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  • Aerodynamics
  • Inorganic Chemistry
  • LUNAR AND PLANETARY EXPLORATION
  • Tectonics
  • Temperatur
  • 2010-2014  (57)
  • 1930-1934
  • 2014  (57)
  • 1
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    Quantitative Entwicklung von Flachseen in Brandenburg im Klimawandel (2014)
    In:  Kaden, S., Dietrich, O., Theobald, S. (Hrsg.). Wassermanagement im Klimawandel – Möglichkeiten und Grenzen von Anpassungsmaßnahmen. Berlin: oekom in Erscheinen
    Details
    Publication Date: 2014
    Description: Teilergebnise des BMBF-Verbundforschungsprojektes INKA BB, des TP 22 „Nachhaltige Managementstrategien für glaziale Seen Brandenburgs im Klimawandel“. Ergebnisse aus Berechnungen mit dem NA-Modell EGMO-D für STAR2-Szenarion mit 0 und 2K Temperaturanstieg zur Wasserstandsentwicklung von Flachseen im Raum Brandenburg KATASTER-BESCHREIBUNG: Absenkung der Seewasserspiegel im Extremfall um mehrere Meter, am größten in den Seen im Südosten Berlins und nehmen zum Nordwesten ab, Zunahme der Absenkungen ist am stärksten für Seen, welche im Vergleich zum Zufluss ein großes Volumen besitzen KATASTER-DETAIL: Tmit für (ECAHM, A1B, STAR 0.0T, d.h. T1 0.0 ... kein weiterer Temperaturanstieg nach 2000 und STAR 2.0K, d.h. T2.0 ... weiterer Temperaturanstieg nach 2000 um ca. 2K bis 2060), bis zum Jahr 2018 Stationarität des Füllungsregimes, erst danach Absenkungen zu erwarten Tmit (0K und 2K), von 0,5 m im Jahr 2018 bis auf 4,2 m im Jahr 2053, diejenigen mit 10% Überschreitungswahrscheinlichkeit von 0,4 m bis auf 2,4 m
    Keywords: Nordostdeutschland ; Szenarien 2004-2053 ; Korrelationsmethode ; Niederschlag ; Temperatur ; Verdunstung ; Abfluss ; Grundwasser
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    BIBLIOGRAPHY INKA BB
  • 2
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    Auswirkungen potenzieller Änderungen des Klimas und der Braunkohleförderung und -verstromung auf natürliche und bewirtschaftete Abflüsse in den Lausitzer Flusseinzugsgebieten (2014)
    In:  Kaden, S., Dietrich, O., Theobald, S. (Hrsg.). Wassermanagement im Klimawandel – Möglichkeiten und Grenzen von Anpassungsmaßnahmen. München: oekom verlag, 2014, S. 121-140
    Details
    Publication Date: 2014
    Description: Ergebnisse von Simulationen der potenziellen zukünftigen natürlichen bzw. bewirtschafteten Abflüsse auf Basis der hydrologischen Modelle SWIM, EGMO, WaSiM-ETH und HBV-light bzw. des Langfristbewirtschaftungsmodells WBalMo für Klimaszenarien von STAR (weiterer Temperaturanstieg um 0 K, 2 K bzw. 3 K) sowie WettReg A1B und eines Szenarios zur Braunkohleförderung und -verstromung (Rückgang des Grundwasserabsenkungsgebiets und der Sümpfungswassereinleitung) KATASTER-BESCHREIBUNG: Rückgang der mittleren natürlichen und bewirtschafteten Abflüsse aufgrund steigender Temperaturen (und der potenziellen Evapotranspiration) sowie teilweise Rückgang der Niederschläge Bei den natürlichen Abflüssen sind die Effekte des Rückgangs des Grundwasserabsenkungsgebiets und der damit verbundenen Vergrößerung der abflusswirksamen Fläche geringer als die Auswirkungen der klimatischen Änderungen. Unsicherheiten der Ergebnisse sowohl aufgrund der Wahl des Klimaszenarios als auch des hydrologischen Modells KATASTER-DETAIL: Delte T (STAR 0K, : 2013-53 vs. 1961-1990) leichter Anstieg der natürlichen Abflüsse durch Rückgang des Grundwasserabsenkungsgebiets; jedoch Rückgang im Vergleich zur Referenzperiode um bis zu 20 % Delte T (STAR 2K, STAR 3K und WettReg A1B: Rückgang der natürlichen Abflüsse im Verlauf des Szenariozeitraums, dabei Rückgang der mittleren Abflüsse (MQ)m Vergleich zur Referenzperiode um bis zu 40 % (STAR 2K) bzw. 50 % (STAR 3K und WettReg A1B), dieser Effekt wird zum Teil durch Rückgang des Grundwasserabsenkungsgebiets gemildert. Bewirtschaftete Abflüsse STAR 0K: z.T. Anstieg bis 2030 STAR 2K, STAR 3K: deutlicher Rückgang ab 2030
    Keywords: Lausitz ; 2013-2053 ; Klima ; Niederschlag ; Temperatur ; Abfluss ; Grundwasser ; Modell
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    BIBLIOGRAPHY INKA BB
  • 3
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    Indoor heat stress: An assessment of human bioclimate using the UTCI in different buildings in Berlin (2014)
    In:  Die Erde, 144, 3, 260-273
    Details
    Publication Date: 2014
    Description: Zusammenhang zwischen Hitzestress in Räumen mit deren Lage. Veranschaulicht durch Messungen von Temperatur, Luftfeuchte und berechnetem UTCI Universal Thermal Climate Index. KATASTER-BESCHREIBUNG: KATASTER-DETAIL:
    Keywords: Berlin ; 2011-2012 ; Umweltmedizin ; Luftfeuchte ; Temperatur ; Witterungsextreme
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    BIBLIOGRAPHY INKA BB
  • 4
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    Spatial analysis of hospital admissions for respiratory diseases during summer months in Berlin taking bioclimatic and socio-economic aspects into account (2014)
    In:  Die Erde, 144, 3, 217-237
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    Publication Date: 2014
    Description: Ansatz zur Identifizierung hitzevulnerabler Stadtgebiete über räumlich-edidemiologische Analysen. KATASTER-BESCHREIBUNG: KATASTER-DETAIL:
    Keywords: Berlin ; 2000-2009 ; Umweltmedizin ; Temperatur ; Witterungsextreme ; Luftverunreinigungen
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    BIBLIOGRAPHY INKA BB
  • 5
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    An Examination of Parameters Affecting Large Eddy Simulations of Flow Past a Square Cylinder (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Separated flow over a bluff body is analyzed via large eddy simulations. The turbulent flow around a square cylinder features a variety of complex flow phenomena such as highly unsteady vortical structures, reverse flow in the near wall region, and wake turbulence. The formation of spanwise vortices is often times artificially suppressed in computations by either insufficient depth or a coarse spanwise resolution. As the resolution is refined and the domain extended, the artificial turbulent energy exchange between spanwise and streamwise turbulence is eliminated within the wake region. A parametric study is performed highlighting the effects of spanwise vortices where the spanwise computational domain's resolution and depth are varied. For Re=22,000, the mean and turbulent statistics computed from the numerical large eddy simulations (NLES) are in good agreement with experimental data. Von-Karman shedding is observed in the wake of the cylinder. Mesh independence is illustrated by comparing a mesh resolution of 2 million to 16 million. Sensitivities to time stepping were minimized and sampling frequency sensitivities were nonpresent. While increasing the spanwise depth and resolution can be costly, this practice was found to be necessary to eliminating the artificial turbulent energy exchange.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-2089 , E-663992 , AIAA Fluid Dynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 6
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    Wedge Shock and Nozzle Exhaust Plume Interaction in a Supersonic Jet Flow (2014)
    In:  CASI
    Details
    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 nozzle exhaust plume. Aft body shock waves that interact with the exhaust plume contribute to the near-field pressure signature of a vehicle. The plume and shock interaction was studied using computational fluid dynamics and compared with experimental data from a coaxial convergent-divergent nozzle flow in an open jet facility. A simple diamond-shaped wedge was used to generate the shock in the outer flow to study its impact on the inner jet flow. Results show that the compression from the wedge deflects the nozzle plume and shocks form on the opposite plume boundary. The sonic boom pressure signature of the nozzle exhaust plume was modified by the presence of the wedge. Both the experimental results and computational predictions show changes in plume deflection.
    Keywords: Aerodynamics
    Type: E-663895 , AIAA SciTech 2014; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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  • 7
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    Dielectric Barrier Discharge (DBD) Plasma Actuators Thrust-Measurement Methodology Incorporating New Anti-Thrust Hypothesis (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a large diameter, grounded, metal sleeve.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0486 , E-18857-1 , AIAA SciTech 2014; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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  • 8
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    Validation of 3-D Ice Accretion Measurement Methodology for Experimental Aerodynamic Simulation (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel testing of artificial, or simulated, ice shapes. Recent developments in ice accretion documentation methods have yielded a laser-scanning capability that can measure highly three-dimensional features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing Research Tunnel using an 18-inch chord, 2-D straight wing with NACA 23012 airfoil section. For six ice accretion cases, a 3-D laser scan was performed to document the ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the University of Illinois low-speed wind tunnel at a Reynolds number of 1.8 x 10(exp 6) and a Mach number of 0.18 with an 18-inch chord NACA 23012 airfoil model that was designed to accommodate the artificial ice shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For four of the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil maximum lift coefficient were less than 3% with corresponding differences in stall angle of approximately one degree or less. The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice accretion for a NACA 23012 airfoil. For several of the ice-accretion cases tested, the aerodynamics is known to depend upon the small, three dimensional features of the ice. These data show that the laser-scan and rapid-prototype manufacturing approach is capable of replicating these ice features within the reported accuracies of the laser-scan measurement and rapid-prototyping method; thus providing a new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing research.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN14961 , AIAA Atmospheric and Space Environments Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 9
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    Contra-Rotating Open Rotor Tone Noise Prediction (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Reliable prediction of contra-rotating open rotor (CROR) noise is an essential element of any strategy for the development of low-noise open rotor propulsion systems that can meet both the community noise regulations and the cabin noise limits. Since CROR noise spectra typically exhibits a preponderance of tones, significant efforts have been directed towards predicting their tone spectra. To that end, there has been an ongoing effort at NASA to assess various in-house open rotor tone noise prediction tools using a benchmark CROR blade set for which significant aerodynamic and acoustic data had been acquired in wind tunnel tests. In the work presented here, the focus is on the near-field noise of the benchmark open rotor blade set at the cruise condition. Using an analytical CROR tone noise model with input from high-fidelity aerodynamic simulations, detailed tone noise spectral predictions have been generated and compared with the experimental data. Comparisons indicate that the theoretical predictions are in good agreement with the data, especially for the dominant CROR tones and their overall sound pressure level. The results also indicate that, whereas individual rotor tones are well predicted by the linear sources (i.e., thickness and loading), for the interaction tones it is essential that the quadrupole sources be included in the analysis.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN15607 , AIAA/CEAS Aeroacoustics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 10
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    Trim and Structural Optimization of Subsonic Transport Wings Using Nonconventional Aeroelastic Tailoring (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Several minimum-mass aeroelastic optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic strength and panel buckling constraints are imposed across a variety of trimmed maneuver loads. Tailoring with metallic thickness variations, functionally graded materials, composite laminates, tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
    Keywords: Aerodynamics
    Type: NF1676L-17757 , AIAA Aviation and Aeronautics Forum and Exposition (AVIATION 2014); Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 11
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    Steady and Unsteady Nozzle Simulations Using the Conservation Element and Solution Element Method (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: This paper presents results from computational fluid dynamic (CFD) simulations of a three-stream plug nozzle. Time-accurate, Euler, quasi-1D and 2D-axisymmetric simulations were performed as part of an effort to provide a CFD-based approach to modeling nozzle dynamics. The CFD code used for the simulations is based on the space-time Conservation Element and Solution Element (CESE) method. Steady-state results were validated using the Wind-US code and a code utilizing the MacCormack method while the unsteady results were partially validated via an aeroacoustic benchmark problem. The CESE steady-state flow field solutions showed excellent agreement with solutions derived from the other methods and codes while preliminary unsteady results for the three-stream plug nozzle are also shown. Additionally, a study was performed to explore the sensitivity of gross thrust computations to the control surface definition. The results showed that most of the sensitivity while computing the gross thrust is attributed to the control surface stencil resolution and choice of stencil end points and not to the control surface definition itself.Finally, comparisons between the quasi-1D and 2D-axisymetric solutions were performed in order to gain insight on whether a quasi-1D solution can capture the steady and unsteady nozzle phenomena without the cost of a 2D-axisymmetric simulation. Initial results show that while the quasi-1D solutions are similar to the 2D-axisymmetric solutions, the inability of the quasi-1D simulations to predict two dimensional phenomena limits its accuracy.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN16144 , AIAA Joint Propulsion Conference; Jul 28, 2014 - Jul 30, 2014; Cleveland, OH; United States
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  • 12
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    The Origin of Inlet Buzz in a Mach 1.7 Low Boom Inlet Design (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Supersonic inlets with external compression, having a good level performance at the critical operating point, exhibit a marked instability of the flow in some subcritical operation below a critical value of the capture mass flow ratio. This takes the form of severe oscillations of the shock system, commonly known as "buzz". The underlying purpose of this study is to indicate how Detached Eddy Simulation (DES) analysis of supersonic inlets will alter how we envision unsteady inlet aerodynamics, particularly inlet buzz. Presented in this paper is a discussion regarding the physical explanation underlying inlet buzz as indicated by DES analysis. It is the normal shock wave boundary layer separation along the spike surface which reduces the capture mass flow that is the controlling mechanism which determines the onset of inlet buzz, and it is the aerodynamic characteristics of a choked nozzle that provide the feedback mechanism that sustains the buzz cycle by imposing a fixed mean corrected inlet weight flow. Comparisons between the DES analysis of the Lockheed Martin Corporation (LMCO) N+2 inlet and schlieren photographs taken during the test of the Gulfstream Large Scale Low Boom (LSLB) inlet in the NASA 8x6 ft. Supersonic Wind Tunnel (SWT) show a strong similarity both in turbulent flow field structure and shock wave formation during the buzz cycle. This demonstrates the value of DES analysis for the design and understanding of supersonic inlets.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN15470 , Joint Propulsion Conference; Jul 28, 2014 - Jul 30, 2014; Cleveland, OH; United States
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  • 13
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    Thrust Removal Scheme for the FAST-MAC Circulation Control Model Tested in the National Transonic Facility (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.
    Keywords: Aerodynamics
    Type: AIAA Paper-2014-2402 , NF1676L-17629 , AIAA Applied Aerodynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA.; United States
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  • 14
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    Evaluation of Linear, Inviscid, Viscous, and Reduced-Order Modeling Aeroelastic Solutions of the AGARD 445.6 Wing Using Root Locus Analysis (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Reduced-order modeling (ROM) methods are applied to the CFD-based aeroelastic analysis of the AGARD 445.6 wing in order to gain insight regarding well-known discrepancies between the aeroelastic analyses and the experimental results. The results presented include aeroelastic solutions using the inviscid CAP-TSD code and the FUN3D code (Euler and Navier-Stokes). Full CFD aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. Important conclusions are drawn from these results including the ability of the linear CAP-TSD code to accurately predict the entire experimental flutter boundary (repeat of analyses performed in the 1980's), that the Euler solutions at supersonic conditions indicate that the third mode is always unstable, and that the FUN3D Navier-Stokes solutions stabilize the unstable third mode seen in the Euler solutions.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0496 , NF1676-16636 , AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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  • 15
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    Nonlinear Aeroelastic Analysis of the HIAD TPS Coupon in the NASA 8' High Temperature Tunnel: Theory and Experiment (2014)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: The purpose of this work is to develop a set of theoretical and experimental techniques to characterize the aeroelasticity of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). A square TPS coupon experiences trailing edge oscillatory behavior during experimental testing in the 8' High Temperature Tunnel (HTT), which may indicate the presence of aeroelastic flutter. Several theoretical aeroelastic models have been developed, each corresponding to a different experimental test configuration. Von Karman large deflection theory is used for the plate-like components of the TPS, along with piston theory for the aerodynamics. The constraints between the individual TPS layers and the presence of a unidirectional foundation at the back of the coupon are included by developing the necessary energy expressions and using the Rayleigh Ritz method to derive the nonlinear equations of motion. Free vibrations and limit cycle oscillations are computed and the frequencies and amplitudes are compared with accelerometer and photogrammetry data from the experiments.
    Keywords: Aerodynamics
    Type: NASA/TM-2014-218267 , L-20356 , NF1676L-18077
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  • 16
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    AFC-Enabled Vertical Tail System Integration Study (2014)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: This document serves as the final report for the SMAAART AFC-Enabled Vertical Tail System Integration Study. Included are the ground rule assumptions which have gone into the study, layouts of the baseline and AFC-enabled configurations, critical sizing information, system requirements and architectures, and assumed system properties that result in an NPV assessment of the two candidate AFC technologies.
    Keywords: Aerodynamics
    Type: NASA/CR-2014-218168 , NF1676L-18162
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  • 17
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    Analysis of Low Speed Stall Aerodynamics of a Swept Wing with Laminar Flow Glove (2014)
    In:  CASI
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    Publication Date: 2019-07-12
    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 American Institute of Aeronautics and Astronautics (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-TN13285
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  • 18
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    Computational Studies for Sonic Boom Prediction (2014)
    In:  CASI
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    Publication Date: 2019-07-20
    Description: Outline: Introduction; Launch Ascent and Vehicle Aerodynamics (LAVA) Framework; 1st AIAA Sonic Boom Prediction Workshop Oblique Shock/Plume Interaction; Summary.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN18916 , Applied Modeling and Simulation Seminar Series; Oct 30, 2014; Moffett Field, CA; United States
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  • 19
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    Delay of Transition Using Forced Damping (2014)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: Several experiments which have reported a delay of transition are analyzed in terms of the frequencies of the induced disturbances generated by different flow control elements. Two of the experiments employed passive stabilizers in the boundary layer, one leading-edge bluntness, and one employed an active spark discharge in the boundary layer. It is found that the frequencies generated by the various elements lie in the damping region of the associated stability curve. It is concluded that the creation of strong disturbances in the damping region stabilizes the boundary-layer and delays the transition from laminar to turbulent flow.
    Keywords: Aerodynamics
    Type: NASA/TM-2014-218551 , L-20496 , NF1676L-20143
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  • 20
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    Robust, Optimal Subsonic Airfoil Shapes (2014)
    In:  CASI
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    Publication Date: 2019-08-13
    Description: A method has been developed to create an airfoil robust enough to operate satisfactorily in different environments. This method determines a robust, optimal, subsonic airfoil shape, beginning with an arbitrary initial airfoil shape, and imposes the necessary constraints on the design. Also, this method is flexible and extendible to a larger class of requirements and changes in constraints imposed.
    Keywords: Aerodynamics
    Type: ARC-14586-2 , NASA Tech Briefs, January 2014; 22
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  • 21
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    Self-Contained Compressed-Flow Generation Device for Use in Making Differential Measurements (2014)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: A device used in making differential measurements of a flow includes a flow obstruction and a support arm. The flow obstruction's forward portion is a nose cone. The flow obstruction's aft portion is coupled to the nose cone. The support arm's first end is coupled to an exterior wall of a conduit, and its second end is coupled to the forward portion of the flow obstruction. The support arm positions the flow obstruction in the conduit such that a flow region is defined around its nose cone, and such that the support arm's first and second end are separated from one another with respect to a length dimension of the conduit. Measurement ports are provided in the support arm and flow obstruction. Manifolds extending through the flow obstruction and support arm couple the ports to points at the exterior wall of the conduit.
    Keywords: Aerodynamics
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    Multi-Element Airfoil System (2014)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: A multi-element airfoil system includes an airfoil element having a leading edge region and a skin element coupled to the airfoil element. A slat deployment system is coupled to the slat and the skin element, and is capable of deploying and retracting the slat and the skin element. The skin element substantially fills the lateral gap formed between the slat and the airfoil element when the slat is deployed. The system further includes an uncoupling device and a sensor to remove the skin element from the gap based on a critical angle-of-attack of the airfoil element. The system can alternatively comprise a trailing edge flap, where a skin element substantially fills the lateral gap between the flap and the trailing edge region of the airfoil element. In each case, the skin element fills a gap between the airfoil element and the deployed flap or slat to reduce airframe noise.
    Keywords: Aerodynamics
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    Advanced Modified High Performance Synthetic Jet Actuator with Curved Chamber (2014)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: The advanced modified high performance synthetic jet actuator with optimized curvature shape chamber (ASJA-M) is a synthetic jet actuator (SJA) with a lower volume reservoir or chamber. A curved chamber is used, instead of the conventional cylinder chamber, to reduce the dead volume of the jet chamber and increase the efficiency of the synthetic jet actuator. The shape of the curvature corresponds to the maximum displacement (deformation) profile of the electroactive diaphragm. The jet velocity and mass flow rate for the ASJA-M will be several times higher than conventional piezoelectric actuators.
    Keywords: Aerodynamics
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    Aeroelastic Tailoring via Tow Steered Composites (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The use of tow steered composites, where fibers follow prescribed curvilinear paths within a laminate, can improve upon existing capabilities related to aeroelastic tailoring of wing structures, though this tailoring method has received relatively little attention in the literature. This paper demonstrates the technique for both a simple cantilevered plate in low-speed flow, as well as the wing box of a full-scale high aspect ratio transport configuration. Static aeroelastic stresses and dynamic flutter boundaries are obtained for both cases. The impact of various tailoring choices upon the aeroelastic performance is quantified: curvilinear fiber steering versus straight fiber steering, certifiable versus noncertifiable stacking sequences, a single uniform laminate per wing skin versus multiple laminates, and identical upper and lower wing skins structures versus individual tailoring.
    Keywords: Aerodynamics
    Type: NASA/TM-2014-218517 , L-20443 , NF1676L-19338
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    Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: This is part 2 of a two part document. Part 1 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression." A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind tunnel model was mated to a new, two degree of freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at10 percent static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free flying wind tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.
    Keywords: Aerodynamics
    Type: NASA/TP-2014-218170 , L-20321 , NF1676L-17355
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    Computational Simulations of Convergent Nozzles for the AIAA 1st Propulsion Aerodynamics Workshop (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Computational Fluid Dynamics (CFD) simulations were completed for a series of convergent nozzles in participation of the American Institute of Aeronautics and Astronautics (AIAA) 1st Propulsion Aerodynamics Workshop. The simulations were performed using the Wind-US flow solver. Discharge and thrust coefficients were computed for four axisymmetric nozzles with nozzle pressure ratios (NPR) ranging from 1.4 to 7.0. The computed discharge coefficients showed excellent agreement with available experimental data; the computed thrust coefficients captured trends observed in the experimental data, but over-predicted the thrust coefficient by 0.25 to 1.0 percent. Sonic lines were computed for cases with NPR 〉= 2.0 and agreed well with experimental data for NPR 〉= 2.5. Simulations were also performed for a 25 deg. conic nozzle bifurcated by a flat plate at NPR = 4.0. The jet plume shock structure was compared with and without the splitter plate to the experimental data. The Wind-US simulations predicted the shock structure well, though lack of grid resolution in the plume reduced the sharpness of the shock waves. Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulations and Detached Eddy Simulations (DES) were performed at NPR = 1.6 for the 25 deg conic nozzle with splitter plate. The simulations predicted vortex shedding from the trailing edge of the splitter plate. However, the vortices of URANS and DES solutions appeared to dissipate earlier than observed experimentally. It is believed that a lack of grid resolution in the region of the vortex shedding may have caused the vortices to break down too soon
    Keywords: Aerodynamics
    Type: NASA/TM-2014-218329 , E-18926 , GRC-E-DAA-TN14457
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    AFC-Enabled Simplified High-Lift System Integration Study (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The primary objective of this trade study report is to explore the potential of using Active Flow Control (AFC) for achieving lighter and mechanically simpler high-lift systems for transonic commercial transport aircraft. This assessment was conducted in four steps. First, based on the Common Research Model (CRM) outer mold line (OML) definition, two high-lift concepts were developed. One concept, representative of current production-type commercial transonic transports, features leading edge slats and slotted trailing edge flaps with Fowler motion. The other CRM-based design relies on drooped leading edges and simply hinged trailing edge flaps for high-lift generation. The relative high-lift performance of these two high-lift CRM variants is established using Computational Fluid Dynamics (CFD) solutions to the Reynolds-Averaged Navier-Stokes (RANS) equations for steady flow. These CFD assessments identify the high-lift performance that needs to be recovered through AFC to have the CRM variant with the lighter and mechanically simpler high-lift system match the performance of the conventional high-lift system. Conceptual design integration studies for the AFC-enhanced high-lift systems were conducted with a NASA Environmentally Responsible Aircraft (ERA) reference configuration, the so-called ERA-0003 concept. These design trades identify AFC performance targets that need to be met to produce economically feasible ERA-0003-like concepts with lighter and mechanically simpler high-lift designs that match the performance of conventional high-lift systems. Finally, technical challenges are identified associated with the application of AFC-enabled highlift systems to modern transonic commercial transports for future technology maturation efforts.
    Keywords: Aerodynamics
    Type: NASA/CR-2014-218521 , NF1676L-19609
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    Aerodynamic Performance of a Compact, High Work-Factor Centrifugal Compressor at the Stage and Subcomponent Level (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The measured aerodynamic performance of a compact, high work factor, single-stage centrifugal compressor, comprising an impeller, diffuser, 90-bend, and exit guide vane (EGV), is reported. Performance levels are based on steady-state total-pressure and total-temperature rake and angularity-probe data acquired at key machine rating planes during recent testing at NASA Glenn Research Center. Aerodynamic performance at the stage level are reported for operation between 70 to 105 of design corrected speed, with subcomponent (impeller, diffuser, and exitguide-vane) detailed flow field measurements presented and discussed at the 100 design-speed condition. Individual component losses from measurements are compared with pre-test predictions on a limited basis.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN16677 , AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 28, 2014 - Jul 30, 2014; Cleveland, OH; United States
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    In-Flight Aeroelastic Stability of the Thermal Protection System on the NASA HIAD, Part I: Linear Theory (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Conical shell theory and piston theory aerodynamics are used to study the aeroelastic stability of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). Structural models of the TPS consist of single or multiple orthotropic conical shell systems resting on several circumferential linear elastic supports. The shells in each model may have pinned (simply-supported) or elastically-supported edges. The Lagrangian is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the equations of motion. The natural modes of vibration and aeroelastic stability boundaries are found by calculating the eigenvalues and eigenvectors of a large coefficient matrix. When the in-flight configuration of the TPS is approximated as a single shell without elastic supports, asymmetric flutter in many circumferential waves is observed. When the elastic supports are included, the shell flutters symmetrically in zero circumferential waves. Structural damping is found to be important in this case. Aeroelastic models that consider the individual TPS layers as separate shells tend to flutter asymmetrically at high dynamic pressures relative to the single shell models. Several parameter studies also examine the effects of tension, orthotropicity, and elastic support stiffness.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-1520 , NF1676L-17623 , AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Jan 13, 2014 - Jan 17, 2014; National Harbor, Md; United States
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    Uncertainty Quantification and Certification Prediction of Low-Boom Supersonic Aircraft Configurations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The primary objective of this work was to develop and demonstrate a process for accurate and efficient uncertainty quantification and certification prediction of low-boom, supersonic, transport aircraft. High-fidelity computational fluid dynamics models of multiple low-boom configurations were investigated including the Lockheed Martin SEEB-ALR body of revolution, the NASA 69 Delta Wing, and the Lockheed Martin 1021-01 configuration. A nonintrusive polynomial chaos surrogate modeling approach was used for reduced computational cost of propagating mixed, inherent (aleatory) and model-form (epistemic) uncertainty from both the computation fluid dynamics model and the near-field to ground level propagation model. A methodology has also been introduced to quantify the plausibility of a design to pass a certification under uncertainty. Results of this study include the analysis of each of the three configurations of interest under inviscid and fully turbulent flow assumptions. A comparison of the uncertainty outputs and sensitivity analyses between the configurations is also given. The results of this study illustrate the flexibility and robustness of the developed framework as a tool for uncertainty quantification and certification prediction of low-boom, supersonic aircraft.
    Keywords: Aerodynamics
    Type: NF1676L-18768 , AIAA Aviation 2014; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States|AIAA Applied Aerodynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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    Analysis of Numerical Simulation Database for Pressure Fluctuations Induced by High-Speed Turbulent Boundary Layers (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Direct numerical simulations (DNS) of Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Re(sub T) approximately 460 are conducted at two wall temperatures (Tw/Tr = 0.25, 0.76) to investigate the generated pressure fluctuations and their dependence on wall temperature. Simulations indicate that the influence of wall temperature on pressure fluctuations is largely limited to the near-wall region, with the characteristics of wall-pressure fluctuations showing a strong temperature dependence. Wall temperature has little influence on the propagation speed of the freestream pressure signal. The freestream radiation intensity compares well between wall-temperature cases when normalized by the local wall shear; the propagation speed of the freestream pressure signal and the orientation of the radiation wave front show little dependence on the wall temperature.
    Keywords: Aerodynamics
    Type: AIAA Paper-2014-2912 , NF1676L-17593 , AIAA Fluid Dynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States|AIAA Aviation and Aeronautics Forum and Exposition (AVIATION 2014); Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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    Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade (2014)
    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 50 deg 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/REV1 , AIAA-2012-3879 , E-18746-1 , Joint Propulsion Conference and Exhibit; Jul 29, 2012 - Aug 01, 2012; Atlanta, GA; United States
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    Unsteady Aerodynamic Validation Experiences From the Aeroelastic Prediction Workshop (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The AIAA Aeroelastic Prediction Workshop (AePW) was held in April 2012, bringing together communities of aeroelasticians, computational fluid dynamicists and experimentalists. The extended objective was to assess the state of the art in computational aeroelastic methods as practical tools for the prediction of static and dynamic aeroelastic phenomena. As a step in this process, workshop participants analyzed unsteady aerodynamic and weakly-coupled aeroelastic cases. Forced oscillation and unforced system experiments and computations have been compared for three configurations. This paper emphasizes interpretation of the experimental data, computational results and their comparisons from the perspective of validation of unsteady system predictions. The issues examined in detail are variability introduced by input choices for the computations, post-processing, and static aeroelastic modeling. The final issue addressed is interpreting unsteady information that is present in experimental data that is assumed to be steady, and the resulting consequences on the comparison data sets.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0203 , NF1676L-16688 , AIAA Aerospace Sciences Meeting; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    Roll Damping Derivatives from Generalized Lifting-Surface Theory and Wind Tunnel Forced-Oscillation Tests (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Improving aerodynamic models for adverse loss-of-control conditions in flight is an area being researched under the NASA Aviation Safety Program. Aerodynamic models appropriate for loss of control conditions require a more general mathematical representation to predict nonlinear unsteady behaviors. As more general aerodynamic models are studied that include nonlinear higher order effects, the possibility of measurements that confound aerodynamic and structural responses are probable. In this study an initial step is taken to look at including structural flexibility in analysis of rigid-body forced-oscillation testing that accounts for dynamic rig, sting and balance flexibility. Because of the significant testing required and associated costs in a general study, it makes sense to capitalize on low cost analytical methods where possible, especially where structural flexibility can be accounted for by a low cost method. This paper provides an initial look at using linear lifting surface theory applied to rigid-body aircraft roll forced-oscillation tests.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0731 , NF1676L-16685 , AIAA Science and Technology Forum and Exposition (SciTech2014); Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    Blade Displacement Predictions for the Full-Scale UH-60A Airloads Rotor (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An unsteady Reynolds-Averaged Navier-Stokes solver for unstructured grids is loosely coupled to a rotorcraft comprehensive code and used to simulate two different test conditions from a wind-tunnel test of a full-scale UH-60A rotor. Performance data and sectional airloads from the simulation are compared with corresponding tunnel data to assess the level of fidelity of the aerodynamic aspects of the simulation. The focus then turns to a comparison of the blade displacements, both rigid (blade root) and elastic. Comparisons of computed root motions are made with data from three independent measurement systems. Finally, comparisons are made between computed elastic bending and elastic twist, and the corresponding measurements obtained from a photogrammetry system. Overall the correlation between computed and measured displacements was good, especially for the root pitch and lag motions and the elastic bending deformation. The correlation of root lead-lag motion and elastic twist deformation was less favorable.
    Keywords: Aerodynamics
    Type: NF1676L-17750 , American Helicopter Society (AHS) Annual Forum; May 20, 2014 - May 22, 2014; Montreal, Quebec; Canada
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    Flight Tests of a Supersonic Natural Laminar Flow Airfoil (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A flight test campaign of a supersonic natural laminar flow airfoil has been recently completed. The test surface was an 80-inch (203 cm) chord and 40-inch (102 cm) span article mounted on the centerline store location of an F-15B airplane. The wing was designed with a leading edge sweep of effectively 0 deg to minimize boundary layer crossflow. The test article surface was coated with an insulating material to avoid significant heat transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared camera system was used to determine boundary layer transition and the extent of laminar flow. The tests were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of the flow to disturbances. Both discrete (trip dots) and 2-D disturbances (forward-facing steps) were tested. A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results indicate the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, were similar to that of subsonic natural laminar flow wings.
    Keywords: Aerodynamics
    Type: DFRC-E-DAA-TN14087 , International Symposium on Flow Visualization; Jun 24, 2014 - Jun 28, 2014; Okinawa; Japan
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    Model-Invariant Hybrid LES-RANS Computation of Separated Flow Past Periodic Hills (2014)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The requirement that physical quantities not vary with a hybrid LESRANS model's blending parameter imposes conditions on the computation that lead to better results across LES-RANS transitions. This promises to allow placement of those transitions so that LES is performed only where required by the physics, improving computational efficiency. The approach is applied to separated flow past periodic hills, where good predictions of separation-bubble size are seen due to the gradual, controlled, LES-RANS transition and the resulting enhanced near-wall eddy viscosity.
    Keywords: Aerodynamics
    Type: NF1676L-17795 , Symposium on Hybrid RANS-LES Methods; Mar 19, 2014 - Mar 21, 2014; College Station, TX; United States
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    Experimental Optimization of a Free-to-Rotate Wing for Small UAS (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper discusses an experimental investigation conducted to optimize a free-to-rotate wing for use on a small unmanned aircraft system (UAS). Although free-to-rotate wings have been used for decades on various small UAS and small manned aircraft, little is known about how to optimize these unusual wings for a specific application. The paper discusses some of the design rationale of the basic wing. In addition, three main parameters were selected for "optimization", wing camber, wing pivot location, and wing center of gravity (c.g.) location. A small apparatus was constructed to enable some simple experimental analysis of these parameters. A design-of-experiment series of tests were first conducted to discern which of the main optimization parameters were most likely to have the greatest impact on the outputs of interest, namely, some measure of "stability", some measure of the lift being generated at the neutral position, and how quickly the wing "recovers" from an upset. A second set of tests were conducted to develop a response-surface numerical representation of these outputs as functions of the three primary inputs. The response surface numerical representations are then used to develop an "optimum" within the trade space investigated. The results of the optimization are then tested experimentally to validate the predictions.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-2273 , NF1676L-17724 , AIAA Applied Aerodynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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    Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to 51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.1210(exp 5) to 2.1210(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.
    Keywords: Aerodynamics
    Type: E-18952 , AIAA Joint Propulsion Conference; Jul 28, 2014 - Jul 30, 2014; Cleveland, OH; United States
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    Tone and Broadband Noise Separation from Acoustic Data of a Scale-Model Contra-Rotating Open Rotor (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Renewed interest in contra-rotating open rotor technology for aircraft propulsion application has prompted the development of advanced diagnostic tools for better design and improved acoustical performance. In particular, the determination of tonal and broadband components of open rotor acoustic spectra is essential for properly assessing the noise control parameters and also for validating the open rotor noise simulation codes. The technique of phase averaging has been employed to separate the tone and broadband components from a single rotor, but this method does not work for the two-shaft contra-rotating open rotor. A new signal processing technique was recently developed to process the contra-rotating open rotor acoustic data. The technique was first tested using acoustic data taken of a hobby aircraft open rotor propeller, and reported previously. The intent of the present work is to verify and validate the applicability of the new technique to a realistic one-fifth scale open rotor model which has 12 forward and 10 aft contra-rotating blades operating at realistic forward flight Mach numbers and tip speeds. The results and discussions of that study are presented in this paper.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN15182 , AIAA Aeroacoustics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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    Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade With Large Incidence and Reynolds Number Variations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to 51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.1210(exp 5) to 2.1210(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.
    Keywords: Aerodynamics
    Type: NASA/TM-2014-218137 , E-18952-1 , AIAA Joint Propulsion Conference; Jul 28, 2014 - Jul 30, 2014; Cleveland, OH; United States
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    Wind Tunnel Testing of a 120th Scale Large Civil Tilt-Rotor Model in Airplane and Helicopter Modes (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In April 2012 and October 2013, NASA and the U.S. Army jointly conducted a wind tunnel test program examining two notional large tilt rotor designs: NASA's Large Civil Tilt Rotor and the Army's High Efficiency Tilt Rotor. The approximately 6%-scale airframe models (unpowered) were tested without rotors in the U.S. Army 7- by 10-foot wind tunnel at NASA Ames Research Center. Measurements of all six forces and moments acting on the airframe were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography and oil flow were used to identify flow trajectories, boundary layer transition and areas of flow separation. The purpose of this test was to collect data for the validation of computational fluid dynamics tools, for the development of flight dynamics simulation models, and to validate performance predictions made during conceptual design. This paper focuses on the results for the Large Civil Tilt Rotor model in an airplane mode configuration up to 200 knots of wind tunnel speed. Results are presented with the full airframe model with various wing tip and nacelle configurations, and for a wing-only case also with various wing tip and nacelle configurations. Key results show that the addition of a wing extension outboard of the nacelles produces a significant increase in the lift-to-drag ratio, and interestingly decreases the drag compared to the case where the wing extension is not present. The drag decrease is likely due to complex aerodynamic interactions between the nacelle and wing extension that results in a significant drag benefit.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN12608 , Decennial AHS Aeromechanics Specialists'' Conference; Jan 22, 2014 - Jan 24, 2014; San Francisco, CA; United States
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    Experimental Study of Slat Noise from 30P30N Three-Element High-Lift Airfoil in JAXA Hard-Wall Low-Speed Wind Tunnel (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Aeroacoustic measurements associated with noise radiation from the leading edge slat of the canonical, unswept 30P30N three-element high-lift airfoil configuration have been obtained in a 2 m x 2 m hard-wall wind tunnel at the Japan Aerospace Exploration Agency (JAXA). Performed as part of a collaborative effort on airframe noise between JAXA and the National Aeronautics and Space Administration (NASA), the model geometry and majority of instrumentation details are identical to a NASA model with the exception of a larger span. For an angle of attack up to 10 degrees, the mean surface Cp distributions agree well with free-air computational fluid dynamics predictions corresponding to a corrected angle of attack. After employing suitable acoustic treatment for the brackets and end-wall effects, an approximately 2D noise source map is obtained from microphone array measurements, thus supporting the feasibility of generating a measurement database that can be used for comparison with free-air numerical simulations. Both surface pressure spectra obtained via KuliteTM transducers and the acoustic spectra derived from microphone array measurements display a mixture of a broad band component and narrow-band peaks (NBPs), both of which are most intense at the lower angles of attack and become progressively weaker as the angle of attack is increased. The NBPs exhibit a substantially higher spanwise coherence in comparison to the broadband portion of the spectrum and, hence, confirm the trends observed in previous numerical simulations. Somewhat surprisingly, measurements show that the presence of trip dots between the stagnation point and slat cusp enhances the NBP levels rather than mitigating them as found in a previous experiment.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-2080 , NF1676L-18887 , AIAA Aviation and Aeronautics Forum and Exposition (AVIATION 2014); Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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    Thrust Measurement of Dielectric Barrier Discharge (DBD) Plasma Actuators: New Anti-Thrust Hypothesis, Frequency Sweeps Methodology, Humidity and Enclosure Effects (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust, or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a grounded large-diameter metal sleeve. Strong dependence on humidity is also shown; the thrust significantly increased with decreasing humidity, e.g., 44 percent increase as relative humidity changed from 18 percent and dew point 33 degF to 50 percent and dew point of 57 degF.
    Keywords: Aerodynamics
    Type: NASA/TM2014-218115 , AIAA Paper 2014-0486 , E-18857 , SciTech 2014 Aerospace Sciences Meeting; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    Static Aeroelastic and Longitudinal Trim Model of Flexible Wing Aircraft Using Finite-Element Vortex-Lattice Coupled Solution (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a static aeroelastic model and longitudinal trim model for the analysis of a flexible wing transport aircraft. The static aeroelastic model is built using a structural model based on finite-element modeling and coupled to an aerodynamic model that uses vortex-lattice solution. An automatic geometry generation tool is used to close the loop between the structural and aerodynamic models. The aeroelastic model is extended for the development of a three degree-of-freedom longitudinal trim model for an aircraft with flexible wings. The resulting flexible aircraft longitudinal trim model is used to simultaneously compute the static aeroelastic shape for the aircraft model and the longitudinal state inputs to maintain an aircraft trim state. The framework is applied to an aircraft model based on the NASA Generic Transport Model (GTM) with wing structures allowed to flexibly deformed referred to as the Elastically Shaped Aircraft Concept (ESAC). The ESAC wing mass and stiffness properties are based on a baseline "stiff" values representative of current generation transport aircraft.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN12487 , SciTech 2014; Jan 13, 2014 - Jan 17, 2014; Baltimore, MD; United States
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    Maintaining Aura's Orbit Requirements Under New Maneuver Operations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Earth Observing System (EOS) Afternoon Constellation consists of five member missions (GCOM-W1, Aqua, CALIPSO, CloudSat, and Aura), each of which maintain a frozen, sun-synchronous orbit with a 16-day repeating ground track that follows the Worldwide Reference System-2 (WRS-2). Under nominal science operations for Aura, the propulsion system is oriented such that the resultant thrust vector is aligned 13.493 degrees away from the velocity vector along the yaw axis. When performing orbit maintenance maneuvers, the spacecraft performs a yaw slew to align the thrust vector in the appropriate direction. A new Drag Make Up (DMU) maneuver operations scheme has been implemented for Aura alleviating the need for the 13.493 degree yaw slew. The focus of this investigation is to assess the impact that no-slew DMU maneuver operations will have on Auras Mean Local Time (MLT) which drives the required along track separation between Aura and the constellation members, as well as Auras frozen orbit properties, eccentricity and argument of perigee. Seven maneuver strategies were analyzed to determine the best operational approach. A mirror pole strategy, with maneuvers alternating at the North and South poles, was implemented operationally to minimize impact to the MLT. Additional analysis determined that the mirror pole strategy could be further modified to include frozen orbit maneuvers and thus maintain both MLT and the frozen orbit properties under no-slew operations
    Keywords: Aerodynamics
    Type: GSFC-E-DAA-TN14703 , International Symposium on Space Flight Dynamics; May 05, 2014 - May 09, 2014; Laurel, MD; United States
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    Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Several topology optimization problems are conducted within the ribs and spars of a wing box. It is desired to locate the best position of lightening holes, truss/cross-bracing, etc. A variety of aeroelastic metrics are isolated for each of these problems: elastic wing compliance under trim loads and taxi loads, stress distribution, and crushing loads. Aileron effectiveness under a constant roll rate is considered, as are dynamic metrics: natural vibration frequency and flutter. This approach helps uncover the relationship between topology and aeroelasticity in subsonic transport wings, and can therefore aid in understanding the complex aircraft design process which must eventually consider all these metrics and load cases simultaneously.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0633 , NF1676L-16611 , AIAA Multidisciplinary Design Optimization Specialist Conference; Jan 13, 2014 - Jan 17, 2014; National Harbor,MD; United States
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    Time-Spectral Rotorcraft Simulations on Overset Grids (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Time-Spectral method is derived as a Fourier collocation scheme and applied to NASA's overset Reynolds-averaged Navier-Stokes (RANS) solver OVERFLOW. The paper outlines the Time-Spectral OVERFLOWimplementation. Successful low-speed laminar plunging NACA 0012 airfoil simulations demonstrate the capability of the Time-Spectral method to resolve the highly-vortical wakes typical of more expensive three-dimensional rotorcraft configurations. Dealiasing, in the form of spectral vanishing viscosity (SVV), facilitates the convergence of Time-Spectral calculations of high-frequency flows. Finally, simulations of the isolated V-22 Osprey tiltrotor for both hover and forward (edgewise) flight validate the three-dimensional Time-Spectral OVERFLOW implementation. The Time-Spectral hover simulation matches the time-accurate calculation using a single harmonic. Significantly more temporal modes and SVV are required to accurately compute the forward flight case because of its more active, high-frequency wake.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN15492 , Applied Aerodynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 49
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    Assessment of Higher-Order RANS Closures in a Decelerated Planar Wall-Bounded Turbulent Flow (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A reference DNS database is presented, which includes third- and fourth-order moment budgets for unstrained and strained planar channel flow. Existing RANS closure models for third- and fourth-order terms are surveyed, and new model ideas are introduced. The various models are then compared with the DNS data term by term using a priori testing of the higher-order budgets of turbulence transport, velocity-pressure-gradient, and dissipation for both the unstrained and strained databases. Generally, the models for the velocity-pressure-gradient terms are most in need of improvement.
    Keywords: Aerodynamics
    Type: NF1676L-17686 , (AVIATION 2014) AIAA Aviation and Aeronautics Forum and Exposition; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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  • 50
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    Aerodynamic Effects of Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.
    Keywords: Aerodynamics
    Type: GRC-E-DAA-TN17526 , Turbine Engine Technical Symposium 2014; Sep 08, 2014 - Sep 11, 2014; Dayton, OH; United States
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    Numerical Simulations For the F-16XL Aircraft Configuration (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Numerical simulations of flow around the F-16XL are presented as a contribution to the Cranked Arrow Wing Aerodynamic Project International II (CAWAPI-II). The NASA Tetrahedral Unstructured Software System (TetrUSS) is used to perform numerical simulations. This CFD suite, developed and maintained by NASA Langley Research Center, includes an unstructured grid generation program called VGRID, a postprocessor named POSTGRID, and the flow solver USM3D. The CRISP CFD package is utilized to provide error estimates and grid adaption for verification of USM3D results. A subsonic high angle-of-attack case flight condition (FC) 25 is computed and analyzed. Three turbulence models are used in the calculations: the one-equation Spalart-Allmaras (SA), the two-equation shear stress transport (SST) and the k turbulence models. Computational results, and surface static pressure profiles are presented and compared with flight data. Solution verification is performed using formal grid refinement studies, the solution of Error Transport Equations, and adaptive mesh refinement. The current study shows that the USM3D solver coupled with CRISP CFD can be used in an engineering environment in predicting vortex-flow physics on a complex configuration at flight Reynolds numbers.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0756 , NF1676L-16701 , AIAA Aerospace Sciences Meeting; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    The NASA High Speed ASE Project: Computational Analyses of a Low-Boom Supersonic Configuration (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A summary of NASA's High Speed Aeroservoelasticity (ASE) project is provided with a focus on a low-boom supersonic configuration developed by Lockheed-Martin and referred to as the N+2 configuration. The summary includes details of the computational models developed to date including a linear finite element model (FEM), linear unsteady aerodynamic models, structured and unstructured CFD grids, and discussion of the FEM development including sizing and structural constraints applied to the N+2 configuration. Linear results obtained to date include linear mode shapes and linear flutter boundaries. In addition to the tasks associated with the N+2 configuration, a summary of the work involving the development of AeroPropulsoServoElasticity (APSE) models is also discussed.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0675 , NF1676L-16637 , AIAA Aerospace Sciences Meeting; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    A Summary of Revisions Applied to a Turbulence Response Analysis Method for Flexible Aircraft Configurations (2014)
    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 additional output data so as to provide a more useful and precise tool for gust load analysis. 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 is included. These revisions and enhancements are implemented and an analysis of the results is used to validate the modifications.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-0600 , NF1676L-16647 , AIAA Guidance, Navigation, and Control Conference; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    An Experimntal Investigation of the 30P30N Multi-Element High-Lift Airfoil (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: High-lift devices often generate an unsteady flow field producing both broadband and tonal noise which radiates from the aircraft. In particular, the leading edge slat is often a dominant contributor to the noise signature. An experimental study of a simplified unswept high-lift configuration, the 30P30N, has been conducted to understand and identify the various flow-induced noise sources around the slat. Closed-wall wind tunnel tests are performed in the Florida State Aeroacoustic Tunnel (FSAT) to characterize the slat cove flow field using a combination of surface and off-body measurements. Mean surface pressures compare well with numerical predictions for the free-air configuration. Consistent with previous measurements and computations for 2D high-lift configurations, the frequency spectra of unsteady surface pressures on the slat surface display several narrowband peaks that decrease in strength as the angle of attack is increased. At positive angles of attack, there are four prominent peaks. The three higher frequency peaks correspond, approximately, to a harmonic sequence related to a feedback resonance involving unstable disturbances in the slat cove shear layer. The Strouhal numbers associated with these three peaks are nearly insensitive to the range of flow speeds (41-58 m/s) and the angles of attack tested (3-8.5 degrees). The first narrow-band peak has an order of magnitude lower frequency than the remaining peaks and displays noticeable sensitivity to the angle of attack. Stereoscopic particle image velocimetry (SPIV) measurements provide supplementary information about the shear layer characteristics and turbulence statistics that may be used for validating numerical simulations.
    Keywords: Aerodynamics
    Type: AIAA Paper 2014-3062 , NF1676L-18896 , AIAA Aviation and Aeronautics Forum and Exposition (AVIATION 2014); Jun 16, 2014 - Jun 20, 2014; Atlanta, GA; United States
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    Space Launch System Ascent Static Aerodynamic Database Development (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes the wind tunnel testing work and data analysis required to characterize the static aerodynamic environment of NASA's Space Launch System (SLS) ascent portion of flight. Scaled models of the SLS have been tested in transonic and supersonic wind tunnels to gather the high fidelity data that is used to build aerodynamic databases. A detailed description of the wind tunnel test that was conducted to produce the latest version of the database is presented, and a representative set of aerodynamic data is shown. The wind tunnel data quality remains very high, however some concerns with wall interference effects through transonic Mach numbers are also discussed. Post-processing and analysis of the wind tunnel dataset are crucial for the development of a formal ascent aerodynamics database.
    Keywords: Aerodynamics
    Type: AIAA Paper-2014-1254 , NF1676L-17910 , AIAA Aerospace Sciences Meeting; Jan 13, 2014 - Jan 17, 2014; National Harbor, MD; United States
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    Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The purpose of this thesis is to document the impact of incidence angle and Reynolds number variations on the three-dimensional flow field and midspan loss and turning of a two-dimensional section of a variable-speed power-turbine (VSPT) rotor blade. Aerodynamic measurements were obtained in a transonic linear cascade at NASA Glenn Research Center in Cleveland, Ohio. Steady-state data were obtained for 10 incidence angles ranging from +15.8deg to -51.0deg. 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 2.12105 to 2.12106. 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 tota lpressure 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 +5.8deg and -36.7deg. 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.
    Keywords: Aerodynamics
    Type: NASA/TM-2014-218133 , E-18929
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    A Modular Approach to Model Oscillating Control Surfaces Using Navier Stokes Equations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-19
    Description: The use of active controls for rotorcraft is becoming more important for modern aerospace configurations. Efforts to reduce the vibrations of helicopter blades with use of active-controls are in progress. Modeling oscillating control surfaces using the linear aerodynamics theory is well established. However, higher-fidelity methods are needed to account for nonlinear effects, such as those that occur in transonic flow. The aeroelastic responses of a wing with an oscillating control surface, computed using the transonic small perturbation (TSP) theory, have been shown to cause important transonic flow effects such as a reversal of control surface effectiveness that occurs as the shock wave crosses the hinge line. In order to account for flow complexities such as blade-vortex interactions of rotor blades higher-fidelity methods based on the Navier-Stokes equations are used. Reference 6 presents a procedure that uses the Navier-Stokes equations with moving-sheared grids and demonstrates up to 8 degrees of control-surface amplitude, using a single grid. Later, this procedure was extended to accommodate larger amplitudes, based on sliding grid zones. The sheared grid method implemented in EulerlNavier-Stokes-based aeroelastic code ENS AERO was successfully applied to active control design by industry. Recently there are several papers that present results for oscillating control surface using Reynolds Averaged Navier-Stokes (RANS) equations. References 9 and 10 report 2-D cases by filling gaps with overset grids. Reference 9 compares integrated forces with the experiment at low oscillating frequencies whereas Ref. 10 reports parametric studies but with no validation. Reference II reports results for a 3D case by modeling the gap region with a deformed grid and compares force results with the experiment only at the mid-span of flap. In Ref. II grid is deformed to match the control surface deflections at the section where the measurements are made. However, there is no indication in Ref. II that the gaps are explicitly modeled as in Ref. 6. Computations using overset grids are reported in Ref. 12 for a case by adding moving control surface to an existing blade but with no validation either with an experiment or another computation.
    Keywords: Aerodynamics
    Type: ARC-E-DAA-TN12352 , Decennial AHS Aeromechanics Specialists Meeting; Jan 22, 2014 - Jan 24, 2014; San Francisco, CA; United States
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