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Sweeping jet actuator in a quiescent environment

Sweeping jet actuator in a quiescent environment

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43rd AIAA Fluid Dynamics Conference

作者： Koklu, Mehti Melton, Latunia P. NASA Langley Research Center Flow Physics and Control Branch MS 170 Hampton VA 23681 United States

ISBN: (纸本)9781624102141

This study presents a detailed analysis of a sweeping jet (fluidic oscillator) actuator. The sweeping jet actuator promises to be a viable flow control actuator candidate due to its simple, no moving part structure and its high momentum, spatially oscillating flow output. Hot-wire anemometer and particle image velocimetry measurements were carried out with an emphasis on understanding the actuator flow field in a quiescent environment. The time averaged, fluctuating, and instantaneous velocity measurements are provided. A modified actuator concept that incorporates high-speed solenoid valves to control the frequency of oscillation enabled phase averaged measurements of the oscillating jet. These measurements reveal that in a given oscillation cycle, the oscillating jet spends more time on each of the Coanda surfaces. In addition, the modified actuator generates four different types of flow fields, namely: a non oscillating downward jet, a non oscillating upward jet, a non oscillating straight jet, and an oscillating jet. The switching from an upward jet to a downward jet is accomplished by providing a single pulse from the solenoid valve. Once the flow is switched, the flow stays there until another pulse is received. The oscillating jet is compared with a non oscillating straight jet, which is a typical planar turbulent jet. The results indicate that the oscillating jet has a higher (5 times) spreading rate, more flow entrainment, and higher velocity fluctuations (equal to the mean velocity).

关键词： Actuators

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DNS, enstrophy balance, and the dissipation equation in a separated turbulent channel flow

DNS, enstrophy balance, and the dissipation equation in a se...

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21st AIAA Computational Fluid Dynamics Conference

作者： Balakumar, P. Rubinstein, R. Rumsey, C.L. Flow Physics and Control Branch NASA Langley Research Center Hampton VA 23681 United States Computational AeroSciences Branch NASA Langley Research Center Hampton VA 23681 United States

The turbulent flows through a plane channel and a channel with a constriction (2-D hill) are numerically simulated using DNS and RANS calculations. The Navier-Stokes equations in the DNS are solved using a higher order kinetic energy preserving central schemes and a fifth order accurate upwind biased WENO scheme for the space discretization. RANS calculations are performed using the NASA code CFL3D with the komega SST two-equation model and a full Reynolds stress model. Using DNS, the magnitudes of different terms that appear in the enstrophy equation are evaluated. The results show that the dissipation and the diffusion terms reach large values at the wall. All the vortex stretching terms have similar magnitudes within the buffer region. Beyond that the triple correlation among the vorticity and strain rate fluctuations becomes the important kinematic term in the enstrophy equation. This term is balanced by the viscous dissipation. In the separated flow, the triple correlation term and the viscous dissipation term peak locally and balance each other near the separated shear layer region. These findings concur with the analysis of Tennekes and Lumley, confirming that the energy transfer terms associated with the small-scale dissipation and the fluctuations of the vortex stretching essentially cancel each other, leaving an equation for the dissipation that is governed by the large-scale motion.

关键词： Navier Stokes equations

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SparkJet efficiency 51

SparkJet efficiency

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51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition 2013

作者： Golbabaei-Asl, Mona Knight, Doyle Anderson, Kellie Wilkinson, Stephen Rutgers - The State University of New Jersey Department of Mechanical and Aerospace Engineering New Brunswick NJ 08903 United States Rutgers - The State University of New Jersey Department of Chemical and Bio-Chemical Engineering New Brunswick NJ 08903 United States NASA Langley Research Center Flow Physics and Control Branch Hampton VA 23681 United States

ISBN: (数字)9781624101816

ISBN: (纸本)9781624101816

A novel method for determining the thermal effiency of the SparkJet is proposed. A SparkJet is attached to the end of a pendulum. The motion of the pendulum subsequent to a single spark discharge is measured using a laser displacement sensor. The measured displacement vs time is compared with the predictions of a theoretical perfect gas model to estimate the fraction of the spark discharge energy which results in heating the gas (i.e., increasing the translational-rotational temperature). The results from multiple runs for different capacitances of c = 3;5;10;20, and 40μF show that the thermal efficiency decreases with higher capacitive discharges. © 2013 by Mona Golbabaei Asl and Doyle Knight.

关键词： Pendulums

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Auto-calibration of spatially-resolving laser-Doppler velocimeters 51

Auto-calibration of spatially-resolving laser-Doppler veloci...

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51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition 2013

作者： Lowe, K. Todd Byun, Gwibo Neuhart, Dan H. Simpson, Roger L. Department of Aerospace and Ocean Engineering Virginia Tech 215 Randolph Hall Blacksburg VA 24060 United States AUR Inc. Blacksburg VA 24060 United States NASA Langley Research Center Flow Physics and Control Branch Hampton VA 23681 United States Department of Aerospace and Ocean Engineering AUR Inc. Virginia Tech Blacksburg VA 24060 United States

ISBN: (数字)9781624101816

ISBN: (纸本)9781624101816

A method for the auto-calibration of laser-Doppler velocimetry (LDV) probe volumes with sub-measurement-volume spatial resolution is presented. The spatially-resolving LDV (SRLDV) technique, with demonstrated spatial resolution better than 3μm, provides low uncertainty velocity statistics in high Reynolds number flows and those with limited optical access. A key enabling aspect of the technique is the spatial characterization of the interference fringes. For SRLDV designs with measurement volume lengths on the order of the beam diameter, a first-order algebraic expansion of the exact analytical solution for interference fringe spacing provides an adequate and readily characterized instrument model. Given this linear approximation, calibration of two parameters, the mean fringe spacing and fringe spacing gradient (FSG) is required. While, the former is required for conventional LDV, calibration of the FSG is unique to SRLDV. Described herein is an auto-calibration scheme for determining the FSG by forcing velocity profile continuity from partially overlapping measurement positions in any flow with a substantial velocity gradient, thereby using the accuracy of a commercial linear traverse as the spatial measurement standard. Results from an application of the auto-calibration method are presented in an ad hoc seeded jet flow, highlighting that a rigorous flow calibration standard is not required to achieve the auto-calibration. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

关键词： Calibration

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Enhancements to the FAST-MAC circulation control model and recent High-Reynolds number testing in the National Transonic Facility 31

Enhancements to the FAST-MAC circulation control model and r...

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31st AIAA Applied Aerodynamics Conference

作者： Milholen II, William E. Jones, Gregory S. Chan, David T. Goodliff, Scott L. Anders, Scott G. Melton, Latunia P. Carter, Melissa B. Allan, Brian G. Capone, Francis J. Configuration Aerodynamics Branch NASA Langley Research Center Mail Stop 499 Hampton VA 23681 United States Flow Physics and Control Branch NASA Langley Research Center Mail Stop 170 Hampton VA 23681 United States Jacobs Technology Inc NASA Langley Research Center Mail Stop 267 Hampton VA 23681 United States Revolutionary Aviation Technologies Branch NASA Langley Research Center Mail Stop 207 Hampton VA 23681 United States AMS NASA Langley Research Center Mail Stop 499 Hampton VA 23681 United States

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.

关键词： Wind tunnels

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Development of the circulation control flow scheme used in the NTF semi-span FAST-MAC model 31

Development of the circulation control flow scheme used in t...

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31st AIAA Applied Aerodynamics Conference

作者： Jones, Gregory S. Milholen II, William E. Chan, David T. Allan, Brian Goodliff, Scott L. Melton, Latunia Anders, Scott G. Carter, Melissa Capone, Francis NASA Langley Research Center Flow Physics and Control Branch Mail Stop 170 Hampton VA 23681 United States NASA Langley Research Center Configuration Aerodynamics Branch Mail Stop 499 Hampton VA 23681 United States NASA Langley Research Center Jacobs Technology Inc. Mail Stop 267 Hampton VA 23681 United States NASA Langley Research Center Revolutionary Aviation Technologies Branch Mail Stop 207 Hampton VA 23681 United States NASA Langley Research Center AMS Mail Stop 499 Hampton VA 23681 United States

The application of a circulation control system for high Reynolds numbers was experimentally validated with the Fundamental Aerodynamic Subsonic Transonic Modular Active control semi-span model in the NASA Langley National Transonic Facility. This model utilized four independent flow paths to modify the lift and thrust performance of a representative advanced transport type of wing. The design of the internal flow paths highlights the challenges associated with high Reynolds number testing in a cryogenic pressurized wind tunnel. Weight flow boundaries for the air delivery system were identified at mildly cryogenic conditions ranging from 0.1 to 10 lbm/sec. Results from the test verified system performance and identified solutions associated with the weight-flow metering system that are linked to internal perforated plates used to achieve flow uniformity at the jet exit.

关键词： Wind tunnels

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Development of advanced high lift leading edge technology for laminar flow wings 51

Development of advanced high lift leading edge technology fo...

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51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition 2013

作者： Bright, Michelle M. Korntheuer, Andrea Komadina, Steve Lin, John C. Northrop Grumman Aerospace Systems Flight Analysis Department El Segundo CA 90245 United States Northrop Grumman Aerospace Systems Aerodynamics Analysis Department El Segundo CA 90245 United States Northrop Grumman Aerospace Systems El Segundo CA 90245 United States NASA Langley Research Center Flow Physics and Control Branch MS170 Hampton VA 23681 United States

ISBN: (数字)9781624101816

ISBN: (纸本)9781624101816

This paper describes the Advanced High Lift Leading Edge (AHLLE) task performed by Northrop Grumman Systems Corporation, Aerospace Systems (NGAS) for the NASA Subsonic Fixed Wing project in an effort to develop enabling high-lift technology for laminar flow wings. Based on a known laminar cruise airfoil that incorporated an NGASdeveloped integrated slot design, this effort involved using Computational Fluid Dynamics (CFD) analysis and quality function deployment (QFD) analysis on several leading edge concepts, and subsequently down-selected to two blown leading-edge concepts for testing. A 7-foot-span AHLLE airfoil model was designed and fabricated at NGAS and then tested at the NGAS 7' x 10' Low Speed Wind Tunnel in Hawthorne, CA. The model configurations tested included: baseline, deflected trailing edge, blown deflected trailing edge, blown leading edge, morphed leading edge, and blown/morphed leading edge. A successful demonstration of high lift leading edge technology was achieved, and the target goals for improved lift were exceeded by 30% with a maximum section lift coefficient (Cl) of 5.2. Maximum incremental section lift coefficients (ΔCl) of 3.5 and 3.1 were achieved for a blown drooped (morphed) leading edge concept and a non-drooped leading edge blowing concept, respectively. The most effective AHLLE design yielded an estimated 94% lift improvement over the conventional high lift Krueger flap configurations while providing laminar flow capability on the cruise configuration. © 2013 by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Laminar flow

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Validation of 3-D ice accretion measurement methodology for experimental aerodynamic simulation

Validation of 3-D ice accretion measurement methodology for ...

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35th Annual International Conference of IEEE Engineering in Medicine and Biology Society, EMBC 2013 in conjunction with 52nd Annual Conference of Japanese Society for Medical and Biological Engineering, JSMBE

作者： Broeren, Andy P. Addy Jr., Harold E. Lee, Sam Monastero, Marianne C. NASA John H. Glenn Research Center Icing Branch 21000 Brookpark Rd. MS 11-2 Cleveland OH 44135 United States LLC Icing Branch 21000 Brookpark Rd. MS 11-2 Cleveland OH 44135 United States University of Illinois at Urbana-Champaign Urbana IL 61801 United States Rensselaer Polytechnic Institute Department of Mechanical Aerospace and Nuclear Engineering Center for Flow Physics and Control Watervliet Facility 110 8th St. Troy NY 12180 United States

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 iceaccretion 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 iceaccretion 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 1.8×106 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,

关键词： Ice

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Trapezoidal wing experimental repeatability and velocity profiles in the 14- by 22-foot subsonic tunnel (Invited) 50

Trapezoidal wing experimental repeatability and velocity pro...

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50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

作者： Hannon, Judith A. Washburn, Anthony E. Jenkins, Luther N. Watson, Ralph D. Flow Physics and Control Branch Mail Stop 170 NASA Langley Research Center Hampton VA 23681 United States

ISBN: (数字)9781600869365

The AIAA Applied Aerodynamics Technical Committee sponsored a High Lift Prediction Workshop held in June 2010. For this first workshop, data from the Trapezoidal Wing experiments were used for comparison to CFD. This paper presents long-term and short-term force and moment repeatability analyses for the Trapezoidal Wing model tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel. This configuration was chosen for its simplified high-lift geometry, publicly available set of test data, and previous CFD experience with this configuration. The Trapezoidal Wing is a three-element semi-span swept wing attached to a body pod. These analyses focus on configuration 1 tested in 1998 (Test 478), 2002 (Test 506), and 2003 (Test 513). This paper also presents model velocity profiles obtained on the main element and on the flap during the 1998 test. These velocity profiles are primarily at an angle of attack of 28 degrees and semi-span station of 83% and show confluent boundary layers and wakes. Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Computational fluid dynamics

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Aerodynamic shaping of spherical turrets to mitigate aero-optic effects 50

Aerodynamic shaping of spherical turrets to mitigate aero-op...

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50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

作者： Crahan, Grady Rennie, Mark Rapagnani, Larry Jumper, Eric J. Gogineni, Siviram Center for Flow Physics and Control University of Notre Dame Notre Dame IN 46556 United States Spectral Energies LLC Dayton OH 45431 United States

ISBN: (数字)9781600869365

Hemispherical turrets provide an efficient means of mounting an optical system on an aircraft, but are susceptible to the formation of flows, such as shocks and separated shear layers that are known to present a severe optical aberration to the outgoing beam. In this paper, methods are presented to prevent or mitigate the formation of these aero-optic flows using improved aerodynamic shaping of the turret itself. The investigation focuses on the use of the "virtual duct" approach, which has been shown previously to successfully prevent aero-optic flows on an underwing pod up to freestream Mach numbers over 0.8. Computational fluid dynamic results are presented to demonstrate the effectiveness of the virtual duct approach, as well as other mitigation strategies. Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc.

关键词： Ducts

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