic behavior is eliminated.
    Keywords: Aerodynamics
    Type: NF1676L-16134 , 22nd AIAA Aerodynamic Decelerator Systems Technology Conference; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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    Overview of the Aeroelastic Prediction Workshop (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The AIAA Aeroelastic Prediction Workshop (AePW) was held in April, 2012, bringing together communities of aeroelasticians and computational fluid dynamicists. The objective in conducting this workshop on aeroelastic prediction was to assess state-of-the-art computational aeroelasticity methods as practical tools for the prediction of static and dynamic aeroelastic phenomena. No comprehensive aeroelastic benchmarking validation standard currently exists, greatly hindering validation and state-of-the-art assessment objectives. The workshop was a step towards assessing the state of the art in computational aeroelasticity. This was an opportunity to discuss and evaluate the effectiveness of existing computer codes and modeling techniques for unsteady flow, and to identify computational and experimental areas needing additional research and development. Three configurations served as the basis for the workshop, providing different levels of geometric and flow field complexity. All cases considered involved supercritical airfoils at transonic conditions. The flow fields contained oscillating shocks and in some cases, regions of separation. The computational tools principally employed Reynolds-Averaged Navier Stokes solutions. The successes and failures of the computations and the experiments are examined in this paper.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-1798 , NF1676L-15301 , 54th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Apr 08, 2013 - Apr 11, 2013; Boston, MA
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    Experimental Data from the Benchmark SuperCritical Wing Wind Tunnel Test on an Oscillating Turntable (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Benchmark SuperCritical Wing (BSCW) wind tunnel model served as a semi-blind testcase for the 2012 AIAA Aeroelastic Prediction Workshop (AePW). The BSCW was chosen as a testcase due to its geometric simplicity and flow physics complexity. The data sets examined include unforced system information and forced pitching oscillations. The aerodynamic challenges presented by this AePW testcase include a strong shock that was observed to be unsteady for even the unforced system cases, shock-induced separation and trailing edge separation. The current paper quantifies these characteristics at the AePW test condition and at a suggested benchmarking test condition. General characteristics of the model's behavior are examined for the entire available data set.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-1802 , NF1676L-15289 , 54th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Apr 08, 2013 - Apr 11, 2013; Boston, MA; United States
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    Multi-Mission Earth Vehicle Subsonic Dynamic Stability Testing and Analyses (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Multi-Mission Earth Entry Vehicles (MMEEVs) are blunt-body vehicles designed with the purpose of transporting payloads from outer space to the surface of the Earth. To achieve high-reliability and minimum weight, MMEEVs avoid use of limited-reliability systems, such as parachutes, retro-rockets, and reaction control systems and rely on the natural aerodynamic stability of the vehicle throughout the Entry, Descent, and Landing (EDL) phase of flight. The Multi-Mission Systems Analysis for Planetary Entry (M-SAPE) parametric design tool is used to facilitate the design of MMEEVs for an array of missions and develop and visualize the trade space. Testing in NASA Langley?s Vertical Spin Tunnel (VST) was conducted to significantly improve M-SAPE?s subsonic aerodynamic models. Vehicle size and shape can be driven by entry flight path angle and speed, thermal protection system performance, terminal velocity limitations, payload mass and density, among other design parameters. The objectives of the VST testing were to define usable subsonic center of gravity limits, and aerodynamic parameters for 6-degree-of-freedom (6-DOF) simulations, for a range of MMEEV designs. The range of MMEEVs tested was from 1.8m down to 1.2m diameter. A backshell extender provided the ability to test a design with a much larger payload for the 1.2m MMEEV.
    Keywords: Aerodynamics
    Type: NF1676L-16196 , International Planetary Probe Workshop (IPPW-10); Jun 17, 2013 - Jun 21, 2013; San Jose, CA; United States
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    Space Shuttle Orbiter Drag Chute Summary (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper summarizes the development history and technical highlights of the Space Shuttle Orbiter Drag Chute Program. Data and references are given on the design, development, and testing of the system, plus several interesting operational issues and solutions. The last Shuttle flight was completed in 2011 and all the Orbiters have now become museum pieces. Before all the data from system development and the 86 Orbiter Drag Chute (ODC) operational landings is lost or forgotten, it may be useful to summarize it here and to identify data sources for future reference. Much has been written about various aspects of the program, and this summary has attempted to cite many such references to make available more detailed information. The ODC program was a high-visibility NASA program that afforded the opportunity to thoroughly engineer and test the chute system, far beyond so many of today s tight-budget programs. So the ODC program was extremely informative--it provided a wide scope of information including protective door jettison issues and solutions, wind tunnel data and analyses on chute stability and drag behind a huge and rather blunt forebody, component and system reuse, and chute cleaning methods. Technology and data created have aided several current and past parachute programs, and will continue to do so in the future. The original Orbiter preliminary design included a drag parachute-- it was deleted early to save weight. But after the 1987 Challenger accident and during the program redefinition phase that followed, Astronaut John Young presented a strong case for enhancing landing safety by adding nosegear steering, brake improvements, and reviving the drag chute.
    Keywords: Aerodynamics
    Type: JSC-CN-28249 , 22nd AIAA Aerodynamic Deceleration System Meeting; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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    Assessment of the Reconstructed Aerodynamics of the Mars Science Laboratory Entry Vehicle (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: On August 5, 2012, the Mars Science Laboratory entry vehicle successfully entered Mars atmosphere, flying a guided entry until parachute deploy. The Curiosity rover landed safely in Gale crater upon completion of the Entry Descent and Landing sequence. This paper compares the aerodynamics of the entry capsule extracted from onboard flight data, including Inertial Measurement Unit (IMU) accelerometer and rate gyro information, and heatshield surface pressure measurements. From the onboard data, static force and moment data has been extracted. This data is compared to preflight predictions. The information collected by MSL represents the most complete set of information collected during Mars entry to date. It allows the separation of aerodynamic performance from atmospheric conditions. The comparisons show the MSL aerodynamic characteristics have been identified and resolved to an accuracy better than the aerodynamic database uncertainties used in preflight simulations. A number of small anomalies have been identified and are discussed. This data will help revise aerodynamic databases for future missions and will guide computational fluid dynamics (CFD) development to improved prediction codes.
    Keywords: Aerodynamics
    Type: AAS13-306 , NF1676L-15415 , 23rd AAS/AIAA Space Flight Mechanics Meeting; Feb 10, 2013 - Feb 14, 2013; Kauai, HI; United States
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    Discrete Surface Evolution and Mesh Deformation for Aircraft Icing Applications (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Robust, automated mesh generation for problems with deforming geometries, such as ice accreting on aerodynamic surfaces, remains a challenging problem. Here we describe a technique to deform a discrete surface as it evolves due to the accretion of ice. The surface evolution algorithm is based on a smoothed, face-offsetting approach. We also describe a fast algebraic technique to propagate the computed surface deformations into the surrounding volume mesh while maintaining geometric mesh quality. Preliminary results presented here demonstrate the ecacy of the approach for a sphere with a prescribed accretion rate, a rime ice accretion, and a more complex glaze ice accretion.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN9768 , Atmospheric and Space Envronments Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    The Effect of Bypass Nozzle Exit Area on Fan Aerodynamic Performance and Noise in a Model Turbofan Simulator (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted called the Fan Broadband Source Diagnostic Test as part of the NASA Quiet Aircraft Technology program. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator which represented a current generation, medium pressure ratio, high bypass turbofan aircraft engine. The investigation focused on simulating in model scale only the bypass section of the turbofan engine. The test objectives were to: identify the noise sources within the model and determine their noise level; investigate several component design technologies by determining their impact on the aerodynamic and acoustic performance of the fan stage; and conduct detailed flow diagnostics within the fan flow field to characterize the physics of the noise generation mechanisms in a turbofan model. This report discusses results obtained for one aspect of the Source Diagnostic Test that investigated the effect of the bypass or fan nozzle exit area on the bypass stage aerodynamic performance, specifically the fan and outlet guide vanes or stators, as well as the farfield acoustic noise level. The aerodynamic performance, farfield acoustics, and Laser Doppler Velocimeter flow diagnostic results are presented for the fan and four different fixed-area bypass nozzle configurations. The nozzles simulated fixed engine operating lines and encompassed the fan stage operating envelope from near stall to cruise. One nozzle was selected as a baseline reference, representing the nozzle area which would achieve the design point operating conditions and fan stage performance. The total area change from the smallest to the largest nozzle was 12.9 percent of the baseline nozzle area. The results will show that there are significant changes in aerodynamic performance and farfield acoustics as the fan nozzle area is increased. The weight flow through the fan model increased between 7 and 9 percent, the fan and stage pressure dropped between 8 and 10 percent, and the adiabatic efficiency increased between 2 and 3 percent--the magnitude of the change dependent on the fan speed. Results from force balance measurements of fan and outlet guide vane thrust will show that as the nozzle exit area is increased the combined thrust of the fan and outlet guide vanes together also increases, between 2 and 3.5 percent, mainly due to the increase in lift from the outlet guide vanes. In terms of farfield acoustics, the overall sound power level produced by the fan stage dropped nearly linearly between 1 dB at takeoff condition and 3.5 dB at approach condition, mainly due to a decrease in the broadband noise levels. Finally, fan swirl angle survey and Laser Doppler Velocimeter mean velocity and turbulence data obtained in the fan wake will show that the swirl angles and turbulence levels within the wake decrease as the fan nozzle area increases, which helps to explain the drop in the fan broadband noise at all fan speeds.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-214029 , GT2005-68573 , E-15384 , Turbo Expo 2005; Jun 06, 2005 - Jun 09, 2005; Reno, NV; United States
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    Investigation of Materials for Boundary Layer Control in a Supersonic Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: During operation of the NASA Glenn Research Center 15- by 15-Centimeter Supersonic Wind Tunnel (SWT), a significant, undesirable corner flow separation is created by the three-dimensional interaction of the wall and floor boundary layers in the tunnel corners following an oblique-shock/ boundary-layer interaction. A method to minimize this effect was conceived by connecting the wall and floor boundary layers with a radius of curvature in the corners. The results and observations of a trade study to determine the effectiveness of candidate materials for creating the radius of curvature in the SWT are presented. The experiments in the study focus on the formation of corner fillets of four different radii of curvature, 6.35 mm (0.25 in.), 9.525 mm (0.375 in.), 12.7 mm (0.5 in.), and 15.875 mm (0.625 in.), based on the observed boundary layer thickness of 11.43 mm (0.45 in.). Tests were performed on ten candidate materials to determine shrinkage, surface roughness, cure time, ease of application and removal, adhesion, eccentricity, formability, and repeatability. Of the ten materials, the four materials which exhibited characteristics most promising for effective use were the heavy body and regular type dental impression materials, the basic sculpting epoxy, and the polyurethane sealant. Of these, the particular material which was most effective, the heavy body dental impression material, was tested in the SWT in Mach 2 flow, and was observed to satisfy all requirements for use in creating the corner fillets in the upcoming experiments on shock-wave/boundary-layer interaction.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217894 , E-18708 , Machinery Failure Prevention Technology (MFPT) 2013; May 13, 2013 - May 17, 2013; Cleveland, OH; United States|International Instrumentaiton Symposium (IIS); May 13, 2013 - May 17, 2013; Cleveland, OH; United States
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    Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade at Low Inlet Turbulence (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to 51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.1210(exp 5) to 2.1210(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and 36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range. deg
    Keywords: Aerodynamics
    Type: NASA/TM-2013-218069 , GT2013-94695 , E-18745 , ASME Turbo Expo 2013; Jun 03, 2013 - Jun 07, 2013; San Antonio, TX; United States
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    Comparison of Force and Moment Coefficients for the Same Test Article in Multiple Wind Tunnels (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper compares the results of force and moment measurements made on the same test article and with the same balance in three transonic wind tunnels. Comparisons are made for the same combination of Reynolds number, Mach number, sideslip angle, control surface configuration, and angle of attack range. Between-tunnel force and moment differences are quantified. An analysis of variance was performed at four unique sites in the design space to assess the statistical significance of between-tunnel variation and any interaction with angle of attack. Tunnel to tunnel differences too large to attribute to random error were detected were observed for all forces and moments. In some cases these differences were independent of angle of attack and in other cases they changed with angle of attack.
    Keywords: Aerodynamics
    Type: NF1676L-16794 , AIAA Ground Testing Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Investigations of Fluid-Structure-Coupling and Turbulence Model Effects on the DLR Results of the Fifth AIAA CFD Drag Prediction Workshop (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The accurate calculation of aerodynamic forces and moments is of significant importance during the design phase of an aircraft. Reynolds-averaged Navier-Stokes (RANS) based Computational Fluid Dynamics (CFD) has been strongly developed over the last two decades regarding robustness, efficiency, and capabilities for aerodynamically complex configurations. Incremental aerodynamic coefficients of different designs can be calculated with an acceptable reliability at the cruise design point of transonic aircraft for non-separated flows. But regarding absolute values as well as increments at off-design significant challenges still exist to compute aerodynamic data and the underlying flow physics with the accuracy required. In addition to drag, pitching moments are difficult to predict because small deviations of the pressure distributions, e.g. due to neglecting wing bending and twisting caused by the aerodynamic loads can result in large discrepancies compared to experimental data. Flow separations that start to develop at off-design conditions, e.g. in corner-flows, at trailing edges, or shock induced, can have a strong impact on the predictions of aerodynamic coefficients too. Based on these challenges faced by the CFD community a working group of the AIAA Applied Aerodynamics Technical Committee initiated in 2001 the CFD Drag Prediction Workshop (DPW) series resulting in five international workshops. The results of the participants and the committee are summarized in more than 120 papers. The latest, fifth workshop took place in June 2012 in conjunction with the 30th AIAA Applied Aerodynamics Conference. The results in this paper will evaluate the influence of static aeroelastic wing deformations onto pressure distributions and overall aerodynamic coefficients based on the NASA finite element structural model and the common grids.
    Keywords: Aerodynamics
    Type: NF1676L-16710 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Towards Bridging the Gaps in Holistic Transition Prediction via Numerical Simulations (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The economic and environmental benefits of laminar flow technology via reduced fuel burn of subsonic and supersonic aircraft cannot be realized without minimizing the uncertainty in drag prediction in general and transition prediction in particular. Transition research under NASA's Aeronautical Sciences Project seeks to develop a validated set of variable fidelity prediction tools with known strengths and limitations, so as to enable "sufficiently" accurate transition prediction and practical transition control for future vehicle concepts. This paper provides a summary of selected research activities targeting the current gaps in high-fidelity transition prediction, specifically those related to the receptivity and laminar breakdown phases of crossflow induced transition in a subsonic swept-wing boundary layer. The results of direct numerical simulations are used to obtain an enhanced understanding of the laminar breakdown region as well as to validate reduced order prediction methods.
    Keywords: Aerodynamics
    Type: NF1676L-15705 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Development of a Multi-Disciplinary Aerothermostructural Model Applicable to Hypersonic Flight (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The harsh and complex hypersonic flight environment has driven design and analysis improvements for many years. One of the defining characteristics of hypersonic flight is the coupled, multi-disciplinary nature of the dominant physics. In an effect to examine some of the multi-disciplinary problems associated with hypersonic flight engineers at the NASA Dryden Flight Research Center developed a non-linear 6 degrees-of-freedom, full vehicle simulation that includes the necessary model capabilities: aerothermal heating, ablation, and thermal stress solutions. Development of the tool and results for some investigations will be presented. Requirements and improvements for future work will also be reviewed. The results of the work emphasize the need for a coupled, multi-disciplinary analysis to provide accurate
    Keywords: Aerodynamics
    Type: DFRC-E-DAA-TN10322 , Thermal and Fluids Analysis Workshop; Jul 29, 2013 - Aug 02, 2013; Daytona Beach, FL; United States
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    The HART II International Workshop: An Assessment of the State-of-the-Art in Comprehensive Code Prediction (2013)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Significant advancements in computational fluid dynamics (CFD) and their coupling with computational structural dynamics (CSD, or comprehensive codes) for rotorcraft applications have been achieved recently. Despite this, CSD codes with their engineering level of modeling the rotor blade dynamics, the unsteady sectional aerodynamics and the vortical wake are still the workhorse for the majority of applications. This is especially true when a large number of parameter variations is to be performed and their impact on performance, structural loads, vibration and noise is to be judged in an approximate yet reliable and as accurate as possible manner. In this article, the capabilities of such codes are evaluated using the HART II International Workshop database, focusing on a typical descent operating condition which includes strong blade-vortex interactions. A companion article addresses the CFD/CSD coupled approach. Three cases are of interest: the baseline case and two cases with 3/rev higher harmonic blade root pitch control (HHC) with different control phases employed. One setting is for minimum blade-vortex interaction noise radiation and the other one for minimum vibration generation. The challenge is to correctly predict the wake physics-especially for the cases with HHC-and all the dynamics, aerodynamics, modifications of the wake structure and the aero-acoustics coming with it. It is observed that the comprehensive codes used today have a surprisingly good predictive capability when they appropriately account for all of the physics involved. The minimum requirements to obtain these results are outlined.
    Keywords: Aerodynamics
    Type: NF1676L-17413 , Annual Forum of the American Helicopter Society (AHS); May 01, 2012 - May 03, 2012; Fort Worth, TX; United States|CEAS Aeronautical Journal: An Official Journal of the Council of European Aerospace Societies ; 4; 3; 223-252
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    Numerical Investigation of a Fuselage Boundary Layer Ingestion Propulsion Concept (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In the present study, a numerical assessment of the performance of fuselage boundary layer ingestion (BLI) propulsion techniques was conducted. This study is an initial investigation into coupling the aerodynamics of the fuselage with a BLI propulsion system to determine if there is sufficient potential to warrant further investigation of this concept. Numerical simulations of flow around baseline, Boundary Layer Controlled (BLC), and propelled boundary layer controlled airships were performed. Computed results showed good agreement with wind tunnel data and previous numerical studies. Numerical simulations and sensitivity analysis were then conducted on four BLI configurations. The two design variables selected for the parametric study of the new configurations were the inlet area and the inlet to exit area ratio. Current results show that BLI propulsors may offer power savings of up to 85% over the baseline configuration. These interim results include the simplifying assumption that inlet ram drag is negligible and therefore likely overstate the reduction in power. It has been found that inlet ram drag is not negligible and should be included in future analysis.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-4402 , NF1676L-17000 , AIAA Aviation Technology, Integration, and Operations (ATIO) Conference; Aug 12, 2013 - Aug 14, 2013; Los Angeles, CA; United States
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    A CFD Database for Airfoils and Wings at Post-Stall Angles of Attack (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents selected results from an ongoing effort to develop an aerodynamic database from Reynolds-Averaged Navier-Stokes (RANS) computational analysis of airfoils and wings at stall and post-stall angles of attack. The data obtained from this effort will be used for validation and refinement of a low-order post-stall prediction method developed at NCSU, and to fill existing gaps in high angle of attack data in the literature. Such data could have potential applications in post-stall flight dynamics, helicopter aerodynamics and wind turbine aerodynamics. An overview of the NASA TetrUSS CFD package used for the RANS computational approach is presented. Detailed results for three airfoils are presented to compare their stall and post-stall behavior. The results for finite wings at stall and post-stall conditions focus on the effects of taper-ratio and sweep angle, with particular attention to whether the sectional flows can be approximated using two-dimensional flow over a stalled airfoil. While this approximation seems reasonable for unswept wings even at post-stall conditions, significant spanwise flow on stalled swept wings preclude the use of two-dimensional data to model sectional flows on swept wings. Thus, further effort is needed in low-order aerodynamic modeling of swept wings at stalled conditions.
    Keywords: Aerodynamics
    Type: NF1676L-16579 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    High-Order Simulation of Induced Disturbance in a Mach 6 Boundary Layer (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Simulations of the early evolution of induced unstable disturbances in a Mach 6 hypersonic boundary layer are presented for the purposes of validating a high-order discontinuous Galerkin Navier-Stokes simulation code. The simulations are modeled after experiments performed in the Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue University. In these experiments, a forcing mechanism was employed to induce reproducible disturbances in a hypersonic boundary layer providing for the controlled study of the growth and breakdown of these disturbances into turbulent spots. Simulations revealed that the form and strength of the excitation can greatly influence the growth of the disturbance. In particular, at large forcing amplitude, the simulated forcing produces large advecting transients that appear to enhance the growth of the wave packet, relative to that of low amplitude forcing. Simulation results with large amplitude forcing agree well with experimental results while results from low amplitude forcing agree with linear stability theory.
    Keywords: Aerodynamics
    Type: NF1676L-15628 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Unique Testing Capabilities of the NASA Langley Transonic Dynamics Tunnel, an Exercise in Aeroelastic Scaling (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA Langley Research Center's Transonic Dynamics Tunnel (TDT) is the world's most capable aeroelastic test facility. Its large size, transonic speed range, variable pressure capability, and use of either air or R-134a heavy gas as a test medium enable unparalleled manipulation of flow-dependent scaling quantities. Matching these scaling quantities enables dynamic similitude of a full-scale vehicle with a sub-scale model, a requirement for proper characterization of any dynamic phenomenon, and many static elastic phenomena. Select scaling parameters are presented in order to quantify the scaling advantages of TDT and the consequence of testing in other facilities. In addition to dynamic testing, the TDT is uniquely well-suited for high risk testing or for those tests that require unusual model mount or support systems. Examples of recently conducted dynamic tests requiring unusual model support are presented. In addition to its unique dynamic test capabilities, the TDT is also evaluated in its capability to conduct aerodynamic performance tests as a result of its flow quality. Results of flow quality studies and a comparison to a many other transonic facilities are presented. Finally, the ability of the TDT to support future NASA research thrusts and likely vehicle designs is discussed.
    Keywords: Aerodynamics
    Type: NF1676L-15532 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    A New Formulation for Hybrid LES-RANS Computations (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Ideally, a hybrid LES-RANS computation would employ LES only where necessary to make up for the failure of the RANS model to provide sufficient accuracy or to provide time-dependent information. Current approaches are fairly restrictive in the placement of LES and RANS regions; an LES-RANS transition in a boundary layer, for example, yields an unphysical log-layer shift. A hybrid computation is formulated here to allow greater control over the placement of LES and RANS regions and the transitions between them. The concept of model invariance is introduced, which provides a basis for interpreting hybrid results within an LES-RANS transition zone. Consequences of imposing model invariance include the addition of terms to the governing equations that compensate for unphysical gradients created as the model changes between RANS and LES. Computational results illustrate the increased accuracy of the approach and its insensitivity to the location of the transition and to the blending function employed.
    Keywords: Aerodynamics
    Type: NF1676L-15665 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Fundamental Inlet Bleed Experiments (FIBE) (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Aerodynamics
    Type: E-664404 , Annual Shock Wave/Boundary Layer Interaction (SBLI) Workshop; Apr 24, 2013 - Apr 25, 2013; Dayton, OH; United States
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    Trajectory Optimization Using Adjoint Method and Chebyshev Polynomial Approximation for Minimizing Fuel Consumption During Climb (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes two methods of trajectory optimization to obtain an optimal trajectory of minimum-fuel- to-climb for an aircraft. The first method is based on the adjoint method, and the second method is based on a direct trajectory optimization method using a Chebyshev polynomial approximation and cubic spine approximation. The approximate optimal trajectory will be compared with the adjoint-based optimal trajectory which is considered as the true optimal solution of the trajectory optimization problem. The adjoint-based optimization problem leads to a singular optimal control solution which results in a bang-singular-bang optimal control.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN10840 , AIAA Infotech@Aerospace 2013 Conference; Aug 19, 2013 - Aug 22, 2013; Boston, MA; United States
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    Supersonic Aerodynamic Characteristics of Blunt Body Trim Tab Configurations (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Trim tabs are aerodynamic control surfaces that can allow an entry vehicle to meet aerodynamic performance requirements while reducing or eliminating the use of ballast mass and providing a capability to modulate the lift-to-drag ratio during entry. Force and moment data were obtained on 38 unique, blunt body trim tab configurations in the NASA Langley Research Center Unitary Plan Wind Tunnel. The data were used to parametrically assess the supersonic aerodynamic performance of trim tabs and to understand the influence of tab area, cant angle, and aspect ratio. Across the range of conditions tested (Mach numbers of 2.5, 3.5, and 4.5; angles of attack from -4deg to +20deg; angles of sideslip from 0deg to +8deg), the effects of varying tab area and tab cant angle were found to be much more significant than effects from varying tab aspect ratio. Aerodynamic characteristics exhibited variation with Mach number and forebody geometry over the range of conditions tested. Overall, the results demonstrate that trim tabs are a viable approach to satisfy aerodynamic performance requirements of blunt body entry vehicles with minimal ballast mass. For a 70deg sphere-cone, a tab with 3% area of the forebody and canted approximately 35deg with no ballast mass was found to give the same trim aerodynamics as a baseline model with ballast mass that was 5% of the total entry mass.
    Keywords: Aerodynamics
    Type: NF1676L-16666 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Nonlinear Modeling by Assembling Piecewise Linear Models (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: To preserve nonlinearity of a full order system over a parameters range of interest, we propose a simple modeling approach by assembling a set of piecewise local solutions, including the first-order Taylor series terms expanded about some sampling states. The work by Rewienski and White inspired our use of piecewise linear local solutions. The assembly of these local approximations is accomplished by assigning nonlinear weights, through radial basis functions in this study. The efficacy of the proposed procedure is validated for a two-dimensional airfoil moving at different Mach numbers and pitching motions, under which the flow exhibits prominent nonlinear behaviors. All results confirm that our nonlinear model is accurate and stable for predicting not only aerodynamic forces but also detailed flowfields. Moreover, the model is robustness-accurate for inputs considerably different from the base trajectory in form and magnitude. This modeling preserves nonlinearity of the problems considered in a rather simple and accurate manner.
    Keywords: Aerodynamics
    Type: E-664078 , AIAA Fluid Dynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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    Turbulence Measurements in Rectangular Jets with Aft Deck (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Aerodynamics
    Type: E-664045 , Acoustics Technical Working Group; Apr 23, 2013 - Apr 24, 2013; Cleveland, OH; United States
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    Aerodynamic Stability and Performance of Next-Generation Parachutes for Mars Descent (2013)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The Low Density Supersonic Decelerator Project is developing a next-generation supersonic parachute for use on future Mars missions. In order to determine the new parachute configuration, a wind tunnel test was conducted at the National Full-scale Aerodynamics Complex 80- by 120-foot Wind Tunnel at the NASA Ames Research Center. The goal of the wind tunnel test was to quantitatively determine the aerodynamic stability and performance of various canopy configurations in order to help select the design to be flown on the Supersonic Flight Dynamics tests. Parachute configurations included the diskgap- band, ringsail, and ringsail-variant designs referred to as a disksail and starsail. During the wind tunnel test, digital cameras captured synchronized image streams of the parachute from three directions. Stereo hotogrammetric processing was performed on the image data to track the position of the vent of the canopy throughout each run. The position data were processed to determine the geometric angular history of the parachute, which were then used to calculate the total angle of attack and its derivatives at each instant in time. Static and dynamic moment coefficients were extracted from these data using a parameter estimation method involving the one-dimensional equation of motion for a rotation of parachute. The coefficients were calculated over all of the available canopy states to reconstruct moment coefficient curves as a function of total angle of attack. From the stability curves, useful metrics such as the trim total angle of attack and pitch stiffness at the trim angle could be determined. These stability metrics were assessed in the context of the parachute's drag load and geometric porosity. While there was generally an inverse relationship between the drag load and the stability of the canopy, the data showed that it was possible to obtain similar stability properties as the disk-gap-band with slightly higher drag loads by appropriately tailoring the geometric porosity distribution.
    Keywords: Aerodynamics
    Type: AIAA Aerodynamic Decelerator Systems Technology Conference; Mar 25, 2013 - Mar 28, 2013; Daytona Beach, FL; United States
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