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  • 1
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    Structural Dynamics Modeling of HIRENASD in Support of the Aeroelastic Prediction Workshop (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: An Aeroelastic Prediction Workshop (AePW) was held in April 2012 using three aeroelasticity case study wind tunnel tests for assessing the capabilities of various codes in making aeroelasticity predictions. One of these case studies was known as the HIRENASD model that was tested in the European Transonic Wind Tunnel (ETW). This paper summarizes the development of a standardized enhanced analytical HIRENASD structural model for use in the AePW effort. The modifications to the HIRENASD finite element model were validated by comparing modal frequencies, evaluating modal assurance criteria, comparing leading edge, trailing edge and twist of the wing with experiment and by performing steady and unsteady CFD analyses for one of the test conditions on the same grid, and identical processing of results.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-1801 , NF1676L-15290 , 54th AIAA/ASME/ASCE/AHS/ASC, Structures, Structural Dynamics, and Materials Conference; 8-1` Apr. 2013; Boston, MA; United States
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  • 2
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    DRE-Enhanced Swept-Wing Natural Laminar Flow at High Reynolds Numbers (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: Nonlinear parabolized stability equations and secondary instability analyses are used to provide a computational assessment of the potential use of the discrete roughness elements (DRE) technology for extending swept-wing natural laminar flow at chord Reynolds numbers relevant to transport aircraft. Computations performed for the boundary layer on a natural laminar flow airfoil with a leading-edge sweep angle of 34.6deg, free-stream Mach number of 0.75 and chord Reynolds numbers of 17 x 10(exp 6), 24 x 10(exp 6) and 30 x 10(exp 6) suggest that DRE could delay laminar-turbulent transition by about 20% when transition is caused by stationary crossflow disturbances. Computations show that the introduction of small wavelength stationary crossflow disturbances (i.e., DRE) also suppresses the growth of most amplified traveling crossflow disturbances.
    Keywords: Aerodynamics
    Type: AIAA Paper 2013-0412 , NF1676L-14855 , 51st AIAA Aerospace Sciences Meeting; 7-110 Jan. 2013; Grapevine, TX; United States
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  • 3
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    Aerodynamic Classification of Swept-Wing Ice Accretion (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely on the advancement of computational and experimental simulation methods for higher fidelity results over an increasing range of aircraft configurations and performance, and icing conditions. The current stateof- the-art in icing aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries that does not currently exist for fundamentally three-dimensional geometries such as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics and to identify the type of research that is required to improve the current understanding. Following the method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: ice roughness, horn ice, streamwise ice and spanwise-ridge ice. In the case of horn ice it is shown that a further subclassification of "nominally 3D" or "highly 3D" horn ice may be necessary. For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data presented in this report can serve as a starting point as new research explores swept-wing aerodynamics with ice shapes. As further results are available, it is expected that these classifications will need to be updated and revised.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-216556 , E-18753 , GRC-E-DAA-TN9727 , AIAA Atmospheric and Space Enviroments Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 4
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    Swept-Wing Ice Accretion Characterization and Aerodynamics (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA, FAA, ONERA, the University of Illinois and Boeing have embarked on a significant, collaborative research effort to address the technical challenges associated with icing on large-scale, three-dimensional swept wings. The overall goal is to improve the fidelity of experimental and computational simulation methods for swept-wing ice accretion formation and resulting aerodynamic effect. A seven-phase research effort has been designed that incorporates ice-accretion and aerodynamic experiments and computational simulations. As the baseline, full-scale, swept-wing-reference geometry, this research will utilize the 65% scale Common Research Model configuration. Ice-accretion testing will be conducted in the NASA Icing Research Tunnel for three hybrid swept-wing models representing the 20%, 64% and 83% semispan stations of the baseline-reference wing. Three-dimensional measurement techniques are being developed and validated to document the experimental ice-accretion geometries. Artificial ice shapes of varying geometric fidelity will be developed for aerodynamic testing over a large Reynolds number range in the ONERA F1 pressurized wind tunnel and in a smaller-scale atmospheric wind tunnel. Concurrent research will be conducted to explore and further develop the use of computational simulation tools for ice accretion and aerodynamics on swept wings. The combined results of this research effort will result in an improved understanding of the ice formation and aerodynamic effects on swept wings. The purpose of this paper is to describe this research effort in more detail and report on the current results and status to date. 1
    Keywords: Aerodynamics
    Type: NASA/TM-2013-216555 , AIAA Paper 2013-28 , E-18752 , GRC-E-DAA-TN9726 , AIAA Atmospheric and Space Enviroment Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 5
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    A Non-Cut Cell Immersed Boundary Method for Use in Icing Simulations (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes a computational fluid dynamic method used for modelling changes in aircraft geometry due to icing. While an aircraft undergoes icing, the accumulated ice results in a geometric alteration of the aerodynamic surfaces. In computational simulations for icing, it is necessary that the corresponding geometric change is taken into consideration. The method used, herein, for the representation of the geometric change due to icing is a non-cut cell Immersed Boundary Method (IBM). Computational cells that are in a body fitted grid of a clean aerodynamic geometry that are inside a predicted ice formation are identified. An IBM is then used to change these cells from being active computational cells to having properties of viscous solid bodies. This method has been implemented in the NASA developed node centered, finite volume computational fluid dynamics code, FUN3D. The presented capability is tested for two-dimensional airfoils including a clean airfoil, an iced airfoil, and an airfoil in harmonic pitching motion about its quarter chord. For these simulations velocity contours, pressure distributions, coefficients of lift, coefficients of drag, and coefficients of pitching moment about the airfoil's quarter chord are computed and used for comparison against experimental results, a higher order panel method code with viscous effects, XFOIL, and the results from FUN3D's original solution process. The results of the IBM simulations show that the accuracy of the IBM compares satisfactorily with the experimental results, XFOIL results, and the results from FUN3D's original solution process.
    Keywords: Aerodynamics
    Type: NASA/TM-2013-217881 , E-18692 , AIAA Paper 2012-1204 , AIAA Aerospace Sciences Meeting; Jan 09, 2012 - Jan 12, 2012; Nashville, TN; United States
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  • 6
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    Testing of the Trim Tab Parametric Model in NASA Langley's Unitary Plan Wind Tunnel (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In support of NASA's Entry, Descent, and Landing technology development efforts, testing of Langley's Trim Tab Parametric Models was conducted in Test Section 2 of NASA Langley's Unitary Plan Wind Tunnel. The objectives of these tests were to generate quantitative aerodynamic data and qualitative surface pressure data for experimental and computational validation and aerodynamic database development. Six component force-and-moment data were measured on 38 unique, blunt body trim tab configurations at Mach numbers of 2.5, 3.5, and 4.5, angles of attack from -4deg to +20deg, and angles of sideslip from 0deg to +8deg. Configuration parameters investigated in this study were forebody shape, tab area, tab cant angle, and tab aspect ratio. Pressure Sensitive Paint was used to provide qualitative surface pressure mapping for a subset of these flow and configuration variables. Over the range of parameters tested, the effects of varying tab area and tab cant angle were found to be much more significant than varying tab aspect ratio relative to key aerodynamic performance requirements. Qualitative surface pressure data supported the integrated aerodynamic data and provided information to aid in future analyses of localized phenomena for trim tab configurations.
    Keywords: Aerodynamics
    Type: NF1676L-15681 , AIAA Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 7
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    Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Laminar-Flow Glove (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This is the presentation related to the paper of the same name describing Reynolds Averaged Navier Stokes (RANS) computational Fluid Dynamics (CFD) analysis of low speed stall aerodynamics of a swept wing with a laminar flow wing glove.
    Keywords: Aerodynamics
    Type: DFRC-E-DAA-TN9982 , DFRC-E-DAA-TN10986 , Applied Aerodynamics Conference; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 8
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    Laminar, Transitional, and Turbulent Heating on Mid Lift-to-Drag Ratio Entry Vehicles (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The boundary-layer transition characteristics and convective aeroheating levels on mid lift-to-drag ratio entry vehicle configurations have been studied through wind-tunnel testing. Several configurations were investigated, including elliptically blunted cylinders with both circular and elliptically flattened cross sections, biconic geometries based on launch vehicle dual-use shrouds, and parametrically optimized analytic geometries. Vehicles of this class have been proposed for high-mass Mars missions, such as sample return and crewed exploration, for which the conventional sphere-cone entry-vehicle geometries of previous Mars missions are insufficient. Testing was conducted at Mach 6 over a range of Reynolds numbers sufficient to generate laminar, transitional, and turbulent flow. Transition onset locations, both straight-line and cross-flow, and heating rates were obtained through global phosphor thermography. Supporting computations were performed to obtain heating rates for comparison with the data. Laminar data and predictions agreed to well within the experimental uncertainty. Fully turbulent data and predictions also agreed well. However, in transitional flow regions, greater differences were observed.
    Keywords: Aerodynamics
    Type: NF1676L-22328 , Journal of Spacecraft and Rockets; 50; 5; 937-959|AIAA Fluid Dynamics Conference and Exhibit; Jun 25, 2012 - Jun 28, 2012; New Orleans, LA; United States
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  • 9
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    Enhancements to the FAST-MAC Circulation Control Model and Recent High-Reynolds Number Testing in the National Transonic Facility (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.
    Keywords: Aerodynamics
    Type: NF1676L-15676 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 10
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    Long Penetration Mode Counterflowing Jets for Supersonic Slender Configurations - A Numerical Study (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A novel approach of using counterflowing jets positioned strategically on the aircraft and exploiting its long penetration mode (LPM) of interaction towards sonic-boom mitigation forms the motivation for this study. Given that most previous studies on the counterflowing LPM jet have all been on blunt bodies and at high supersonic or hypersonic flow conditions, exploring the feasibility to obtain a LPM jet issuing from a slender body against low supersonic freestream conditions is the main focus of this study. Computational fluid dynamics computations of axisymmetric models (cone-cylinder and quartic geometry), of relevance to NASA's High Speed project, are carried out using the space-time conservation element solution element viscous flow solver with unstructured meshes. A systematic parametric study is conducted to determine the optimum combination of counterflowing jet size, mass flow rate, and nozzle geometry for obtaining LPM jets. Details from these computations will be used to assess the potential of the LPM counterflowing supersonic jet as a means of active flow control for enabling supersonic flight over land and to establish the knowledge base for possible future implementation of such technologies.
    Keywords: Aerodynamics
    Type: NF1676L-15643 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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