Leading-edge singularities in thin-airfoil theory

NASA Technical Reports Server (NTRS)

Jones, R. T.

1976-01-01

If the thin airfoil theory is applied to an airfoil having a rounded leading edge, a certain error will arise in the determination of the pressure distribution around the nose. It is shown that the evaluation of the drag of such a blunt nosed airfoil by the thin airfoil theory requires the addition of a leading edge force, analogous to the leading edge thrust of the lifting airfoil. The method of calculation is illustrated by application to: (1) The Joukowski airfoil in subsonic flow; and (2) the thin elliptic cone in supersonic flow. A general formula for the edge force is provided which is applicable to a variety of wing forms.

Development of X-43A Mach 10 Leading Edges

NASA Technical Reports Server (NTRS)

Ohlhorst, Craig W.; Glass, David E.; Bruce, Walter E., III; Lindell, Michael C.; Vaughn, Wallace L.; Dirling, R. B., Jr.; Hogenson, P. A.; Nichols, J. M.; Risner, N. W.; Thompson, D. R.

2005-01-01

The nose leading edge of the Hyper-X Mach 10 vehicle was orginally anticipated to reach temperatures near 4000 F at the leading-edge stagnation line. A SiC coated carbon/carbon (C/C) leading-edge material will not survive that extreme temperature for even a short duration single flight. To identify a suitable leading edge for the Mach 10 vehicle, arc-jet testing was performed on thirteen leading-edge segments fabricated from different material systems to evaluate their performance in a simulated flight environment. Hf, Zr, Si, and Ir based materials, in most cases as a coating on C/C, were included in the evaluation. Afterwards, MER, Tucson, AZ was selected as the supplier of the flight vehicle leading edges. The nose and the vertical and horizontal tail leading edges were fabricated out of a 3:1 biased high thermal conductivity C/C. The leading edges were coated with a three layer coating comprised of a SiC conversion of the top surface of the C/C, followed by a chemical vapor deposited layer of SiC, followed by a thin chemical vapor deposited layer of HfC. This paper will describe the fabrication of the Mach 10 C/C leading edges and the testing performed to validate performance.

Aerodynamic and Aerothermodynamic Layout of the Hypersonic Flight Experiment Shefex

NASA Astrophysics Data System (ADS)

Eggers, Th.

2005-02-01

The purpose of the SHarp Edge Flight EXperiment SHEFEX is the investigation of possible new shapes for future launcher or reentry vehicles [1]. The main focus is the improvement of common space vehicle shapes by application of facetted surfaces and sharp edges. The experiment will enable the time accurate investigation of the flow effects and their structural answer during the hypersonic flight from 90 km down to an altitude of 20 km. The project, being performed under responsibility of the German Aerospace Center (DLR) is scheduled to fly on top of a two-stage solid propellant sounding rocket for the first half of 2005. The paper contains a survey of the aerodynamic and aerothermodynamic layout of the experimental vehicle. The results are inputs for the definition of the structural layout, the TPS and the flight instrumentation as well as for the preparation of the flight test performed by the Mobile Rocket Base of DLR.

Space Shuttle aerothermodynamic data report, phase C

NASA Technical Reports Server (NTRS)

1985-01-01

Space shuttle aerothermodynamic data, collected from a continuing series of wind tunnel tests, are permanently stored with the Data Management Services (DMS) system. Information pertaining to current baseline configuration definition is also stored. Documentation of DMS processed data arranged sequentially and by space shuttle configuration are included. An up-to-date record of all applicable aerothermodynamic data collected, processed, or summarized during the space shuttle program is provided. Tables are designed to provide suvery information to the various space shuttle managerial and technical levels.

Experimental investigation of leading-edge thrust at supersonic speeds

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1983-01-01

Wings, designed for leading edge thrust at supersonic speeds, were investigated in the Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, 2.16, and 2.36. Experimental data were obtained on a uncambered wing which had three interchangeable leading edges that varied from sharp to blunt. The leading edge thrust concept was evaluated. Results from the investigation showed that leading edge flow separation characteristics of all wings tested agree well with theoretical predictions. The experimental data showed that significant changes in wing leading edge bluntness did not affect the zero lift drag of the uncambered wings.

Aerothermodynamic Insight From The HIFIRE Program

NASA Astrophysics Data System (ADS)

Kimmel, Roger L.; Adamczak, David; Dolvin, Douglas; Borg, Matthew; Stanfield, Scott

2011-05-01

The HIFiRE (Hypersonic International Flight Research and Experimentation) program is a joint venture of the United States Air Force Research Laboratory and Australian Defence Science and Technology Organisation to utilize economical flight research opportunities in the exploration of flight science issues for space access systems. Flights 1 and 5 focus on collecting high-resolution experimental data on critical aerothermodynamic phenomena, including laminar-turbulent transition and shock/boundary layer interactions. Flight 1, successfully flown in March 2010, employed a test article composed of a 7-deg right angle cone, followed by a cylinder and flare. The test article remained attached to the second-stage booster throughout the ballistic trajectory. Flight 5, to be launched in a similar fashion, will feature a 2:1 elliptic cross-section cone as the test article. For both flights significant resources have been invested in pre-flight aerothermodynamic analysis and testing. This manuscript will summarize the overall strategy of the HIFiRE program, review the pre-flight aerothermodynamic analysis for Flights 1 and 5, and present a brief look at preliminary results from the post-flight analysis of Flight 1.

Dynamic Stall Characteristics of Drooped Leading Edge Airfoils

NASA Technical Reports Server (NTRS)

Sankar, Lakshmi N.; Sahin, Mehmet; Gopal, Naveen

2000-01-01

Helicopters in high-speed forward flight usually experience large regions of dynamic stall over the retreating side of the rotor disk. The rapid variations in the lift and pitching moments associated with the stall process can result in vibratory loads, and can cause fatigue and failure of pitch links. In some instances, the large time lag between the aerodynamic forces and the blade motion can trigger stall flutter. A number of techniques for the alleviation of dynamic stall have been proposed and studied by researchers. Passive and active control techniques have both been explored. Passive techniques include the use of high solidity rotors that reduce the lift coefficients of individual blades, leading edge slots and leading edge slats. Active control techniques include steady and unsteady blowing, and dynamically deformable leading edge (DDLE) airfoils. Considerable amount of experimental and numerical data has been collected on the effectiveness of these concepts. One concept that has not received as much attention is the drooped-leading edge airfoil idea. It has been observed in wind tunnel studies and flight tests that drooped leading edge airfoils can have a milder dynamic stall, with a significantly milder load hysteresis. Drooped leading edge airfoils may not, however, be suitable at other conditions, e.g. in hover, or in transonic flow. Work needs to be done on the analysis and design of drooped leading edge airfoils for efficient operation in a variety of flight regimes (hover, dynamic stall, and transonic flow). One concept that is worthy of investigation is the dynamically drooping airfoil, where the leading edge shape is changed roughly once-per-rev to mitigate the dynamic stall.

Symmetric airfoil geometry effects on leading edge noise.

PubMed

Gill, James; Zhang, X; Joseph, P

2013-10-01

Computational aeroacoustic methods are applied to the modeling of noise due to interactions between gusts and the leading edge of real symmetric airfoils. Single frequency harmonic gusts are interacted with various airfoil geometries at zero angle of attack. The effects of airfoil thickness and leading edge radius on noise are investigated systematically and independently for the first time, at higher frequencies than previously used in computational methods. Increases in both leading edge radius and thickness are found to reduce the predicted noise. This noise reduction effect becomes greater with increasing frequency and Mach number. The dominant noise reduction mechanism for airfoils with real geometry is found to be related to the leading edge stagnation region. It is shown that accurate leading edge noise predictions can be made when assuming an inviscid meanflow, but that it is not valid to assume a uniform meanflow. Analytic flat plate predictions are found to over-predict the noise due to a NACA 0002 airfoil by up to 3 dB at high frequencies. The accuracy of analytic flat plate solutions can be expected to decrease with increasing airfoil thickness, leading edge radius, gust frequency, and Mach number.

An Overview of the Space Shuttle Aerothermodynamic Design

NASA Technical Reports Server (NTRS)

Martin, Fred

2011-01-01

The Space Shuttle Thermal Protection System was one of the three areas that required the development of new technology. The talk discusses the pre-flight development of the aerothermodynamic environment which was based on Mach 8 wind tunnel data. A high level overview of the pre-flight heating rate predictions and comparison to the Orbiter Flight Test (OFT) data is presented, along with a discussion of the dramatic improvement in the state-of-the-art in aerothermodynamic capability that has been used to support the Shuttle Program. A high level review of the Orbiter aerothermodynamic design is discussed, along with improvements in Computational Fluid Dynamics and wind tunnel testing that was required for flight support during the last 30 years. The units have been removed from the plots, and the discussion is kept at a high level.

Wing Leading Edge Debris Analysis

NASA Technical Reports Server (NTRS)

Shah, Sandeep; Jerman, Gregory

2004-01-01

This is a slide presentation showing the Left Wing Leading Edge (WLE) heat damage observations: Heavy "slag" deposits on select RCC panels. Eroded and knife-edged RCC rib sections. Excessive overheating and slumping of carrier panel tiles. Missing or molten attachment bolts but intact bushing. Deposit mainly on "inside" RCC panel. Deposit on some fractured RCC surface

Leading-edge flow criticality as a governing factor in leading-edge vortex initiation in unsteady airfoil flows

NASA Astrophysics Data System (ADS)

Ramesh, Kiran; Granlund, Kenneth; Ol, Michael V.; Gopalarathnam, Ashok; Edwards, Jack R.

2018-04-01

A leading-edge suction parameter (LESP) that is derived from potential flow theory as a measure of suction at the airfoil leading edge is used to study initiation of leading-edge vortex (LEV) formation in this article. The LESP hypothesis is presented, which states that LEV formation in unsteady flows for specified airfoil shape and Reynolds number occurs at a critical constant value of LESP, regardless of motion kinematics. This hypothesis is tested and validated against a large set of data from CFD and experimental studies of flows with LEV formation. The hypothesis is seen to hold except in cases with slow-rate kinematics which evince significant trailing-edge separation (which refers here to separation leading to reversed flow on the aft portion of the upper surface), thereby establishing the envelope of validity. The implication is that the critical LESP value for an airfoil-Reynolds number combination may be calibrated using CFD or experiment for just one motion and then employed to predict LEV initiation for any other (fast-rate) motion. It is also shown that the LESP concept may be used in an inverse mode to generate motion kinematics that would either prevent LEV formation or trigger the same as per aerodynamic requirements.

A Perspective on Computational Aerothermodynamics at NASA

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2007-01-01

The evolving role of computational aerothermodynamics (CA) within NASA over the past 20 years is reviewed. The presentation highlights contributions to understanding the Space Shuttle pitching moment anomaly observed in the first shuttle flight, prediction of a static instability for Mars Pathfinder, and the use of CA for damage assessment in post-Columbia mission support. In the view forward, several current challenges in computational fluid dynamics and aerothermodynamics for hypersonic vehicle applications are discussed. Example simulations are presented to illustrate capabilities and limitations. Opportunities to advance the state-of-art in algorithms, grid generation and adaptation, and code validation are identified.

Shuttle Tethered Aerothermodynamics Research Facility (STARFAC) Instrumentation Requirements

NASA Technical Reports Server (NTRS)

Wood, George M.; Siemers, Paul M.; Carlomagno, Giovanni M.; Hoffman, John

1986-01-01

The instrumentation requirements for the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC) are presented. The typical physical properties of the terrestrial atmosphere are given along with representative atmospheric daytime ion concentrations and the equilibrium and nonequilibrium gas property comparison from a point away from a wall. STARFAC science and engineering measurements are given as are the TSS free stream gas analysis. The potential nonintrusive measurement techniques for hypersonic boundary layer research are outlined along with the quantitative physical measurement methods for aerothermodynamic studies.

Aerodynamics and Aerothermodynamics of undulated re-entry vehicles

NASA Astrophysics Data System (ADS)

Kaushikh, K.; Arunvinthan, S.; Pillai, S. Nadaraja

2018-01-01

Aerodynamic and aerothermodynamic analysis is a fundamental basis for the design of a hypersonic vehicle. In this work, aerodynamic and aerothermodynamic analyses of a blunt body vehicle with undulations on its after-body are studied with the help of numerical simulations. A crew exploration vehicle (CEV) is taken for initial analysis and undulations with varying amplitude and wavelength are introduced on CEV's after-body. Numerical simulations were carried out for CEV and for CEV with undulations at Mach 3.0 and 7.0 for angles of attack ranging from -20° to +20° with increments of +5°. The results show that introduction of undulations did not have a significant impact on mono stability and lift-drag characteristics of the vehicle. It was also observed that introduction of undulations improved the aerothermodynamic characteristics of CEV. A reduction of about 36% in maximum heat flux at Mach 3.0 and about 21% at Mach 7.0 compared to the maximum heat flux for CEV was observed.

Oceanic Extreme Model Atmospheres for Aerothermodynamic Calculations,

DTIC Science & Technology

Atmospheric temperature, Atmospheric sounding, Regression analysis, Aerothermodynamics, Marine meteorology, Radiosondes, Weather stations, Newfoundland(Province), Marshall Islands , Arabia, Iran, Coastal regions

Sharp Refractory Composite Leading Edges on Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Walker, Sandra P.; Sullivan, Brian J.

2003-01-01

On-going research of advanced sharp refractory composite leading edges for use on hypersonic air-breathing vehicles is presented in this paper. Intense magnitudes of heating and of heating gradients on the leading edge lead to thermal stresses that challenge the survivability of current material systems. A fundamental understanding of the problem is needed to further design development. Methodology for furthering the technology along with the use of advanced fiber architectures to improve the thermal-structural response is explored in the current work. Thermal and structural finite element analyses are conducted for several advanced fiber architectures of interest. A tailored thermal shock parameter for sharp orthotropic leading edges is identified for evaluating composite material systems. The use of the tailored thermal shock parameter has the potential to eliminate the need for detailed thermal-structural finite element analyses for initial screening of material systems being considered for a leading edge component.

CFD Aerothermodynamic Characterization Of The IXV Hypersonic Vehicle

NASA Astrophysics Data System (ADS)

Roncioni, P.; Ranuzzi, G.; Marini, M.; Battista, F.; Rufolo, G. C.

2011-05-01

In this paper, and in the framework of the ESA technical assistance activities for IXV project, the numerical activities carried out by ASI/CIRA to support the development of Aerodynamic and Aerothermodynamic databases, independent from the ones developed by the IXV Industrial consortium, are reported. A general characterization of the IXV aerothermodynamic environment has been also provided for cross checking and verification purposes. The work deals with the first year activities of Technical Assistance Contract agreed between the Italian Space Agency/CIRA and ESA.

NASA's hypersonic fluid and thermal physics program (Aerothermodynamics)

NASA Technical Reports Server (NTRS)

Graves, R. A.; Hunt, J. L.

1985-01-01

This survey paper gives an overview of NASA's hypersonic fluid and thermal physics program (recently renamed aerothermodynamics). The purpose is to present the elements of, example results from, and rationale and projection for this program. The program is based on improving the fundamental understanding of aerodynamic and aerothermodynamic flow phenomena over hypersonic vehicles in the continuum, transitional, and rarefied flow regimes. Vehicle design capabilities, computational fluid dynamics, computational chemistry, turbulence modeling, aerothermal loads, orbiter flight data analysis, orbiter experiments, laser photodiagnostics, and facilities are discussed.

Shock Interaction Control for Scramjet Cowl Leading Edges

NASA Technical Reports Server (NTRS)

Albertson, Cindy W.; Venkat, Venki, S.

2005-01-01

An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface.

Performance of hydrofoils with humpback whale-like leading edge protuberances.

NASA Astrophysics Data System (ADS)

Levshin, Alexandra; Henoch, Charles; Johari, Hamid

2005-11-01

The humpback whale (Megaptera novaeangliae) is extremely maneuverable, compared to other whale species, despite its large size and rigid body. Turning maneuvers are especially evident during pursuit of prey. The agility of humpback whale has been attributed to their use of pectoral flippers. The thick flippers have large aspect ratios, and large scale protuberances are present on the leading edge. The flippers do not flap during turning maneuvers. The cross-section of the flipper has a profile similar to a NACA 634-021 airfoil. The amplitude of leading edge protuberances ranges from 2.5 to 12% of the chord, with a spanwise extent of 10 to 50% the chord depending on the location along the span. It has been hypothesized that the `bumpy' leading edge is used for flow control. To examine the effects of protuberances on the leading edge of hydrofoils, a series of rectangular foils with bumpy leading edges were manufactured. The leading edge is sinusoidal in the spanwise direction with amplitudes and wavelengths comparable to that of humpback whale's flippers. The forces and moments on these bumpy foils were measured in a water tunnel and compared with a smooth leading edge foil.

Aerothermodynamic data base. Data file contents report, phase C

NASA Technical Reports Server (NTRS)

Lutz, G. R.

1983-01-01

Space shuttle aerothermodynamic data, collected from a continuing series of wind tunnel tests, are permanently stored with the Data Management Services (DMS) system. Information pertaining to current baseline configuration definition is also stored. Documentation of DMS processed data arranged sequentially and by space shuttle configuration is listed to provide an up-to-date record of all applicable aerothermodynamic data collected, processed, or summarized during the space shuttle program. Tables provide survey information to the various space shuttle managerial and technical levels.

Span efficiency of wings with leading edge protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2013-11-01

Past work has shown that sinusoidal leading edge protuberances resembling those found on humpback whale flippers alter the lift and drag coefficients of full- and finite-span foils and wings depending on the angle of attack and leading edge geometry. Although the load characteristics of protuberance modified finite-span wings have been reported for flipper-like geometries at higher Reynolds numbers and for rectangular planforms at lower Reynolds numbers, the effects of leading edge geometry on the span efficiency, which is indicative of the deviation of the spanwise lift distribution from elliptical and the viscous effects, for a range of planforms and Reynolds numbers have not been addressed. The lift and drag coefficients of 7 rectangular, 2 swept, and 2 flipper-like planform models with aspect ratios of 4.3, 4.0, and 8.86, respectively, were used to compute the span efficiency at Reynolds numbers ranging from 0.9 to 4.5 × 105. The span efficiency, based on the data at lower angles of attack, of modified wings was compared with the unmodified models. For the cases considered, the span efficiencies of the leading edge modified models were less than those of the equivalent unmodified models. The dependence of span efficiency on the leading edge geometry, planform, and Reynolds number will be presented. Supported by the ONR-ULI program.

Heat-Pipe-Cooled Leading Edges for Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Glass, David E.

2006-01-01

Heat pipes can be used to effectively cool wing leading edges of hypersonic vehicles. . Heat-pipe leading edge development. Design validation heat pipe testing confirmed design. Three heat pipes embedded and tested in C/C. Single J-tube heat pipe fabricated and testing initiated. HPCLE work is currently underway at several locations.

Prediction of the Aerothermodynamic Environment of the Huygens Probe

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Striepe, Scott A.; Wright, Michael J.; Bose, Deepak; Sutton, Kenneth; Takashima, Naruhisa

2005-01-01

An investigation of the aerothermodynamic environment of the Huygens entry probe has been conducted. A Monte Carlo simulation of the trajectory of the probe during entry into Titan's atmosphere was performed to identify a worst-case heating rate trajectory. Flowfield and radiation transport computations were performed at points along this trajectory to obtain convective and radiative heat-transfer distributions on the probe's heat shield. This investigation identified important physical and numerical factors, including atmospheric CH4 concentration, transition to turbulence, numerical diffusion modeling, and radiation modeling, which strongly influenced the aerothermodynamic environment.

A Thermostructural Analysis of a Diboride Composite Leading Edge

NASA Technical Reports Server (NTRS)

Kowalski, Tom; Buesking, Kent; Kolodziej, Paul; Bull, Jeff

1996-01-01

In an effort to support the design of zirconium diboride composite leading edges for hypersonic vehicles, a finite element model (FEM) of a prototype leading edge was created and finite element analysis (FEA) was employed to assess its thermal and structural response to aerothermal boundary conditions. Unidirectional material properties for the structural components of the leading edge, a continuous fiber reinforced diboride composite, were computed with COSTAR. These properties agree well with those experimentally measured. To verify the analytical approach taken with COSMOS/M, an independent FEA of one of the leading edge assembly components was also done with COSTAR. Good agreement was obtained between the two codes. Both showed that a unidirectional lay-up had the best margin of safety for a simple loading case. Both located the maximum stress in the same region and ply. The magnitudes agreed within 4 percent. Trajectory based aerothermal heating was then applied to the leading edge assembly FEM created with COSMOS/M to determine steady state temperature response, displacement, stresses, and contact forces due to thermal expansion and thermal strains. Results show that the leading edge stagnation line temperature reached 4700 F. The maximum computed failure index for the laminated composite components peaks at 4.2, and is located at the bolt flange in layer 2 of the side bracket. The temperature gradient in the tip causes a compressive stress of 279 ksi along its width and substantial tensile stresses within its depth.

Hypersonic aerothermodynamic and scramjet research using high enthalpy shock tunnel

NASA Astrophysics Data System (ADS)

Itoh, K.; Ueda, S.; Tanno, H.; Komuro, T.; Sato, K.

A high enthalpy shock tunnel is a potential facility for gaining knowledge to develop modern aerothermodynamic and propulsion technologies. The largest high enthalpy shock tunnel HIEST was built at NAL Kakuda in 1997, aiming for aerothermodynamic tests of Japan's space vehicle HOPE and scramjet propulsion systems. Selected topics from the experimental studies carried out using HIEST so far, such as the nonequilibrium aerodynamics of HOPE, the surface catalytic effect on aerodynamic heating and scramjet performance are described.

Laminar Flow Control Leading Edge Systems in Simulated Airline Service

NASA Technical Reports Server (NTRS)

Wagner, R. D.; Maddalon, D. V.; Fisher, D. F.

1988-01-01

Achieving laminar flow on the wings of a commercial transport involves difficult problems associated with the wing leading edge. The NASA Leading Edge Flight Test Program has made major progress toward the solution of these problems. The effectiveness and practicality of candidate laminar flow leading edge systems were proven under representative airline service conditions. This was accomplished in a series of simulated airline service flights by modifying a JetStar aircraft with laminar flow leading edge systems and operating it out of three commercial airports in the United States. The aircraft was operated as an airliner would under actual air traffic conditions, in bad weather, and in insect infested environments.

Development of Advanced High Lift Leading Edge Technology for Laminar Flow Wings

NASA Technical Reports Server (NTRS)

Bright, Michelle M.; Korntheuer, Andrea; Komadina, Steve; Lin, John C.

2013-01-01

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

Leading-edge receptivity for blunt-nose bodies

NASA Technical Reports Server (NTRS)

Kerschen, Edward J.

1991-01-01

This research program investigates boundary-layer receptivity in the leading-edge region for bodies with blunt leading edges. Receptivity theory provides the link between the unsteady distrubance environment in the free stream and the initial amplitudes of the instability waves in the boundary layer. This is a critical problem which must be addressed in order to develop more accurate prediction methods for boundary-layer transition. The first phase of this project examines the effects of leading-edge bluntness and aerodynamic loading for low Mach number flows. In the second phase of the project, the investigation is extended to supersonic Mach numbers. Singular perturbation techniques are utilized to develop an asymptotic theory for high Reynolds numbers. In the first year, the asymptotic theory was developed for leading-edge receptivity in low Mach number flows. The case of a parabolic nose is considered. Substantial progress was made on the Navier-Sotkes computations. Analytical solutions for the steady and unsteady potential flow fields were incorporated into the code, greatly expanding the types of free-stream disturbances that can be considered while also significantly reducing the the computational requirements. The time-stepping algorithm was modified so that the potential flow perturbations induced by the unsteady pressure field are directly introduced throughout the computational domain, avoiding an artificial 'numerical diffusion' of these from the outer boundary. In addition, the start-up process was modified by introducing the transient Stokes wave solution into the downstream boundary conditions.

Wing Leading Edge Concepts for Noise Reduction

NASA Technical Reports Server (NTRS)

Shmilovich, Arvin; Yadlin, Yoram; Pitera, David M.

2010-01-01

This study focuses on the development of wing leading edge concepts for noise reduction during high-lift operations, without compromising landing stall speeds, stall characteristics or cruise performance. High-lift geometries, which can be obtained by conventional mechanical systems or morphing structures have been considered. A systematic aerodynamic analysis procedure was used to arrive at several promising configurations. The aerodynamic design of new wing leading edge shapes is obtained from a robust Computational Fluid Dynamics procedure. Acoustic benefits are qualitatively established through the evaluation of the computed flow fields.

Method for a Leading Edge Slat on a Wing of an Aircraft

NASA Technical Reports Server (NTRS)

Pitt, Dale M. (Inventor); Eckstein, Nicholas Stephen (Inventor)

2016-01-01

A method for managing a flight control surface system. A leading edge device is moved on a leading edge from an undeployed position to a deployed position. The leading edge device has an outer surface, an inner surface, and a deformable fairing attached to the leading edge device such that the deformable fairing covers at least a portion of the inner surface. The deformable fairing changes from a deformed shape to an original shape when the leading edge device is moved to the deployed position. The leading edge device is then moved from the deployed position to the undeployed position, wherein the deformable fairing changes from the original shape to the deformed shape.

Overview of aerothermodynamic loads definition study

NASA Technical Reports Server (NTRS)

Gaugler, Raymond E.

1989-01-01

Over the years, NASA has been conducting the Advanced Earth-to-Orbit (AETO) Propulsion Technology Program to provide the knowledge, understanding, and design methodology that will allow the development of advanced Earth-to-orbit propulsion systems with high performance, extended service life, automated operations, and diagnostics for in-flight health monitoring. The objective of the Aerothermodynamic Loads Definition Study is to develop methods to more accurately predict the operating environment in AETO propulsion systems, such as the Space Shuttle Main Engine (SSME) powerhead. The approach taken consists of 2 parts: to modify, apply, and disseminate existing computational fluid dynamics tools in response to current needs and to develop new technology that will enable more accurate computation of the time averaged and unsteady aerothermodynamic loads in the SSME powerhead. The software tools are detailed. Significant progress was made in the area of turbomachinery, where there is an overlap between the AETO efforts and research in the aeronautical gas turbine field.

The leading-edge vortex of yacht sails

NASA Astrophysics Data System (ADS)

Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

We experimentally show, for the first time, that a stable Leading-Edge Vortex (LEV) can be formed on an asymmetric spinnaker, which is a high-lift sail used by yachts to sail downwind. We tested a 3D printed rigid sail in a water flume at a chord-based Reynolds number of ca. 104. We found that on the leeward side of the sail (the suction side), the flow separates at the leading edge reattaching further downstream and forming a stable LEV. The LEV grows in diameter from the root to the tip of the sail, where it merges with the tip vortex. We detected the LEV using the γ criterion, and we verified its stability over time. The lift contribution provided by the LEV was computed solving a complex potential model of each sail section. This analysis indicated that the LEV provides a substantial contribution to the total sail's lift. These findings suggest that the maximum lift of low-aspect-ratio wings with a sharp leading edge, such as spinnakers, can be enhanced by promoting a stable LEV. This work was funded by the Consejo Nacional de Ciencia y Tecnologia (CONACYT).

Structural Health Monitoring Analysis for the Orbiter Wing Leading Edge

NASA Technical Reports Server (NTRS)

Yap, Keng C.

2010-01-01

This viewgraph presentation reviews Structural Health Monitoring Analysis for the Orbiter Wing Leading Edge. The Wing Leading Edge Impact Detection System (WLE IDS) and the Impact Analysis Process are also described to monitor WLE debris threats. The contents include: 1) Risk Management via SHM; 2) Hardware Overview; 3) Instrumentation; 4) Sensor Configuration; 5) Debris Hazard Monitoring; 6) Ascent Response Summary; 7) Response Signal; 8) Distribution of Flight Indications; 9) Probabilistic Risk Analysis (PRA); 10) Model Correlation; 11) Impact Tests; 12) Wing Leading Edge Modeling; 13) Ascent Debris PRA Results; and 14) MM/OD PRA Results.

Development of X-33/X-34 Aerothermodynamic Data Bases: Lessons Learned and Future Enhancements

NASA Technical Reports Server (NTRS)

Miller, C. G.

1999-01-01

A synoptic of programmatic and technical lessons learned in the development of aerothermodynamic data bases for the X-33 and X-34 programs is presented in general terms and from the perspective of the NASA Langley Research Center Aerothermodynamics Branch. The format used is that of the aerothermodynamic chain, the links of which are personnel, facilities, models/test articles, instrumentation, test techniques, and computational fluid dynamics (CFD). Because the aerodynamic data bases upon which the X-33 and X-34 vehicles will fly are almost exclusively from wind tunnel testing, as opposed to CFD, the primary focus of the lessons learned is on ground-based testing.

The Flow Field on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2008-11-01

The agility of the humpback whale has been attributed to the use of its pectoral flippers, on which protuberances are present along the leading edge. The forces and moments on hydrofoils with leading edge protuberances were measured in a water tunnel and were compared to a baseline NACA 63(4)-021 hydrofoil revealing significant performance differences. Three protuberance amplitudes and two spanwise wavelengths, closely resembling the morphology found in nature, were examined. Qualitative flow visualization techniques were used to examine flow patterns surrounding the hydrofoils, and Particle Image Velocimetry (PIV) was used to quantify these patterns. Flow visualizations have revealed counter-rotating vortices stemming from the shoulders of the protuberances. These streamwise vortices are a result of the spanwise pressure gradient brought about by the varying leading edge curvature. PIV was used to quantify the strength of these vortices as a function of angle of attack and leading edge geometry. At low angles of attack, these vortices are symmetric with respect to the protuberances; however, the symmetry is lost at high angles of attack. The loss of symmetry can be correlated with the separation point location on the hydrofoil.

Study of supersonic wings employing the attainable leading-edge thrust concept

NASA Technical Reports Server (NTRS)

Middleton, W. D.

1982-01-01

A theoretical study was made of supersonic wing geometries at Mach 1.8, using the attainable leading-edge thrust concept. The attainable thrust method offers a powerful means to improve overall aerodynamic efficiency by identifying wing leading-edge geometries that promote attached flow and by defining a local angle-of-attack range over which attached flow may be obtained. The concept applies to flat and to cambered wings, which leads to the consideration of drooped-wing leading edges for attached flow at high lift coefficients.

Cavitation on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid; Office of Naval Research Collaboration

2012-11-01

The effects of spanwise-uniform sinusoidal leading edge protuberances on the flow characteristics and forces of finite-span hydrofoils under vaporous cavitation conditions were examined experimentally over angles of attack ranging from -9° α <= 27°. Two planforms were studied, rectangular and swept, at a Reynolds number of ~ 720,000. Two protuberance wavelengths, λ = 0.25 c and 0.50 c, and three amplitudes, A = 0.025 c, 0.05 c, and 0.12 c, were examined as they resemble the humpback whale flipper morphology. All hydrofoils retain a mean NACA 634-021 profile. The forces and moments were measured at a freestream velocity of 7.2 m/s, and high-speed digital photography was used to capture flow field images at several angles of attack. The cavitation number corresponding to incipient leading edge cavitation was also calculated. As far as forces and cavitation number are concerned, results show that the baseline hydrofoil tends to have nearly equal or improved performance over the modified hydrofoils at most angles of attack tested. Flow images reveal that it is possible that the extent of sheet and tip vortex cavitation can be reduced with the introduction of leading edge protuberances. The forces and cavitation characteristics will be presented. Sponsored by the ONR-ULI program.

Computational aerothermodynamics

NASA Technical Reports Server (NTRS)

Deiwert, George S.

1989-01-01

Computational aerothermodynamics concerns the coupling of real gas effects with equations of motion to include thermochemical rate processes for chemical and energy exchange phenomena. These processes concern the creation and destruction of gas species by chemical reactions and the transfer of energy between the various species and between the various energy modes (e.g., translation, rotation, vibration, ionization, dissociation/recombination, etc.) of the species. To gain some insight into when such phenomena occur for current and future aerospace flight vehicles the author shows the flight regimes of some typical vehicles (e.g., Concord, aerospace plane, Space Shuttle, associated space transfer vehicles, Apollo entry vehicle, etc.) in terms of flight altitude and flight speed. Also indicated are regimes where chemical reactions such as dissociation and ionization are important and where nonequilibrium thermochemical phenomena are important.

Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

NASA Technical Reports Server (NTRS)

Arnold, James O.; Deiwert, G. S.

1997-01-01

The dream of producing an air-breathing, hydrogen fueled, hypervelocity aircraft has been before the aerospace community for decades. However, such a craft has not yet been realized, even in an experimental form. Despite the simplicity and beauty of the concept, many formidable problems must be overcome to make this dream a reality. This paper summarizes the aero/aerothermodynamic issues that must be addressed to make the dream a reality and discusses how aerothermodynamics facilities and their modem companion, real-gas computational fluid dynamics (CFD), can help solve the problems blocking the way to realizing the dream. The approach of the paper is first to outline the concept of an air-breathing hypersonic vehicle and then discuss the nose-to-tail aerothermodynamics issues and special aerodynamic problems that arise with such a craft. Then the utility of aerothermodynamic facilities and companion CFD analysis is illustrated by reviewing results from recent United States publications wherein these problems have been addressed. Papers selected for the discussion have k e n chosen such that the review will serve to survey important U.S. aero/aerothermodynamic real gas and conventional wind tunnel facilities that are useful in the study of hypersonic, hydrogen propelled hypervelocity vehicles.

Hypersonic aerospace vehicle leading edge cooling using heat pipe, transpiration and film cooling techniques

NASA Astrophysics Data System (ADS)

Modlin, James Michael

An investigation was conducted to study the feasibility of cooling hypersonic vehicle leading edge structures exposed to severe aerodynamic surface heat fluxes using a combination of liquid metal heat pipes and surface mass transfer cooling techniques. A generalized, transient, finite difference based hypersonic leading edge cooling model was developed that incorporated these effects and was demonstrated on an assumed aerospace plane-type wing leading edge section and a SCRAMJET engine inlet leading edge section. The hypersonic leading edge cooling model was developed using an existing, experimentally verified heat pipe model. Two applications of the hypersonic leading edge cooling model were examined. An assumed aerospace plane-type wing leading edge section exposed to a severe laminar, hypersonic aerodynamic surface heat flux was studied. A second application of the hypersonic leading edge cooling model was conducted on an assumed one-quarter inch nose diameter SCRAMJET engine inlet leading edge section exposed to both a transient laminar, hypersonic aerodynamic surface heat flux and a type 4 shock interference surface heat flux. The investigation led to the conclusion that cooling leading edge structures exposed to severe hypersonic flight environments using a combination of liquid metal heat pipe, surface transpiration, and film cooling methods appeared feasible.

Thermal management of tungsten leading edges in DIII-D

DOE PAGES

Nygren, Richard E.; Rudakov, Dmitry L.; Murphy, Christopher; ...

2017-04-29

The DiMES materials probe exposed tungsten blocks with 0.3 and 1 mm high leading edges to DIII-D He plasmas in 2015 and 2016 viewed with high resolution IRTV. The 1-mm edge may have reached >2400° C in a 3-s shot with a (parallel) heat load of ~50 MW/m 2 and ~10 MW/m 2 on the surface based on modeling. The experiments support ITER. Leading edges were also a concern in the DIII-D Metal Tile Experiment in 2016. Two toroidal rings of divertor tiles had W-coated molybdenum inserts 50 mm wide radially. This study presents data and thermal analyses.

Thermal management of tungsten leading edges in DIII-D

DOE Office of Scientific and Technical Information (OSTI.GOV)

Nygren, Richard E.; Rudakov, Dmitry L.; Murphy, Christopher

The DiMES materials probe exposed tungsten blocks with 0.3 and 1 mm high leading edges to DIII-D He plasmas in 2015 and 2016 viewed with high resolution IRTV. The 1-mm edge may have reached >2400° C in a 3-s shot with a (parallel) heat load of ~50 MW/m 2 and ~10 MW/m 2 on the surface based on modeling. The experiments support ITER. Leading edges were also a concern in the DIII-D Metal Tile Experiment in 2016. Two toroidal rings of divertor tiles had W-coated molybdenum inserts 50 mm wide radially. This study presents data and thermal analyses.

On the study of wavy leading-edge vanes to achieve low fan interaction noise

NASA Astrophysics Data System (ADS)

Tong, Fan; Qiao, Weiyang; Xu, Kunbo; Wang, Liangfeng; Chen, Weijie; Wang, Xunnian

2018-04-01

The application of wavy leading-edge vanes to reduce a single-stage axial fan noise is numerically studied. The aerodynamic and acoustic performance of the fan is numerically investigated using a hybrid unsteady Reynolds averaged Navier-Stokes (URANS)/acoustic analogy method (Goldstein equations). First, the hybrid URANS/Goldstein method is developed and successfully validated against experiment results. Next, numerical simulations are performed to investigate the noise reduction effects of the wavy leading-edge vanes. The aerodynamic and acoustic performance is assessed for a fan with vanes equipped with two different wavy leading-edge profiles and compared with the performance of conventional straight leading-edge vanes. Results indicate that a fan with wavy leading-edge vanes produces lower interaction noise than the baseline fan without a significant loss in aerodynamic performance. In fact, it is demonstrated that wavy leading-edge vanes have the potential to lead to both aerodynamic and acoustic improvements. The two different wavy leading-edge profiles are shown to successfully reduce the fan tone sound power level by 1.2 dB and 4.3 dB, respectively. Fan efficiency is also improved by about 1% with one of the tested wavy leading-edge profiles. Large eddy simulation (LES) is also performed for a simplified fan stage model to assess the effects of wavy leading-edge vanes on the broadband fan noise. Results indicate that the overall sound power level of a fan can be reduced by about 4 dB with the larger wavy leading-edge profile. Finally, the noise reduction mechanisms are investigated and analysed. It is found that the wavy leading-edge profiles can induce significant streamwise vorticity around the leading-edge protuberances and reduce pressure fluctuations (especially at locations of wavy leading-edge hills) and unsteady forces on the stator vanes. The underlying mechanism of the reduced pressure fluctuations is also discussed by examining the magnitude

Aerothermodynamic environment for a Titan probe with deployable decelerator

NASA Technical Reports Server (NTRS)

Green, M. J.; Swenson, B. L.; Balakrishnan, A.

1985-01-01

It is pointed out that further exploration of Titan, Saturn's largest moon, is of current interest to the scientific community, particularly from the standpoint of the organic chemical evolution of its atmosphere. For a suitable study of this Saturnian satellite, a mission involving a Titan atmospheric entry probe is to be conducted. The probe is to employ a deployable decelerator with the aim to allow scientific measurements in the haze layer. The present investigation is concerned with an assessment of the aerothermodynamic environment for the considered probe during its hypervelocity, low-Reynolds-number entry. Attention is given to the employed computational method, the Titan probe configuration, the Titan probe trajectory, the viscous-layer regime of the aerothermodynamic environment, and the incipient merged-layer regime.

Impact of ETO propellants on the aerothermodynamic analyses of propulsion components

NASA Technical Reports Server (NTRS)

Civinskas, K. C.; Boyle, R. J.; Mcconnaughey, H. V.

1988-01-01

The operating conditions and the propellant transport properties used in Earth-to-Orbit (ETO) applications affect the aerothermodynamic design of ETO turbomachinery in a number of ways. Some aerodynamic and heat transfer implications of the low molecular weight fluids and high Reynolds number operating conditions on future ETO turbomachinery are discussed. Using the current SSME high pressure fuel turbine as a baseline, the aerothermodynamic comparisons are made for two alternate fuel turbine geometries. The first is a revised first stage rotor blade designed to reduce peak heat transfer. This alternate design resulted in a 23 percent reduction in peak heat transfer. The second design concept was a single stage rotor to yield the same power output as the baseline two stage rotor. Since the rotor tip speed was held constant, the turbine work factor doubled. In this alternate design, the peak heat transfer remained the same as the baseline. While the efficiency of the single stage design was 3.1 points less than the baseline two stage turbine, the design was aerothermodynamically feasible, and may be structurally desirable.

Vortex interaction with a leading-edge of finite thickness

NASA Technical Reports Server (NTRS)

Sohn, D.; Rockwell, Donald

1987-01-01

Vortex interaction with a thick elliptical leading-edge at zero relative offset produces a pronounced secondary vortes of opposite sense that travels with the same phase speed as the primaty vortex along the lower surface of the edge. The edge thickness (scale) relative to the incident vorticity field has a strong effect on the distortion of the incident primary vortex during the impingement processs. When the thickness is sufficiently small, there is a definite severing of the incident vortex and the portion of the incident vortex that travels along the upper part of the elliptical surface has a considerably larger phase speed than that along the lower surface; this suggests that the integrated loading along the upper surface is more strongly correlated. When the thickness becomes too large, then most, if not all, of the incident vortex passes below the leading-edge. On the other hand, the relative tranverse offset of the edge with respect to the center of the incident vortex has a significant effect on the secondary vortex formation.

Theory of step on leading edge of negative corona current pulse

NASA Astrophysics Data System (ADS)

Gupta, Deepak K.; Mahajan, Sangeeta; John, P. I.

2000-03-01

Theoretical models taking into account different feedback source terms (e.g., ion-impact electron emission, photo-electron emission, field emission, etc) have been proposed for the existence and explanation of the shape of negative corona current pulse, including the step on the leading edge. In the present work, a negative corona current pulse with the step on the leading edge is obtained in the presence of ion-impact electron emission feedback source only. The step on the leading edge is explained in terms of the plasma formation process and enhancement of the feedback source. Ionization wave-like movement toward the cathode is observed after the step. The conditions for the existence of current pulse, with and without the step on the leading edge, are also described. A qualitative comparison with earlier theoretical and experimental work is also included.

Method and Apparatus for a Leading Edge Slat on a Wing of an Aircraft

NASA Technical Reports Server (NTRS)

Pitt, Dale M. (Inventor); Eckstein, Nicholas Stephen (Inventor)

2013-01-01

A method and apparatus for managing a flight control surface system. A leading edge device is moved on a leading edge from an undeployed position to a deployed position. The leading edge device has an outer surface, an inner surface, and a deformable fairing attached to the leading edge device such that the deformable fairing covers at least a portion of the inner surface. The deformable fairing changes from a deformed shape to an original shape when the leading edge device is moved to the deployed position. The leading edge device is then moved from the deployed position to the undeployed position, wherein the deformable fairing changes from the original shape to the deformed shape.

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

Robins, A. W.; Carlson, H. W.; Mack, R. J.

1980-01-01

Experimental/theoretical correlations are presented which show that significant levels of leading-edge thrust are possible at supersonic speeds for certain planforms having the geometry to support the theoretical thrust-distribution potential. The new analytical process employed provides not only the level of leading-edge thrust attainable but also the spanwise distribution of both it and that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.

Simulated airline service experience with laminar-flow control leading-edge systems

NASA Technical Reports Server (NTRS)

Maddalon, Dal V.; Fisher, David F.; Jennett, Lisa A.; Fischer, Michael C.

1987-01-01

The first JetStar leading edge flight test was made November 30, 1983. The JetStar was flown for more than 3 years. The titanium leading edge test articles today remain in virtually the same condition as they were in on that first flight. No degradation of laminar flow performance has occurred as a result of service. The JetStar simulated airline service flights have demonstrated that effective, practical leading edge systems are available for future commercial transports. Specific conclusions based on the results of the simulated airline service test program are summarized.

Air-Cooled Turbine Blades with Tip Cap For Improved Leading-Edge Cooling

NASA Technical Reports Server (NTRS)

Calvert, Howard F.; Meyer, Andre J., Jr.; Morgan, William C.

1959-01-01

An investigation was conducted in a modified turbojet engine to determine the cooling characteristics of the semistrut corrugated air- cooled turbine blade and to compare and evaluate a leading-edge tip cap as a means for improving the leading-edge cooling characteristics of cooled turbine blades. Temperature data were obtained from uncapped air-cooled blades (blade A), cooled blades with the leading-edge tip area capped (blade B), and blades with slanted corrugations in addition to leading-edge tip caps (blade C). All data are for rated engine speed and turbine-inlet temperature (1660 F). A comparison of temperature data from blades A and B showed a leading-edge temperature reduction of about 130 F that could be attributed to the use of tip caps. Even better leading-edge cooling was obtained with blade C. Blade C also operated with the smallest chordwise temperature gradients of the blades tested, but tip-capped blade B operated with the lowest average chordwise temperature. According to a correlation of the experimental data, all three blade types 0 could operate satisfactorily with a turbine-inlet temperature of 2000 F and a coolant flow of 3 percent of engine mass flow or less, with an average chordwise temperature limit of 1400 F. Within the range of coolant flows investigated, however, only blade C could maintain a leading-edge temperature of 1400 F for a turbine-inlet temperature of 2000 F.

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2017-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

Design modification of airfoil by integrating sinusoidal leading edge and dimpled surface

NASA Astrophysics Data System (ADS)

Masud, M. H.; Naim-Ul-Hasan, Arefin, Amit Md. Estiaque; Joardder, Mohammad U. H.

2017-06-01

Airfoil is widely used for aircraft wings and blades of helicopters, turbines, propellers, fans and compressors. Many researches have been conducted on focusing the leading edge, surface and trailing edge of airfoil in order to maximize airfoil lift and to reduce drag. Literature shows that using protuberances along the leading edge of NACA 2412, it is possible to attain better performance from the baseline. Besides, the inward dimpled surface of NACA 0018 produces lesser drag at a positive angle of attacks. However, there is no literature that integrates sinusoidal leading edge and dimpled to attain the benefits of the both. In this study, simulation has been done for design improvement of airfoil by integrating sinusoidal leading edge and dimpled surface. Simulations have been run using finite element method environment. Significant improvement has been observed from the simulation results.

Laminar flow control leading edge glove flight test article development

NASA Technical Reports Server (NTRS)

Pearce, W. E.; Mcnay, D. E.; Thelander, J. A.

1984-01-01

A laminar flow control (LFC) flight test article was designed and fabricated to fit into the right leading edge of a JetStar aircraft. The article was designed to attach to the front spar and fill in approx. 70 inches of the leading edge that are normally occupied by the large slipper fuel tank. The outer contour of the test article was constrained to align with an external fairing aft of the front spar which provided a surface pressure distribution over the test region representative of an LFC airfoil. LFC is achieved by applying suction through a finely perforated surface, which removes a small fraction of the boundary layer. The LFC test article has a retractable high lift shield to protect the laminar surface from contamination by airborne debris during takeoff and low altitude operation. The shield is designed to intercept insects and other particles that could otherwise impact the leading edge. Because the shield will intercept freezing rain and ice, a oozing glycol ice protection system is installed on the shield leading edge. In addition to the shield, a liquid freezing point depressant can be sprayed on the back of the shield.

Aerothermodynamic Design of the Mars Science Laboratory Heatshield

NASA Technical Reports Server (NTRS)

Edquist, Karl T.; Dyakonov, Artem A.; Wright, Michael J.; Tang, Chun Y.

2009-01-01

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule heatshield. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux, shear stress, and pressure) entry trajectories from a 2009 launch. Boundary layer transition is expected prior to peak heat flux, a first for Mars entry, and the heatshield environments were defined for a fully-turbulent heat pulse. The effects of distributed surface roughness on turbulent heat flux and shear stress peaks are included using empirical correlations. Additional biases and uncertainties are based on computational model comparisons with experimental data and sensitivity studies. The peak design conditions are 197 W/sq cm for heat flux, 471 Pa for shear stress, 0.371 Earth atm for pressure, and 5477 J/sq cm for total heat load. Time-varying conditions at fixed heatshield locations were generated for thermal protection system analysis and flight instrumentation development. Finally, the aerothermodynamic effects of delaying launch until 2011 are previewed.

Leading edge flap system for aircraft control augmentation

NASA Technical Reports Server (NTRS)

Rao, D. M. (Inventor)

1984-01-01

Traditional roll control systems such as ailerons, elevons or spoilers are least effective at high angles of attack due to boundary layer separation over the wing. This invention uses independently deployed leading edge flaps on the upper surfaces of vortex stabilized wings to shift the center of lift outboard. A rolling moment is created that is used to control roll in flight at high angles of attack. The effectiveness of the rolling moment increases linearly with angle of attack. No adverse yaw effects are induced. In an alternate mode of operation, both leading edge flaps are deployed together at cruise speeds to create a very effective airbrake without appreciable modification in pitching moment. Little trim change is required.

Evaluation of cloud detection instruments and performance of laminar-flow leading-edge test articles during NASA Leading-Edge Flight-Test Program

NASA Technical Reports Server (NTRS)

Davis, Richard E.; Maddalon, Dal V.; Wagner, Richard D.; Fisher, David F.; Young, Ronald

1989-01-01

Summary evaluations of the performance of laminar-flow control (LFC) leading edge test articles on a NASA JetStar aircraft are presented. Statistics, presented for the test articles' performance in haze and cloud situations, as well as in clear air, show a significant effect of cloud particle concentrations on the extent of laminar flow. The cloud particle environment was monitored by two instruments, a cloud particle spectrometer (Knollenberg probe) and a charging patch. Both instruments are evaluated as diagnostic aids for avoiding laminar-flow detrimental particle concentrations in future LFC aircraft operations. The data base covers 19 flights in the simulated airline service phase of the NASA Leading-Edge Flight-Test (LEFT) Program.

Aerothermodynamics of the Mars Global Surveyor Spacecraft

NASA Technical Reports Server (NTRS)

Shane, Russell W.; Tolson, Robert H.

1998-01-01

The aerothermodynamics characteristics of the Mars Global Surveyor spacecraft are investigated and reported. These results have been used by the Mars Global Surveyor mission planners to design the aerobraking phase of the mission. Analytical and Direct Simulation Monte Carlo computer codes were used with a detailed, three dimensional model of the spacecraft to evaluate spacecraft aerobraking characteristics for flight in free molecular and transitional flow regimes. The spacecraft is found to be aerodynamically stable in aerobraking and planned contingency configurations. Aerodynamic forces, moments, and heating are found to be highly dependent on atmospheric density. Accommodation coefficient. is seen to strongly influence drag coefficient. Transitional flow effects are found to reduce overall solar panel heating. Attitude control thruster plumes are shown to interact with the freestream, diminishing the effectiveness of the attitude control system and even leading to thrust reversal. These plume-freestream interaction effects are found to be highly dependent on freestream density.

The fish tail motion forms an attached leading edge vortex

PubMed Central

Borazjani, Iman; Daghooghi, Mohsen

2013-01-01

The tail (caudal fin) is one of the most prominent characteristics of fishes, and the analysis of the flow pattern it creates is fundamental to understanding how its motion generates locomotor forces. A mechanism that is known to greatly enhance locomotor forces in insect and bird flight is the leading edge vortex (LEV) reattachment, i.e. a vortex (separation bubble) that stays attached at the leading edge of a wing. However, this mechanism has not been reported in fish-like swimming probably owing to the overemphasis on the trailing wake, and the fact that the flow does not separate along the body of undulating swimmers. We provide, to our knowledge, the first evidence of the vortex reattachment at the leading edge of the fish tail using three-dimensional high-resolution numerical simulations of self-propelled virtual swimmers with different tail shapes. We show that at Strouhal numbers (a measure of lateral velocity to the axial velocity) at which most fish swim in nature (approx. 0.25) an attached LEV is formed, whereas at a higher Strouhal number of approximately 0.6 the LEV does not reattach. We show that the evolution of the LEV drastically alters the pressure distribution on the tail and the force it generates. We also show that the tail's delta shape is not necessary for the LEV reattachment and fish-like kinematics is capable of stabilising the LEV. Our results suggest the need for a paradigm shift in fish-like swimming research to turn the focus from the trailing edge to the leading edge of the tail. PMID:23407826

Numerical Investigations of the Influence of Unsteady Vane Trailing Edge Shock Wave on Film Cooling Effectiveness of Rotor Blade Leading Edge

NASA Astrophysics Data System (ADS)

Wang, Yufeng; Cai, Le; Wang, Songtao; Zhou, Xun

2018-04-01

Unsteady numerical simulations of a high-load transonic turbine stage have been carried out to study the influences of vane trailing edge outer-extending shockwave on rotor blade leading edge film cooling performance. The turbine stage used in this paper is composed of a vane section and a rotor one which are both near the root section of a transonic high-load turbine stage. The Mach number is 0.94 at vane outlet, and the relative Mach number is above 1.10 at rotor outlet. Various positions and oblique angles of film cooling holes were investigated in this research. Results show that the cooling efficiency on the blade surface of rotor near leading edge is significantly affected by vane trailing edge outer-extending shockwave in some cases. In the cases that film holes are close to leading edge, cooling performance suffers more from the sweeping vane trailing edge outer-extending shockwave. In addition, coolant flow ejected from oblique film holes is harder to separate from the blade surface of rotor, and can cover more blade area even under the effects of sweeping vane trailing edge shockwave. As a result, oblique film holes can provide better film cooling performance than vertical film holes do near the leading edge on turbine blade which is swept by shockwaves.

Numerical Simulations Of High-Altitude Aerothermodynamics Of A Prospective Spacecraft Model

NASA Astrophysics Data System (ADS)

Vashchenkov, P. V.; Kaskovsky, A. V.; Krylov, A. N.; Ivanov, M. S.

2011-05-01

The paper describes the computations of aerothermodynamic characteristics of a promising spacecraft (Prospective Piloted Transport System) along its de- scent trajectory at altitudes from 120 to 60 km. The computations are performed by the DSMC method with the use of the SMILE software system and by the engineering technique (local bridging method) with the use of the RuSat software system. The influence of real gas effects (excitation of rotational and vibrational energy modes and chemical reactions) on aerothermodynamic characteristics of the vehicle is studied. A comparison of results obtained by the approximate engineering method and the DSMC method allow the accuracy of prediction of aerodynamic characteristics by the local bridging method to be estimated.

The leading-edge vortex of swift-wing shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the Leading-Edge Vortex (LEV) for lift generation in a variety of flight conditions. In this investigation, a model non-slender delta shaped wing with a sharp leading-edge is tested at low Reynolds Number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the un-modified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift-wing shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds Number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta shaped wing. This work received funding from the Engineering and Physical Sciences Research Council [EP/M506515/1] and the Consejo Nacional de Ciencia y Tecnología (CONACYT).

Fabrication and Testing of a Leading-Edge-Shaped Heat Pipe

NASA Technical Reports Server (NTRS)

Glass, David E.; Merrigan, Michael A.; Sena, J. Tom; Reid, Robert S.

1998-01-01

The development of a refractory-composite/heat-pipe-cooled leading edge has evolved from the design stage to the fabrication and testing of a full size, leading-edge-shaped heat pipe. The heat pipe had a 'D-shaped' cross section and was fabricated from arc cast Mo-4lRe. An artery was included in the wick. Several issues were resolved with the fabrication of the sharp leading edge radius heat pipe. The heat pipe was tested in a vacuum chamber at Los Alamos National Laboratory using induction heating and was started up from the frozen state several times. However, design temperatures and heat fluxes were not obtained due to premature failure of the heat pipe resulting from electrical discharge between the induction heating apparatus and the heat pipe. Though a testing anomaly caused premature failure of the heat pipe, successful startup and operation of the heat pipe was demonstrated.

A flight test of laminar flow control leading-edge systems

NASA Technical Reports Server (NTRS)

Fischer, M. C.; Wright, A. S., Jr.; Wagner, R. D.

1983-01-01

NASA's program for development of a laminar flow technology base for application to commercial transports has made significant progress since its inception in 1976. Current efforts are focused on development of practical reliable systems for the leading-edge region where the most difficult problems in applying laminar flow exist. Practical solutions to these problems will remove many concerns about the ultimate practicality of laminar flow. To address these issues, two contractors performed studies, conducted development tests, and designed and fabricated fully functional leading-edge test articles for installation on the NASA JetStar aircraft. Systems evaluation and performance testing will be conducted to thoroughly evaluate all system capabilities and characteristics. A simulated airline service flight test program will be performed to obtain the operational sensitivity, maintenance, and reliability data needed to establish that practical solutions exist for the difficult leading-edge area of a future commercial transport employing laminar flow control.

Improved Method for Prediction of Attainable Wing Leading-Edge Thrust

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; McElroy, Marcus O.; Lessard, Wendy B.; McCullers, L. Arnold

1996-01-01

Prediction of the loss of wing leading-edge thrust and the accompanying increase in drag due to lift, when flow is not completely attached, presents a difficult but commonly encountered problem. A method (called the previous method) for the prediction of attainable leading-edge thrust and the resultant effect on airplane aerodynamic performance has been in use for more than a decade. Recently, the method has been revised to enhance its applicability to current airplane design and evaluation problems. The improved method (called the present method) provides for a greater range of airfoil shapes from very sharp to very blunt leading edges. It is also based on a wider range of Reynolds numbers than was available for the previous method. The present method, when employed in computer codes for aerodynamic analysis, generally results in improved correlation with experimental wing-body axial-force data and provides reasonable estimates of the measured drag.

Development of a cyber physical apparatus for investigating fluid structure interaction on leading edge vortex evolution

NASA Astrophysics Data System (ADS)

Raghu Gowda, Belagumba Venkatachalaiah

This dissertation examines how simple structural compliance impacts a specific transient vortex phenomenon that occurs on high angle of attack lifting surfaces termed dynamic stall. In many Fluid structure interaction (FSI) research efforts, a purely physical or purely computational approach is taken. In this work a low cost cyber-physical (CPFD) system is designed and developed for representing the FSI in the leading edge vortex (LEV) development problem. The leading edge compliance appears to be favorable in a specific spring constant range for a given wing. When the leading edge compliance prescribed via CPFD system is too low compared with the moment due to dynamic pressure or fluid unsteady effect, the LEV behavior is similar to that of a rigid wing system. When the leading edge compliance is too high, excessive compliance is introduced into the wing system and the leading edge vortex evolution is affected by the large change in wing angle. At moderate leading edge compliance, a balance appears to be achieved in which the leading edge vorticity shedding rate supports the long term evolution of the leading edge vortex. Further investigation is required to determine specific parameters governing these leading edge compliance ranges.

Experimental And Numerical Study Of CMC Leading Edges In Hypersonic Flows

NASA Astrophysics Data System (ADS)

Kuhn, Markus; Esser, Burkard; Gulhan, Ali; Dalenbring, Mats; Cavagna, Luca

2011-05-01

Future transportation concepts aim at high supersonic or hypersonic speeds, where the formerly sharp boundaries between aeronautic and aerospace applications become blurred. One of the major issues involved to high speed flight are extremely high aerothermal loads, which especially appear at the leading edges of the plane’s wings and at sharp edged air intake components of the propulsion system. As classical materials like metals or simple ceramics would thermally and structurally fail here, new materials have to be applied. In this context, lightweight ceramic matrix composites (CMC) seem to be prospective candidates as they are high-temperature resistant and offer low thermal expansion along with high specific strength at elevated temperature levels. A generic leading edge model with a ceramic wing assembly with a sweep back angle of 53° was designed, which allowed for easy leading edge sample integration of different CMC materials. The samples consisted of the materials C/C-SiC (non-oxide), OXIPOL and WHIPOX (both oxide) with a nose radius of 2 mm. In addition, a sharp edged C/C-SiC sample was prepared to investigate the nose radius influence. Overall, 13 thermocouples were installed inside the entire model to measure the temperature evolution at specific locations, whereby 5 thermocouples were placed inside the leading edge sample itself. In addition, non-intrusive techniques were applied for surface temperature measurements: An infrared camera was used to measure the surface temperature distribution and at specific spots, the surface temperature was also measured by pyrometers. Following, the model was investigated in DLR’s arc-heated facility L3K at a total enthalpy of 8.5 MJ/kg, Mach number of 7.8, different angles of attack and varying wing inclination angles. These experiments provide a sound basis for the simulation of aerothermally loaded CMC leading edge structures. Such fluid-structure coupled approaches have been performed by FOI, basing on a

Aerothermal Performance Constraint Analysis of Sharp Nosecaps and Leading Edges

NASA Technical Reports Server (NTRS)

Rizk, Yehia; Gee, Ken

2004-01-01

The main objective of this work is to predict the Aerothermal Performance Constraint (APC) for a class of Crew Transfer Vehicles (CTV) with shap noses and wing leading edges made out of UHTC which is a family of Ultra High Temperature Ceramics materials developed at NASA Ames. The APC is based on the theoretical temperature limit of the material which is usually encountered at the CTV nose or wing leading edge. The APC places a lower limit on the trajectory of the CTV in the altitude velocity space. The APC is used as one of the constraints in developing reentry and abort trajectories for the CTV. The trajectories are then used to generate transient thermal response of the nosecaps and wing leading edges which are represented as either a one piece of UHTC or two piece (UHTC + RCC) with perfect axial contact. The final paper will include more details about the analysis procedure and will also include results for reentry and abort design trajectories.

X-38 NASA/DLR/ESA-Dassault Aviation Integrated Aerodynamic and Aerothermodynamic Activities

NASA Technical Reports Server (NTRS)

Labbe, Steve G.; Perez, Leo F.; Fitzgerald, Steve; Longo, Jose; Rapuc, Marc; Molina, Rafael; Nicholson, Leonard S. (Technical Monitor)

1999-01-01

The characterization of the aeroshape selected for the X-38 [Crew Return Vehicle (CRV) demonstrator] is presently being performed as a cooperative endeavour between NASA, DLR (through its TETRA Program), and European Space Agency (ESA) with Dassault Aviation integrating the aerodynamic and aerothermodynamic activities. The methodologies selected for characterizing the aerodynamic and aerothermodynamic environment of the X-38 are presented. Also, the implications for related disciplines such as Guidance Navigation and Control (GN&C) with its corresponding Flight Control System (FCS), Structural, and Thermal Protection System (TPS) design are discussed. An attempt is made at defining the additional activities required to support the design of a derived operational CRV.

Hypersonic Engine Leading Edge Experiments in a High Heat Flux, Supersonic Flow Environment

NASA Technical Reports Server (NTRS)

Gladden, Herbert J.; Melis, Matthew E.

1994-01-01

A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Three aerothermal load related concerns are the boundary layer transition from laminar to turbulent flow, articulating panel seals in high temperature environments, and strut (or cowl) leading edges with shock-on-shock interactions. A multidisciplinary approach is required to address these technical concerns. A hydrogen/oxygen rocket engine heat source has been developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to experimentally evaluate the heat transfer and structural response of the strut (or cowl) leading edge. A recent experimental program conducted in this facility is discussed and related to cooling technology capability. The specific objective of the experiment discussed is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Heat transfer analyses of a similar leading edge concept cooled with gaseous hydrogen is included to demonstrate the complexity of the problem resulting from plastic deformation of the structures. Macro-photographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight.

Effects of leading-edge devices on the low-speed aerodynamic characteristics of a highly-swept arrow-wing

NASA Technical Reports Server (NTRS)

Scott, S. J.; Nicks, O. W.; Imbrie, P. K.

1985-01-01

An investigation was conducted in the Texas A&M University 7 by 10 foot Low Speed Wind Tunnel to provide a direct comparison of the effect of several leading edge devices on the aerodynamic performance of a highly swept wing configuration. Analysis of the data indicates that for the configuration with undeflected leading edges, vortex separation first occurs on the outboard wing panel for angles of attack of approximately 2, and wing apex vorticies become apparent for alpha or = 4 deg. However, the occurrence of the leading edge vortex flow may be postponed with leading edge devices. Of the devices considered, the most promising were a simple leading edge deflection of 30 deg and a leading edge slat system. The trailing edge flap effectiveness was found to be essentially the same for the configuration employing either of these more promising leading edge devices. Analysis of the lateral directional data showed that for all of the concepts considered, deflecting leading edge downward in an attempt to postpone leading edge vortex flows, has the favorable effect of reducing the effective dihedral.

Membrane tension controls adhesion positioning at the leading edge of cells

PubMed Central

Pontes, Bruno; Gole, Laurent; Kosmalska, Anita Joanna; Tam, Zhi Yang; Luo, Weiwei; Kan, Sophie; Viasnoff, Virgile; Roca-Cusachs, Pere; Tucker-Kellogg, Lisa

2017-01-01

Cell migration is dependent on adhesion dynamics and actin cytoskeleton remodeling at the leading edge. These events may be physically constrained by the plasma membrane. Here, we show that the mechanical signal produced by an increase in plasma membrane tension triggers the positioning of new rows of adhesions at the leading edge. During protrusion, as membrane tension increases, velocity slows, and the lamellipodium buckles upward in a myosin II–independent manner. The buckling occurs between the front of the lamellipodium, where nascent adhesions are positioned in rows, and the base of the lamellipodium, where a vinculin-dependent clutch couples actin to previously positioned adhesions. As membrane tension decreases, protrusion resumes and buckling disappears, until the next cycle. We propose that the mechanical signal of membrane tension exerts upstream control in mechanotransduction by periodically compressing and relaxing the lamellipodium, leading to the positioning of adhesions at the leading edge of cells. PMID:28687667

Airfoil noise reductions through leading edge serrations

NASA Astrophysics Data System (ADS)

Narayanan, S.; Chaitanya, P.; Haeri, S.; Joseph, P.; Kim, J. W.; Polacsek, C.

2015-02-01

This paper provides an experimental investigation into the use of leading edge (LE) serrations as a means of reducing the broadband noise generated due to the interaction between the aerofoil's LE and impinging turbulence. Experiments are performed on a flat plate in an open jet wind tunnel. Grids are used to generate isotropic homogeneous turbulence. The leading edge serrations are in the form of sinusoidal profiles of wavelengths, λ, and amplitudes, 2h. The frequency and amplitude characteristics are studied in detail in order to understand the effect of LE serrations on noise reduction characteristics and are compared with straight edge baseline flat plates. Noise reductions are found to be insignificant at low frequencies but significant in the mid frequency range (500 Hz-8 kHz) for all the cases studied. The flat plate results are also compared to the noise reductions obtained on a serrated NACA-65 aerofoil with the same serration profile. Noise reductions are found to be significantly higher for the flat plates with a maximum noise reduction of around 9 dB compared with about 7 dB for the aerofoil. In general, it is observed that the sound power reduction level (ΔPWL) is sensitive to the amplitude, 2h of the LE serrations but less sensitive to the serration wavelength, λ. Thus, this paper sufficiently demonstrates that the LE amplitude acts as a key parameter for enhancing the noise reduction levels in flat plates and aerofoils.

Development of X-33/X-34 Aerothermodynamic Data Bases: Lessons Learned and Future Enhancements

NASA Technical Reports Server (NTRS)

Miller, C. G.

2000-01-01

A synoptic of programmatic and technical lessons learned in the development of aerothermodynamic data bases for the X-33 and X-34 programs is presented in general terms and from the perspective of the NASA Langley Research Center Aerothermodynamics Branch. The format used is that of the "aerothermodynamic chain," the links of which are personnel, facilities, models/test articles, instrumentation, test techniques, and computational fluid dynamics (CFD). Because the aerodynamic data bases upon which the X-33 and X-34 vehicles will fly are almost exclusively from wind tunnel testing, as opposed to CFD, the primary focus of the lessons learned is on ground-based testing. The period corresponding to the development of X-33 and X-34 aerothermodynamic data bases was challenging, since a number of other such programs (e.g., X-38, X-43) competed for resources at a time of downsizing of personnel, facilities, etc., outsourcing, and role changes as NASA Centers served as subcontractors to industry. The impact of this changing environment is embedded in the lessons learned. From a technical perspective, the relatively long times to design and fabricate metallic force and moment models, delays in delivery of models, and a lack of quality assurance to determine the fidelity of model outer mold lines (OML) prior to wind tunnel testing had a major negative impact on the programs. On the positive side, the application of phosphor thermography to obtain global, quantitative heating distributions on rapidly fabricated ceramic models revolutionized the aerothermodynamic optimization of vehicle OMLs, control surfaces, etc. Vehicle designers were provided with aeroheating information prior to, or in conjunction with, aerodynamic information early in the program, thereby allowing trades to be made with both sets of input; in the past only aerodynamic data were available as input. Programmatically, failure to include transonic aerodynamic wind tunnel tests early in the assessment phase

Effect of leading-edge geometry on boundary-layer receptivity to freestream sound

NASA Technical Reports Server (NTRS)

Lin, Nay; Reed, Helen L.; Saric, W. S.

1991-01-01

The receptivity to freestream sound of the laminar boundary layer over a semi-infinite flat plate with an elliptic leading edge is simulated numerically. The incompressible flow past the flat plate is computed by solving the full Navier-Stokes equations in general curvilinear coordinates. A finite-difference method which is second-order accurate in space and time is used. Spatial and temporal developments of the Tollmien-Schlichting wave in the boundary layer, due to small-amplitude time-harmonic oscillations of the freestream velocity that closely simulate a sound wave travelling parallel to the plate, are observed. The effect of leading-edge curvature is studied by varying the aspect ratio of the ellipse. The boundary layer over the flat plate with a sharper leading edge is found to be less receptive. The relative contribution of the discontinuity in curvature at the ellipse-flat-plate juncture to receptivity is investigated by smoothing the juncture with a polynomial. Continuous curvature leads to less receptivity. A new geometry of the leading edge, a modified super ellipse, which provides continuous curvature at the juncture with the flat plate, is used to study the effect of continuous curvature and inherent pressure gradient on receptivity.

Aerothermodynamics of blunt body entry vehicles

NASA Astrophysics Data System (ADS)

Hollis, Brian R.; Borrelli, Salvatore

2012-01-01

In this chapter, the aerothermodynamic phenomena of blunt body entry vehicles are discussed. Four topics will be considered that present challenges to current computational modeling techniques for blunt body environments: turbulent flow, non-equilibrium flow, rarefied flow, and radiation transport. Examples of comparisons between computational tools to ground and flight-test data will be presented in order to illustrate the challenges existing in the numerical modeling of each of these phenomena and to provide test cases for evaluation of computational fluid dynamics (CFD) code predictions.

The effects of leading edge modifications on the post-stall characteristics of wings

NASA Technical Reports Server (NTRS)

Winkelmann, A. E.; Barlow, J. B.; Saini, J. K.; Anderson, J. D., Jr.; Jones, E.

1980-01-01

An investigation of the effects of leading edge modifications on the post-stall characteristics of two rectangular planform wings in a series of low speed wind tunnel tests is presented. Abrupt discontinuities in the leading edge shape of the wings were produced by placing a nose glove over a portion of the span or by deflecting sections of a segmented leading edge flap. Six component balance data, oil flow visualization photographs, and pressure distribution measurements were obtained, and tests made to study the development of flow separation at stall on small scale planform wing models. Results of oil flow visualization tests at and beyond stall showed the formation of counter-rotating swirl patterns on the upper surface of the '2-D' and '3-D' wings, and results of a numerical lifting line technique applied to wings with leading edge modifications are included.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1985-01-01

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Experimental Investigation of Dynamic Stall on an Airfoil with Leading Edge Tubercles

NASA Astrophysics Data System (ADS)

Hrynuk, John; Bohl, Douglas

2013-11-01

Humpback whales are unique in that their flippers have leading edge ``bumps'' or tubercles. Past work on airfoils modeled after whale flippers has centered on the static aerodynamic characteristics of these airfoils. In the current work, NACA 0012 airfoils modified with leading edge tubercles are investigated to determine the effect of the tubercles on the dynamic characteristics, specifically on dynamic stall vortex formation, of the airfoils. Molecular Tagging Velocimetry (MTV) is used to measure the flow field around the modified airfoils at nondimensional pitch rates of Ω = 0.1, 0.2, and 0.4. The results show that the characteristics of the dynamics stall vortex are dependent on the location relative to the peak or valley of the leading edge bumps. These characteristics are also found to be different than those observed in dynamic stall on a smooth leading edge airfoil. In specific, the location of the dynamic stall vortex appears to form further aft on the airfoil for the tubercle case versus the smooth case. This work supported by NSF Grant # 0845882.

The leading-edge vortex of swift wing-shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

The leading-edge vortex of swift wing-shaped delta wings

PubMed Central

Muir, Rowan Eveline; Arredondo-Galeana, Abel

2017-01-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing. PMID:28878968

The leading-edge vortex of swift wing-shaped delta wings.

PubMed

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus . The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

Spanwise visualization of the flow around a three-dimensional foil with leading edge protuberances

NASA Astrophysics Data System (ADS)

Stanway, M. J.; Techet, A. H.

2006-11-01

Studies of model humpback whale fins have shown that leading edge protuberances, or tubercles, can lead to delayed stall and increased lift at higher angles of attack, compared to foils with geometrically smooth leading edges. Such enhanced performance characteristics could prove highly useful in underwater vehicles such as gliders or long range AUVs (autonomous underwater vehicles). In this work, Particle Imaging Velocimetry (PIV) is performed on two static wings in a water tunnel over a range of angles of attack. These three- dimensional, finite-aspect ratio wings are modeled after a humpback whale flipper and are identical in shape, tapered from root to tip, except for the leading edge. In one of the foils the leading edge is smooth, whereas in the other, regularly spaced leading edge bumps are machined to simulate the whale’s fin tubercles. Results from these PIV tests reveal distinct cells where coherent flow structures are destroyed as a result of the leading edge perturbations. Tests are performed at Reynolds numbers Re ˜ O(10^5), based on chordlength, in a recirculating water tunnel. An inline six-axis load cell is mounted to measure the forces on the foil over a range of static pitch angles. It is hypothesized that this spanwise breakup of coherent vortical structures is responsible for the delayed angle of stall. These quantitative experiments complement exiting qualitative studies with two dimensional foils.

Turbine Airfoil With CMC Leading-Edge Concept Tested Under Simulated Gas Turbine Conditions

NASA Technical Reports Server (NTRS)

Robinson, R. Craig; Hatton, Kenneth S.

2000-01-01

Silicon-based ceramics have been proposed as component materials for gas turbine engine hot-sections. When the Navy s Harrier fighter experienced engine (Pegasus F402) failure because of leading-edge durability problems on the second-stage high-pressure turbine vane, the Office of Naval Research came to the NASA Glenn Research Center at Lewis Field for test support in evaluating a concept for eliminating the vane-edge degradation. The High Pressure Burner Rig (HPBR) was selected for testing since it could provide temperature, pressure, velocity, and combustion gas compositions that closely simulate the engine environment. The study focused on equipping the stationary metal airfoil (Pegasus F402) with a ceramic matrix composite (CMC) leading-edge insert and evaluating the feasibility and benefits of such a configuration. The test exposed the component, with and without the CMC insert, to the harsh engine environment in an unloaded condition, with cooling to provide temperature relief to the metal blade underneath. The insert was made using an AlliedSignal Composites, Inc., enhanced HiNicalon (Nippon Carbon Co. LTD., Yokohama, Japan) fiber-reinforced silicon carbide composite (SiC/SiC CMC) material fabricated via chemical vapor infiltration. This insert was 45-mils thick and occupied a recessed area in the leading edge and shroud of the vane. It was designed to be free floating with an end cap design. The HPBR tests provided a comparative evaluation of the temperature response and leading-edge durability and included cycling the airfoils between simulated idle, lift, and cruise flight conditions. In addition, the airfoils were aircooled, uniquely instrumented, and exposed to the exact set of internal and external conditions, which included gas temperatures in excess of 1370 C (2500 F). In addition to documenting the temperature response of the metal vane for comparison with the CMC, a demonstration of improved leading-edge durability was a primary goal. First, the

Acoustic Receptivity of Mach 4.5 Boundary Layer with Leading- Edge Bluntness

NASA Technical Reports Server (NTRS)

Malik, Mujeeb R.; Balakumar, Ponnampalam

2007-01-01

Boundary layer receptivity to two-dimensional slow and fast acoustic waves is investigated by solving Navier-Stokes equations for Mach 4.5 flow over a flat plate with a finite-thickness leading edge. Higher order spatial and temporal schemes are employed to obtain the solution whereby the flat-plate leading edge region is resolved by providing a sufficiently refined grid. The results show that the instability waves are generated in the leading edge region and that the boundary-layer is much more receptive to slow acoustic waves (by almost a factor of 20) as compared to the fast waves. Hence, this leading-edge receptivity mechanism is expected to be more relevant in the transition process for high Mach number flows where second mode instability is dominant. Computations are performed to investigate the effect of leading-edge thickness and it is found that bluntness tends to stabilize the boundary layer. Furthermore, the relative significance of fast acoustic waves is enhanced in the presence of bluntness. The effect of acoustic wave incidence angle is also studied and it is found that the receptivity of the boundary layer on the windward side (with respect to the acoustic forcing) decreases by more than a factor of 4 when the incidence angle is increased from 0 to 45 deg. However, the receptivity coefficient for the leeward side is found to vary relatively weakly with the incidence angle.

Space shuttle orbiter leading-edge flight performance compared to design goals

NASA Technical Reports Server (NTRS)

Curry, D. M.; Johnson, D. W.; Kelly, R. E.

1983-01-01

Thermo-structural performance of the Space Shuttle orbiter Columbia's leading-edge structural subsystem for the first five (5) flights is compared with the design goals. Lessons learned from thse initial flights of the first reusable manned spacecraft are discussed in order to assess design maturity, deficiencies, and modifications required to rectify the design deficiencies. Flight data and post-flight inspections support the conclusion that the leading-edge structural subsystem hardware performance was outstanding for the initial five (5) flights.

Experimental evaluation of joint designs for a space-shuttle orbiter ablative leading edge

NASA Technical Reports Server (NTRS)

Tompkins, S. S.; Kabana, W. P.

1975-01-01

The thermal performance of two types of ablative leading-edge joints for a space-shuttle orbiter were tested and evaluated. Chordwise joints between ablative leading-edge segments, and spanwise joints between ablative leading-edge segments and reusable surface insulation tiles were exposed to simulated shuttle heating environments. The data show that the thermal performance of models with chordwise joints to be as good as jointless models in simulated ascent-heating and orbital cold-soak environments. The suggestion is made for additional work on the joint seals, and, in particular, on the effects of heat-induced seal-material surface irregularities on the local flow.

A theoretical investigation of the aerodynamics of low-aspect-ratio wings with partial leading-edge separation

NASA Technical Reports Server (NTRS)

Mehrotra, S. C.; Lan, C. E.

1978-01-01

A numerical method is developed to predict distributed and total aerodynamic characteristics for low aspect-ratio wings with partial leading-edge separation. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the quasi-vortex-lattice method. The leading-edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at mid-points to satisfy the force free condition. The wake behind the trailing-edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading- and trailing-edges. Comparison of the predicted results with complete leading-edge separation has shown reasonably good agreement. For cases with partial leading-edge separation, the lift is found to be highly nonlinear with angle of attack.

Asymmetric distribution of Echinoid defines the epidermal leading edge during Drosophila dorsal closure

PubMed Central

Laplante, Caroline

2011-01-01

During Drosophila melanogaster dorsal closure, lateral sheets of embryonic epidermis assemble an actomyosin cable at their leading edge and migrate dorsally over the amnioserosa, converging at the dorsal midline. We show that disappearance of the homophilic cell adhesion molecule Echinoid (Ed) from the amnioserosa just before dorsal closure eliminates homophilic interactions with the adjacent dorsal-most epidermal (DME) cells, which comprise the leading edge. The resulting planar polarized distribution of Ed in the DME cells is essential for the localized accumulation of actin regulators and for actomyosin cable formation at the leading edge and for the polarized localization of the scaffolding protein Bazooka/PAR-3. DME cells with uniform Ed fail to assemble a cable and protrude dorsally, suggesting that the cable restricts dorsal migration. The planar polarized distribution of Ed in the DME cells thus provides a spatial cue that polarizes the DME cell actin cytoskeleton, defining the epidermal leading edge and establishing its contractile properties. PMID:21263031

X-38 Experimental Aerothermodynamics

NASA Technical Reports Server (NTRS)

Horvath, Thomas J.; Berry, Scott A.; Merski, N. Ronald; Fitzgerald, Steve M.

2000-01-01

The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were

Aerodynamic robustness in owl-inspired leading-edge serrations: a computational wind-gust model.

PubMed

Rao, Chen; Liu, Hao

2018-06-08

Owls are a master to achieve silent flight in gliding and flapping flights under natural turbulent environments owing to their unique wing morphologies. While the leading-edge serrations are recently revealed, as a passive flow control micro-device, to play a crucial role in aerodynamic force production and sound suppression [25], the characteristics of wind-gust rejection associated with leading-edge serrations remain unclear. Here we address a large-eddy simulation (LES)-based study of aerodynamic robustness in owl-inspired leading-edge serrations, which is conducted with clean and serrated wing models through mimicking wind-gusts under a longitudinal fluctuation in free-stream inflow and a lateral fluctuation in pitch angle over a broad range of angles of attack (AoAs) over 0° ≤ Φ ≤ 20°. Our results show that the leading-edge serration-based passive flow control mechanisms associated with laminar-turbulent transition work effectively under fluctuated inflow and wing pitch, indicating that the leading-edge serrations are of potential gust fluctuation rejection or robustness in aerodynamic performance. Moreover, it is revealed that the tradeoff between turbulent flow control (i.e., aero-acoustic suppression) and force production in the serrated model holds independently to the wind-gust environments: poor at lower AoAs but capable of achieving equivalent aerodynamic performance at higher AoAs > 15o compared to the clean model. Our results reveal that the owl-inspired leading-edge serrations can be a robust micro-device for aero-acoustic control coping with unsteady and complex wind environments in biomimetic rotor designs for various fluid machineries. © 2018 IOP Publishing Ltd.

Experimental Stage Separation Tool Development in NASA Langley's Aerothermodynamics Laboratory

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Scallion, William I.

2005-01-01

As part of the research effort at NASA in support of the stage separation and ascent aerothermodynamics research program, proximity testing of a generic bimese wing-body configuration was conducted in NASA Langley's Aerothermodynamics Laboratory in the 20-Inch Mach 6 Air Tunnel. The objective of this work is the development of experimental tools and testing methodologies to apply to hypersonic stage separation problems for future multi-stage launch vehicle systems. Aerodynamic force and moment proximity data were generated at a nominal Mach number of 6 over a small range of angles of attack. The generic bimese configuration was tested in a belly-to-belly and back-to-belly orientation at 86 relative proximity locations. Over 800 aerodynamic proximity data points were taken to serve as a database for code validation. Longitudinal aerodynamic data generated in this test program show very good agreement with viscous computational predictions. Thus a framework has been established to study separation problems in the hypersonic regime using coordinated experimental and computational tools.

HEART Aerothermodynamic Analysis

NASA Technical Reports Server (NTRS)

Mazaheri, Alireza

2012-01-01

This paper presents an assessment of the aerothermodynamic environment around an 8.3 meter High Energy Atmospheric Reentry Test (HEART) vehicle. This study generated twelve nose shape configurations and compared their responses at the peak heating trajectory point against the baseline nose shape. The heat flux sensitivity to the angle of attack variations are also discussed. The possibility of a two-piece Thermal Protection System (TPS) design at the nose is also considered, as are the surface catalytic affects of the aeroheating environment of such configuration. Based on these analyses, an optimum nose shape is proposed to minimize the surface heating. A recommendation is also made for a two-piece TPS design, for which the surface catalytic uncertainty associated with the jump in heating at the nose-IAD juncture is reduced by a minimum of 93%. In this paper, the aeroshell is assumed to be rigid and the inflatable fluid interaction effect is left for future investigations.

Heat pipes for wing leading edges of hypersonic vehicles

NASA Technical Reports Server (NTRS)

Boman, B. L.; Citrin, K. M.; Garner, E. C.; Stone, J. E.

1990-01-01

Wing leading edge heat pipes were conceptually designed for three types of vehicle: an entry research vehicle, aero-space plane, and advanced shuttle. A full scale, internally instrumented sodium/Hastelloy X heat pipe was successfully designed and fabricated for the advanced shuttle application. The 69.4 inch long heat pipe reduces peak leading edge temperatures from 3500 F to 1800 F. It is internally instrumented with thermocouples and pressure transducers to measure sodium vapor qualities. Large thermal gradients and consequently large thermal stresses, which have the potential of limiting heat pipe life, were predicted to occur during startup. A test stand and test plan were developed for subsequent testing of this heat pipe. Heat pipe manufacturing technology was advanced during this program, including the development of an innovative technique for wick installation.

Application of local indentations for film cooling of gas turbine blade leading edge

NASA Astrophysics Data System (ADS)

Petelchyts, V. Yu.; Khalatov, A. A.; Pysmennyi, D. N.; Dashevskyy, Yu. Ya.

2016-09-01

The paper presents results of computer simulation of the film cooling on the turbine blade leading edge model where the air coolant is supplied through radial holes and row of cylindrical inclined holes placed inside hemispherical dimples or trench. The blowing factor was varied from 0.5 to 2.0. The model size and key initial parameters for simulation were taken as for a real blade of a high-pressure high-performance gas turbine. Simulation was performed using commercial software code ANSYS CFX. The simulation results were compared with reference variant (no dimples or trench) both for the leading edge area and for the flat plate downstream of the leading edge.

PRE_X Programme: Aerothermodynamic Objectives and Aeroshape Definition for in Flight Experiments

NASA Astrophysics Data System (ADS)

Lambert, O.; Tribot, J.-P.; Saint-Cloud, F.

2002-01-01

As the expendable launch vehicles (ELV) are limited in their trend to lower costs, the reusability (Reusable Launch Vehicle, RLV) could be the way to make drastic step. By the year 2001, CNES proposed through the ANGEL phase 1 programme to preprare the required technical maturity before that RLV's become alternatives to ELV's. In such way, system ,propulsion, ground based demonstrations, aero-thermo-dynamics as well as in flight experimentation are planned. This paper is focused on the aero-thermo-dynamics (ATD) and in flight demonstration activities with emphasis on the better understanding of ATD problems emerging from past programmes among them shock wave transitionnal boundary layer interaction on surface control, boundary layer transition, local aerothermodynamic effects, gas- surface interaction, catalycity, base flow prediction,...In order to minimize as small as possible the management risk a first generation of vehicle dubbed Pre_X is designed to validate technological choices and to have as soon as possible re-entry data to calibrate the various tools involved in the future RLV definition. In addition, the main requirement for PRE_X aeroshape definition and the two different design approaches considered by Dassault Aviation and EADS-LV are discussed. Then, the more promising concept for the PRE_X application is presented. Finally, the current status of the ATD activities is given as well as the perspectives.

Opportunities for research in aerothermodynamics

NASA Technical Reports Server (NTRS)

Graham, R. W.

1983-01-01

"Aerothermodynamics' involves the disciplines of chemistry, thermodynamics, fluid mechanics and heat transfer which have collaborative importance in propulsion systems. There are growing opportunities for the further application of these disciplines to improve the methodology for the design of advanced gas turbines; particularly, the combustor and turbine. Design procedures follow empirical or cut and try guidelines. The tremendous advances in computational analysis and in instrumentation techniques hold promise for research answers to complex physical processes that are currently not well understood. The transfer of basic research understanding to engineering design should result in shorter, less expensive development commitments for engines. The status and anticipated opportunities in research topics relevant to combustors and turbines is reviewed.

Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Covell, Peter F.; Wood, Richard M.; Miller, David S.

1987-01-01

An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

Wind-tunnel based definition of the AFE aerothermodynamic environment. [Aeroassist Flight Experiment

NASA Technical Reports Server (NTRS)

Miller, Charles G.; Wells, W. L.

1992-01-01

The Aeroassist Flight Experiment (AFE), scheduled to be performed in 1994, will serve as a precursor for aeroassisted space transfer vehicles (ASTV's) and is representative of entry concepts being considered for missions to Mars. Rationale for the AFE is reviewed briefly as are the various experiments carried aboard the vehicle. The approach used to determine hypersonic aerodynamic and aerothermodynamic characteristics over a wide range of simulation parameters in ground-based facilities is presented. Facilities, instrumentation and test procedures employed in the establishment of the data base are discussed. Measurements illustrating the effects of hypersonic simulation parameters, particularly normal-shock density ratio (an important parameter for hypersonic blunt bodies), and attitude on aerodynamic and aerothermodynamic characteristics are presented, and predictions from computational fluid dynamic (CFD) computer codes are compared with measurement.

On the leading edge vortex of thin wings

NASA Astrophysics Data System (ADS)

Arredondo, Abel; Viola, Ignazio Maria

2016-11-01

On thin wings, the sharp leading edge triggers laminar separation followed by reattachment, forming a Leading Edge Vortex (LEV). This flow feature is of paramount importance because, if periodically shed, it leads to large amplitude load fluctuations, while if stably attached to the wing, it can provide lift augmentation. We found that on asymmetric-spinnaker-type yacht sails, the LEV can be stable despite the relatively low sweep (30°). This finding, which was recently predicted numerically by Viola et al., has been confirmed through current flume tests on a 1:115th model scale sail. Forces were measured and Particle Image Velocimetry was performed on four horizontal sail sections at a Reynolds number of 1.7x104. Vortex detection revealed that the LEV becomes progressively larger and more stable towards the highest sections, where its axis has a smaller angle with respect to the freestream velocity. Mapping the sail section on a rotating cylinder through a Joukowski transformation, we quantified the lift augmentation provided by the LEV on each sail section. These results open up new sail design strategies based on the manipulation of the LEV and can be applicable to the wings of unmanned aerial vehicles and underwater vehicles. Project funded by Conacyt.

Ablative overlays for Space Shuttle leading edge ascent heat protection

NASA Technical Reports Server (NTRS)

Strauss, E. L.

1975-01-01

Ablative overlays were evaluated via a plasma-arc simulation of the ascent pulse on the leading edge of the Space Shuttle Orbiter. Overlay concepts included corkboard, polyisocyanurate foam, low-density Teflon, epoxy, and subliming salts. Their densities ranged from 4.9 to 81 lb per cu ft, and the thicknesses varied from 0.107 to 0.330 in. Swept-leading-edge models were fabricated from 30-lb per cu ft silicone-based ablators. The overlays were bonded to maintain the surface temperature of the base ablator below 500 F during ascent. Foams provided minimum-weight overlays, and subliming salts provided minimum-thickness overlays. Teflon left the most uniform surface after ascent heating.

Reynolds Number Effects on Leading Edge Radius Variations of a Supersonic Transport at Transonic Conditions

NASA Technical Reports Server (NTRS)

Rivers, S. M. B.; Wahls, R. A.; Owens, L. R.

2001-01-01

A computational study focused on leading-edge radius effects and associated Reynolds number sensitivity for a High Speed Civil Transport configuration at transonic conditions was conducted as part of NASA's High Speed Research Program. The primary purposes were to assess the capabilities of computational fluid dynamics to predict Reynolds number effects for a range of leading-edge radius distributions on a second-generation supersonic transport configuration, and to evaluate the potential performance benefits of each at the transonic cruise condition. Five leading-edge radius distributions are described, and the potential performance benefit including the Reynolds number sensitivity for each is presented. Computational results for two leading-edge radius distributions are compared with experimental results acquired in the National Transonic Facility over a broad Reynolds number range.

Effect of leading-edge load constraints on the design and performance of supersonic wings

NASA Technical Reports Server (NTRS)

Darden, C. M.

1985-01-01

A theoretical and experimental investigation was conducted to assess the effect of leading-edge load constraints on supersonic wing design and performance. In the effort to delay flow separation and the formation of leading-edge vortices, two constrained, linear-theory optimization approaches were used to limit the loadings on the leading edge of a variable-sweep planform design. Experimental force and moment tests were made on two constrained camber wings, a flat uncambered wing, and an optimum design with no constraints. Results indicate that vortex strength and separation regions were mildest on the severely and moderately constrained wings.

A feasibility study of heat-pipe-cooled leading edges for hypersonic cruise aircraft

NASA Technical Reports Server (NTRS)

Silverstein, C. C.

1971-01-01

A theoretical study of the use of heat pipe structures for cooling the leading edges of hypersonic cruise aircraft was carried out over a Mach number range of 6 to 12. Preliminary design studies showed that a heat pipe cooling structure with a 33-in. chordwise length could maintain the maximum temperature of a 65 deg sweepback wing with a 0.5-in. leading edge radius below 1600 F during cruise at Mach 8. A few relatively minor changes in the steady-state design of the structure were found necessary to insure satisfactory cooling during the climb to cruise speed and altitude. It was concluded that heat pipe cooling is an attractive, feasible technique for limiting leading edge temperatures of hypersonic cruise aircraft.

Effects of Leading Edge Defect on the Aerodynamic and Flow Characteristics of an S809 Airfoil.

PubMed

Wang, Yan; Zheng, Xiaojing; Hu, Ruifeng; Wang, Ping

Unexpected performance degradation occurs in wind turbine blades due to leading edge defect when suffering from continuous impacts with rain drops, hails, insects, or solid particles during its operation life. To assess this issue, this paper numerically investigates the steady and dynamic stall characteristics of an S809 airfoil with various leading edge defects. More leading edge defect sizes and much closer to practical parameters are investigated in the paper. Numerical computation is conducted using the SST k-ω turbulence model, and the method has been validated by comparison with existed published data. In order to ensure the calculation convergence, the residuals for the continuity equation are set to be less than 10-7 and 10-6 in steady state and dynamic stall cases. The simulations are conducted with the software ANSYS Fluent 13.0. It is found that the characteristics of aerodynamic coefficients and flow fields are sensitive to leading edge defect both in steady and dynamic conditions. For airfoils with the defect thickness of 6%tc, leading edge defect has a relative small influence on the aerodynamics of S809 airfoil. For other investigated defect thicknesses, leading edge defect has much greater influence on the flow field structures, pressure coefficients and aerodynamic characteristics of airfoil at relative small defect lengths. For example, the lift coefficients decrease and drag coefficients increase sharply after the appearance of leading edge defect. However, the aerodynamic characteristics could reach a constant value when the defect length is large enough. The flow field, pressure coefficient distribution and aerodynamic coefficients do not change a lot when the defect lengths reach to 0.5%c,1%c, 2%c and 3%c with defect thicknesses of 6%tc, 12%tc,18%tc and 25%tc, respectively. In addition, the results also show that the critical defect length/thickness ratio is 0.5, beyond which the aerodynamic characteristics nearly remain unchanged. In

Leading-Edge Flow Sensing for Aerodynamic Parameter Estimation

NASA Astrophysics Data System (ADS)

Saini, Aditya

The identification of inflow air data quantities such as airspeed, angle of attack, and local lift coefficient on various sections of a wing or rotor blade provides the capability for load monitoring, aerodynamic diagnostics, and control on devices ranging from air vehicles to wind turbines. Real-time measurement of aerodynamic parameters during flight provides the ability to enhance aircraft operating capabilities while preventing dangerous stall situations. This thesis presents a novel Leading-Edge Flow Sensing (LEFS) algorithm for the determination of the air -data parameters using discrete surface pressures measured at a few ports in the vicinity of the leading edge of a wing or blade section. The approach approximates the leading-edge region of the airfoil as a parabola and uses pressure distribution from the exact potential-ow solution for the parabola to _t the pressures measured from the ports. Pressures sensed at five discrete locations near the leading edge of an airfoil are given as input to the algorithm to solve the model using a simple nonlinear regression. The algorithm directly computes the inflow velocity, the stagnation-point location, section angle of attack and lift coefficient. The performance of the algorithm is assessed using computational and experimental data in the literature for airfoils under different ow conditions. The results show good correlation between the actual and predicted aerodynamic quantities within the pre-stall regime, even for a rotating blade section. Sensing the deviation of the aerodynamic behavior from the linear regime requires additional information on the location of ow separation on the airfoil surface. Bio-inspired artificial hair sensors were explored as a part of the current research for stall detection. The response of such artificial micro-structures can identify critical ow characteristics, which relate directly to the stall behavior. The response of the microfences was recorded via an optical microscope for

Real-Gas Flow Properties for NASA Langley Research Center Aerothermodynamic Facilities Complex Wind Tunnels

NASA Technical Reports Server (NTRS)

Hollis, Brian R.

1996-01-01

A computational algorithm has been developed which can be employed to determine the flow properties of an arbitrary real (virial) gas in a wind tunnel. A multiple-coefficient virial gas equation of state and the assumption of isentropic flow are used to model the gas and to compute flow properties throughout the wind tunnel. This algorithm has been used to calculate flow properties for the wind tunnels of the Aerothermodynamics Facilities Complex at the NASA Langley Research Center, in which air, CF4. He, and N2 are employed as test gases. The algorithm is detailed in this paper and sample results are presented for each of the Aerothermodynamic Facilities Complex wind tunnels.

A Reynolds Number Study of Wing Leading-Edge Effects on a Supersonic Transport Model at Mach 0.3

NASA Technical Reports Server (NTRS)

Williams, M. Susan; Owens, Lewis R., Jr.; Chu, Julio

1999-01-01

A representative supersonic transport design was tested in the National Transonic Facility (NTF) in its original configuration with small-radius leading-edge flaps and also with modified large-radius inboard leading-edge flaps. Aerodynamic data were obtained over a range of Reynolds numbers at a Mach number of 0.3 and angles of attack up to 16 deg. Increasing the radius of the inboard leading-edge flap delayed nose-up pitching moment to a higher lift coefficient. Deflecting the large-radius leading-edge flap produced an overall decrease in lift coefficient and delayed nose-up pitching moment to even higher angles of attack as compared with the undeflected large- radius leading-edge flap. At angles of attack corresponding to the maximum untrimmed lift-to-drag ratio, lift and drag coefficients decreased while lift-to-drag ratio increased with increasing Reynolds number. At an angle of attack of 13.5 deg., the pitching-moment coefficient was nearly constant with increasing Reynolds number for both the small-radius leading-edge flap and the deflected large-radius leading-edge flap. However, the pitching moment coefficient increased with increasing Reynolds number for the undeflected large-radius leading-edge flap above a chord Reynolds number of about 35 x 10 (exp 6).

The Flow Field on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2009-11-01

The exceptional mobility of the humpback whale has been linked to the use of its unique pectoral flippers. Biologists speculate that the flippers leading edge protuberances are a form of passive flow control. Force measurements on 2D hydrofoils with spanwise uniform leading edge protuberances, resembling those seen on the humpback whale flipper, were taken in a water tunnel and have revealed performance modifications when compared to a baseline NACA 63(4)-021 hydrofoil model. Qualitative flow visualization techniques and Particle Image Velocimetry (PIV) flow field measurements on the modified hydrofoils have shown that streamwise vortices originating from the shoulders of the protuberances are the likely cause of performance changes. Varying levels of interaction among adjacent streamwise vortices have been observed as a function of angle of attack and chord location. The circulation of these vortices as a function of angle of attack and spatial location was measured and an analysis of the vortex interactions will be presented.

Design and fabrication of a high temperature leading edge heating array, phase 1

NASA Technical Reports Server (NTRS)

1972-01-01

Progress during a Phase 1 program to design a high temperature heating array is reported for environmentally testing full-scale shuttle leading edges (30 inch span, 6 to 15 inch radius) at flight heating rates and pressures. Heat transfer analyses of the heating array, individual modules, and the shuttle leading edge were performed, which influenced the array design, and the design, fabrication, and testing of a prototype heater module.

Potential environmental effects of the leading edge hydrokinetic energy technology.

DOT National Transportation Integrated Search

2017-05-01

The Volpe Center evaluated potential environmental challenges and benefits of the ARPA-E funded research project, Marine Hydrokinetic Energy Harvesting Using Cyber-Physical Systems, led by Brown University. The Leading Edge research team develo...

Probabilistic Structural Health Monitoring of the Orbiter Wing Leading Edge

NASA Technical Reports Server (NTRS)

Yap, Keng C.; Macias, Jesus; Kaouk, Mohamed; Gafka, Tammy L.; Kerr, Justin H.

2011-01-01

A structural health monitoring (SHM) system can contribute to the risk management of a structure operating under hazardous conditions. An example is the Wing Leading Edge Impact Detection System (WLEIDS) that monitors the debris hazards to the Space Shuttle Orbiter s Reinforced Carbon-Carbon (RCC) panels. Since Return-to-Flight (RTF) after the Columbia accident, WLEIDS was developed and subsequently deployed on board the Orbiter to detect ascent and on-orbit debris impacts, so as to support the assessment of wing leading edge structural integrity prior to Orbiter re-entry. As SHM is inherently an inverse problem, the analyses involved, including those performed for WLEIDS, tend to be associated with significant uncertainty. The use of probabilistic approaches to handle the uncertainty has resulted in the successful implementation of many development and application milestones.

Modifications of W and Mo leading edges under plasma loads in DIII-D divertor

NASA Astrophysics Data System (ADS)

Rudakov, D. L.; Bykov, I.; Moyer, R. A.; Abrams, T.; Chrobak, C. P.; Guo, H. Y.; Stahl, B.; Thomas, D. M.; Barton, J. L.; Nygren, R. E.; Watkins, J. G.; Lasnier, C. J.; Litnovsky, Andrey; Stangeby, P. C.; Unterberg, E. A.

2017-10-01

Cracking and melting of W and Mo leading edges were observed in the lower divertor of DIII-D during experiments with intentionally misaligned W monoblocks (MBs) and in the course of the Metal Rings Campaign involving W-coated Mo tile inserts (TIs). MBs were exposed near the attached outer strike point during deuterium and helium L- and H-mode discharges using DiMES. Two of the MBs were misaligned by 0.3 mm and 1 mm, forming leading edges. Particulate ejection from a 1 mm leading edge was observed during the exposure, and evidence of melting and cracking was found post mortem. Two toroidal rings of TIs were installed in the lower outer divertor, the inner one at the floor and the outer one at the shelf. The floor TIs bowed during plasma exposure forming leading edges up to 1.2 mm high; about 40% of these edges experienced melting. Re-solidified melt layers up to 1 mm thick were observed, their shape being consistent with motion in the jx B direction with j driven by electron emission. Work supported by US DOE under DE-FC02-04ER54698, DE-FG02-07ER54917, DE-AC04-94AL85000, DE-AC52-07NA27344, and DE-AC05-00OR22725.

Chronophin coordinates cell leading edge dynamics by controlling active cofilin levels

PubMed Central

Delorme-Walker, Violaine; Seo, Ji-Yeon; Gohla, Antje; Fowler, Bruce; Bohl, Ben; DerMardirossian, Céline

2015-01-01

Cofilin, a critical player of actin dynamics, is spatially and temporally regulated to control the direction and force of membrane extension required for cell locomotion. In carcinoma cells, although the signaling pathways regulating cofilin activity to control cell direction have been established, the molecular machinery required to generate the force of the protrusion remains unclear. We show that the cofilin phosphatase chronophin (CIN) spatiotemporally regulates cofilin activity at the cell edge to generate persistent membrane extension. We show that CIN translocates to the leading edge in a PI3-kinase–, Rac1-, and cofilin-dependent manner after EGF stimulation to activate cofilin, promotes actin free barbed end formation, accelerates actin turnover, and enhances membrane protrusion. In addition, we establish that CIN is crucial for the balance of protrusion/retraction events during cell migration. Thus, CIN coordinates the leading edge dynamics by controlling active cofilin levels to promote MTLn3 cell protrusion. PMID:26324884

Effects of Leading Edge Defect on the Aerodynamic and Flow Characteristics of an S809 Airfoil

PubMed Central

Wang, Yan; Zheng, Xiaojing; Hu, Ruifeng; Wang, Ping

2016-01-01

Background Unexpected performance degradation occurs in wind turbine blades due to leading edge defect when suffering from continuous impacts with rain drops, hails, insects, or solid particles during its operation life. To assess this issue, this paper numerically investigates the steady and dynamic stall characteristics of an S809 airfoil with various leading edge defects. More leading edge defect sizes and much closer to practical parameters are investigated in the paper. Methodology Numerical computation is conducted using the SST k-ω turbulence model, and the method has been validated by comparison with existed published data. In order to ensure the calculation convergence, the residuals for the continuity equation are set to be less than 10−7 and 10−6 in steady state and dynamic stall cases. The simulations are conducted with the software ANSYS Fluent 13.0. Results It is found that the characteristics of aerodynamic coefficients and flow fields are sensitive to leading edge defect both in steady and dynamic conditions. For airfoils with the defect thickness of 6%tc, leading edge defect has a relative small influence on the aerodynamics of S809 airfoil. For other investigated defect thicknesses, leading edge defect has much greater influence on the flow field structures, pressure coefficients and aerodynamic characteristics of airfoil at relative small defect lengths. For example, the lift coefficients decrease and drag coefficients increase sharply after the appearance of leading edge defect. However, the aerodynamic characteristics could reach a constant value when the defect length is large enough. The flow field, pressure coefficient distribution and aerodynamic coefficients do not change a lot when the defect lengths reach to 0.5%c,1%c, 2%c and 3%c with defect thicknesses of 6%tc, 12%tc,18%tc and 25%tc, respectively. In addition, the results also show that the critical defect length/thickness ratio is 0.5, beyond which the aerodynamic characteristics

Design & fabrication of two seated aircraft with an advanced rotating leading edge wing

NASA Astrophysics Data System (ADS)

Al Ahmari, Saeed Abdullah Saeed

The title of this thesis is "Design & Fabrication of two Seated Aircraft with an Advanced Rotating Leading Edge Wing", this gives almost a good description of the work has been done. In this research, the moving surface boundary-layer control (MSBC) concept was investigated and implemented. An experimental model was constructed and tested in wind tunnel to determine the aerodynamic characteristics using the leading edge moving surface of modified semi-symmetric airfoil NACA1214. The moving surface is provided by a high speed rotating cylinder, which replaces the leading edge of the airfoil. The angle of attack, the cylinder surfaces velocity ratio Uc/U, and the flap deflection angle effects on the lift and drag coefficients and the stall angle of attack were investigated. This new technology was applied to a 2-seat light-sport aircraft that is designed and built in the Aerospace Engineering Department at KFUPM. The project team is led by the aerospace department chairman Dr. Ahmed Z. AL-Garni and Dr. Wael G. Abdelrahman and includes graduate and under graduate student. The wing was modified to include a rotating cylinder along the leading edge of the flap portion. This produced very promising results such as the increase of the maximum lift coefficient at Uc/U=3 by 82% when flaps up and 111% when flaps down at 40° and stall was delayed by 8degrees in both cases. The laboratory results also showed that the effective range of the leading-edge rotating cylinder is at low angles of attack which reduce the need for higher angles of attack for STOL aircraft.

Closed Form Equations for the Preliminary Design of a Heat-Pipe-Cooled Leading Edge

NASA Technical Reports Server (NTRS)

Glass, David E.

1998-01-01

A set of closed form equations for the preliminary evaluation and design of a heat-pipe-cooled leading edge is presented. The set of equations can provide a leading-edge designer with a quick evaluation of the feasibility of using heat-pipe cooling. The heat pipes can be embedded in a metallic or composite structure. The maximum heat flux, total integrated heat load, and thermal properties of the structure and heat-pipe container are required input. The heat-pipe operating temperature, maximum surface temperature, heat-pipe length, and heat pipe-spacing can be estimated. Results using the design equations compared well with those from a 3-D finite element analysis for both a large and small radius leading edge.

Surface thermochemical effects on TPS-coupled aerothermodynamics in hypersonic Martian gas flow

NASA Astrophysics Data System (ADS)

Yang, Xiaofeng; Gui, Yewei; Tang, Wei; Du, Yanxia; Liu, Lei; Xiao, Guangming; Wei, Dong

2018-06-01

This paper deals with the surface thermochemical effects on TPS-coupled aerothermodynamics in hypersonic Martian gas flow. An interface condition with finite-rate thermochemistry was established to balance the three-dimensional Navier-Stokes solver and TPS thermal response solver, and a series of coupled simulations of chemical non-equilibrium aerothermodynamics and structure heat transfer with various surface catalycities were performed for hypersonic Mars entries. The analysis of surface thermochemistry reveals that the surface chemical reactions have great contribution to aerodynamic heating, and the temperature-dependence of finite-rate catalysis highly influences the evolution of the coupling aerodynamic heating in the coupling process. For fixed free stream parameters with proper catalytic excitation energy, a "leap" phenomenon of the TPS-coupled heat flux with the coupling time appears in the initial stage of the coupling process, due to the strong thermochemical effects on the TPS surface.

Effect of canard position and wing leading-edge flap deflection on wing buffet at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Henderson, W. P.; Huffman, J. K.

1974-01-01

A generalized wind-tunnel model, with canard and wing planform typical of highly maneuverable aircraft, was tested. The addition of a canard above the wing chord plane, for the configuration with leading-edge flaps undeflected, produced substantially higher total configuration lift coefficients before buffet onset than the configuration with the canard off and leading-edge flaps undeflected. The wing buffet intensity was substantially lower for the canard-wing configuration than the wing-alone configuration. The low-canard configuration generally displayed the poorest buffet characteristics. Deflecting the wing leading-edge flaps substantially improved the wing buffet characteristics for canard-off configurations. The addition of the high canard did not appear to substantially improve the wing buffet characteristics of the wing with leading-edge flaps deflected.

Periodic and aperiodic flow patterns around an airfoil with leading-edge protuberances

NASA Astrophysics Data System (ADS)

Cai, Chang; Zuo, Zhigang; Maeda, Takao; Kamada, Yasunari; Li, Qing'an; Shimamoto, Kensei; Liu, Shuhong

2017-11-01

Recently leading-edge protuberances have attracted great attention as a passive method for separation control. In this paper, the effect of multiple leading-edge protuberances on the performance of a two-dimensional airfoil is investigated through experimental measurement of aerodynamic forces, surface tuft visualization, and numerical simulation. In contrast to the sharp stall of the baseline airfoil with large hysteresis effect during AOA (angle of attack) increasing and decreasing, the stall process of the modified airfoil with leading-edge protuberances is gentle and stable. Flow visualization revealed that the flow past each protuberance is periodic and symmetric at small AOAs. Streamwise vortices are generated on the shoulders of the protuberance, leading to a larger separation around the valley sections and a longer attachment along the peak sections. When some critical AOA is exceeded, aperiodic and asymmetric flow patterns occur on the protuberances at different spanwise positions, with leading-edge separation on some of the valley sections and non-stalled condition elsewhere. A combined mechanism, involving both the compartmentalization effect of the slender momentum-enhanced attached flows on the protuberance peaks and the downwash effect of the local stalled region with low circulation, is proposed to explain the generation of the aperiodic flow patterns. The influence of the number of protuberances is also investigated, which shows similar aperiodic flow patterns. The distance between the neighboring local stalled valley sections is found to be in the range of 4-7 times the protuberance wavelength. According to the proposed mechanism, it is speculated that the distance between the neighboring local stalled valley sections is inclined to increase with a smaller protuberance amplitude or at a larger AOA.

SiC/SiC Leading Edge Turbine Airfoil Tested Under Simulated Gas Turbine Conditions

NASA Technical Reports Server (NTRS)

Robinson, R. Craig; Hatton, Kenneth S.

1999-01-01

Silicon-based ceramics have been proposed as component materials for use in gas turbine engine hot-sections. A high pressure burner rig was used to expose both a baseline metal airfoil and ceramic matrix composite leading edge airfoil to typical gas turbine conditions to comparatively evaluate the material response at high temperatures. To eliminate many of the concerns related to an entirely ceramic, rotating airfoil, this study has focused on equipping a stationary metal airfoil with a ceramic leading edge insert to demonstrate the feasibility and benefits of such a configuration. Here, the idea was to allow the SiC/SiC composite to be integrated as the airfoil's leading edge, operating in a "free-floating" or unrestrained manner. and provide temperature relief to the metal blade underneath. The test included cycling the airfoils between simulated idle, lift, and cruise flight conditions. In addition, the airfoils were air-cooled, uniquely instrumented, and exposed to the same internal and external conditions, which included gas temperatures in excess of 1370 C (2500 F). Results show the leading edge insert remained structurally intact after 200 simulated flight cycles with only a slightly oxidized surface. The instrumentation clearly suggested a significant reduction (approximately 600 F) in internal metal temperatures as a result of the ceramic leading edge. The object of this testing was to validate the design and analysis done by Materials Research and Design of Rosemont, PA and to determine the feasibility of this design for the intended application.

Effects of Nose Radius and Aerodynamic Loading on Leading Edge Receptivity

NASA Technical Reports Server (NTRS)

Hammerton, P. W.; Kerschen, E. J.

1998-01-01

An analysis is presented of the effects of airfoil thickness and mean aerodynamic loading on boundary-layer receptivity in the leading-edge region. The case of acoustic free-stream disturbances, incident on a thin cambered airfoil with a parabolic leading edge in a low Mach number flow, is considered. An asymptotic analysis based on large Reynolds number is developed, supplemented by numerical results. The airfoil thickness distribution enters the theory through a Strouhal number based on the nose radius of the airfoil, S = (omega)tau(sub n)/U, where omega is the frequency of the acoustic wave and U is the mean flow speed. The influence of mean aerodynamic loading enters through an effective angle-of-attack parameter ti, related to flow around the leading edge from the lower surface to the upper. The variation of the receptivity level is analyzed as a function of S, mu, and characteristics of the free-stream acoustic wave. For an unloaded leading edge, a finite nose radius dramatically reduces the receptivity level compared to that for a flat plate, the amplitude of the instability waves in the boundary layer being decreased by an order of magnitude when S = 0.3. Modest levels of aerodynamic loading are found to further decrease the receptivity level for the upper surface of the airfoil, while an increase in receptivity level occurs for the lower surface. For larger angles of attack close to the critical angle for boundary layer separation, a local rise in the receptivity level occurs for the upper surface, while for the lower surface the receptivity decreases. The effects of aerodynamic loading are more pronounced at larger values of S. Oblique acoustic waves produce much higher receptivity levels than acoustic waves propagating downstream parallel to the airfoil chord.

Shuttle Wing Leading Edge Root Cause NDE Team Findings and Implementation of Quantitative Flash Infrared Thermography

NASA Technical Reports Server (NTRS)

Burke, Eric R.

2009-01-01

Comparison metrics can be established to reliably and repeatedly establish the health of the joggle region of the Orbiter Wing Leading Edge reinforced carbon carbon (RCC) panels. Using these metrics can greatly reduced the man hours needed to perform, wing leading edge scanning for service induced damage. These time savings have allowed for more thorough inspections to be preformed in the necessary areas with out affecting orbiter flow schedule. Using specialized local inspections allows for a larger margin of safety by allowing for more complete characterizations of panel defects. The presence of the t-seal during thermographic inspection can have adverse masking affects on ability properly characterize defects that exist in the joggle region of the RCC panels. This masking affect dictates the final specialized inspection should be preformed with the t-seal removed. Removal of the t-seal and use of the higher magnification optics has lead to the most effective and repeatable inspection method for characterizing and tracking defects in the wing leading edge. Through this study some inadequacies in the main health monitoring system for the orbiter wing leading edge have been identified and corrected. The use of metrics and local specialized inspection have lead to a greatly increased reliability and repeatable inspection of the shuttle wing leading edge.

Overview of aerothermodynamic loads definition study

NASA Technical Reports Server (NTRS)

Gaugler, Raymond E.

1991-01-01

The objective of the Aerothermodynamic Loads Definition Study is to develop methods of accurately predicting the operating environment in advanced Earth-to-Orbit (ETO) propulsion systems, such as the Space Shuttle Main Engine (SSME) powerhead. Development of time averaged and time dependent three dimensional viscous computer codes as well as experimental verification and engine diagnostic testing are considered to be essential in achieving that objective. Time-averaged, nonsteady, and transient operating loads must all be well defined in order to accurately predict powerhead life. Described here is work in unsteady heat flow analysis, improved modeling of preburner flow, turbulence modeling for turbomachinery, computation of three dimensional flow with heat transfer, and unsteady viscous multi-blade row turbine analysis.

Simulation of Flow Through Breach in Leading Edge at Mach 24

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Alter, Stephen J.

2004-01-01

A baseline solution for CFD Point 1 (Mach 24) in the STS-107 accident investigation was modified to include effects of holes through the leading edge into a vented cavity. The simulations were generated relatively quickly and early in the investigation by making simplifications to the leading edge cavity geometry. These simplifications in the breach simulations enabled: 1) A very quick grid generation procedure; 2) High fidelity corroboration of jet physics with internal surface impingements ensuing from a breach through the leading edge, fully coupled to the external shock layer flow at flight conditions. These simulations provided early evidence that the flow through a 2 inch diameter (or larger) breach enters the cavity with significant retention of external flow directionality. A normal jet directed into the cavity was not an appropriate model for these conditions at CFD Point 1 (Mach 24). The breach diameters were of the same order or larger than the local, external boundary-layer thickness. High impingement heating and pressures on the downstream lip of the breach were computed. It is likely that hole shape would evolve as a slot cut in the direction of the external streamlines. In the case of the 6 inch diameter breach the boundary layer is fully ingested.

Advanced leading edge thermal-structure concept. Direct bond reusable surface insulation to a composite structure

NASA Technical Reports Server (NTRS)

Riccitiello, S. R.; Figueroa, H.; Coe, C. F.; Kuo, C. P.

1984-01-01

An advanced leading-edge concept was analyzed using the space shuttle leading edge system as a reference model. The comparison indicates that a direct-bond system utilizing a high temperature (2700 F) fibrous refractory composite insulation tile bonded to a high temperature (PI/graphite) composite structure can result in a weight savings of up to 800 lb. The concern that tile damage or loss during ascent would result in adverse entry aerodynamics if a leading edge tile system were used is addressed. It was found from experiment that missing tiles (as many as 22) on the leading edge would not significantly affect the basic force-and-moment aerodynamic coefficients. Additionally, this concept affords a degree of redundancy to a thermal protection system in that the base structure (being a composite material) ablates and neither melts nor burns through when subjected to entry heating in the event tiles are actually lost or damaged during ascent.

Manipulation of upstream rotor leading edge vortex and its effects on counter rotating propeller noise

NASA Technical Reports Server (NTRS)

Squires, Becky

1993-01-01

The leading edge vortex of a counter rotating propeller (CRP) model was altered by using shrouds and by turning the upstream rotors to a forward sweep configuration. Performance, flow, and acoustic data were used to determine the effect of vortex impingement on the noise signature of the CRP system. Forward sweep was found to eliminate the leading edge vortex of the upstream blades. Removal of the vortex had little effect on the tone noise at the forward and rear blade passing frequencies (BPF's) but significantly altered both the sound pressure level and directivity of the interaction tone which occurs at the sum of the two BPF's. A separate manipulation of the leading edge vortex performed by installing shrouds of various inlet length on the CRP verified that diverting the vortex path increases the noise level of the interaction tone. An unexpected link has been established between the interaction tone and the leading edge vortex-blade interaction phenomenon.

Owl-inspired leading-edge serrations play a crucial role in aerodynamic force production and sound suppression.

PubMed

Rao, Chen; Ikeda, Teruaki; Nakata, Toshiyuki; Liu, Hao

2017-07-04

Owls are widely known for silent flight, achieving remarkably low noise gliding and flapping flights owing to their unique wing morphologies, which are normally characterized by leading-edge serrations, trailing-edge fringes and velvet-like surfaces. How these morphological features affect aerodynamic force production and sound suppression or noise reduction, however, is still not well known. Here we address an integrated study of owl-inspired single feather wing models with and without leading-edge serrations by combining large-eddy simulations (LES) with particle-image velocimetry (PIV) and force measurements in a low-speed wind tunnel. With velocity and pressure spectra analysis, we demonstrate that leading-edge serrations can passively control the laminar-turbulent transition over the upper wing surface, i.e. the suction surface at all angles of attack (0°  <  AoA  <  20°), and hence play a crucial role in aerodynamic force and sound production. We find that there exists a tradeoff between force production and sound suppression: serrated leading-edges reduce aerodynamic performance at lower AoAs  <  15° compared to clean leading-edges but are capable of achieving both noise reduction and aerodynamic performance at higher AoAs  >  15° where owl wings often reach in flight. Our results indicate that the owl-inspired leading-edge serrations may be a useful device for aero-acoustic control in biomimetic rotor designs for wind turbines, aircrafts, multi-rotor drones as well as other fluid machinery.

Aerothermodynamic Testing of Aerocapture and Planetary Probe Geometries in Hypersonic Ballistic-Range Environments

NASA Technical Reports Server (NTRS)

Wilder, M. C.; Reda, D. C.; Bogdanoff, D. W.; Olejniczak, J.

2005-01-01

A viewgraph presentation on aerothermodynamic testing of aerocapture and planetary probe design methods in hypersonic ballistic range environments is shown. The topics include: 1) Ballistic Range Testing; 2) NASA-Ames Hypervelocity Free Flight Facility; and 3) Representative Results.

Aerothermodynamics of Blunt Body Entry Vehicles. Chapter 3

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Borrelli, Salvatore

2011-01-01

In this chapter, the aerothermodynamic phenomena of blunt body entry vehicles are discussed. Four topics will be considered that present challenges to current computational modeling techniques for blunt body environments: turbulent flow, non-equilibrium flow, rarefied flow, and radiation transport. Examples of comparisons between computational tools to ground and flight-test data will be presented in order to illustrate the challenges existing in the numerical modeling of each of these phenomena and to provide test cases for evaluation of Computational Fluid Dynamics (CFD) code predictions.

Reynolds Number, Compressibility, and Leading-Edge Bluntness Effects on Delta-Wing Aerodynamics

NASA Technical Reports Server (NTRS)

Luckring, James M.

2004-01-01

An overview of Reynolds number, compressibility, and leading edge bluntness effects is presented for a 65 degree delta wing. The results of this study address both attached and vortex-flow aerodynamics and are based upon a unique data set obtained in the NASA-Langley National Transonic Facility (NTF) for i) Reynolds numbers ranging from conventional wind-tunnel to flight values, ii) Mach numbers ranging from subsonic to transonic speeds, and iii) leading-edge bluntness values that span practical slender wing applications. The data were obtained so as to isolate the subject effects and they present many challenges for Computational Fluid Dynamics (CFD) studies.

Development flight tests of JetStar LFC leading-edge flight test experiment

NASA Technical Reports Server (NTRS)

Fisher, David F.; Fischer, Michael C.

1987-01-01

The overall objective of the flight tests on the JetStar aircraft was to demonstrate the effectiveness and reliability of laminar flow control under representative flight conditions. One specific objective was to obtain laminar flow on the JetStar leading-edge test articles for the design and off-design conditions. Another specific objective was to obtain operational experience on a Laminar Flow Control (LFC) leading-edge system in a simulated airline service. This included operational experience with cleaning requirements, the effect of clogging, possible foreign object damage, erosion, and the effects of ice particle and cloud encounters. Results are summarized.

Subsonic balance and pressure investigation of a 60 deg delta wing with leading edge devices

NASA Technical Reports Server (NTRS)

Tingas, S. A.; Rao, D. M.

1982-01-01

Low supersonic wave drag makes the thin highly swept delta wing the logical choice for use on aircraft designed for supersonic cruise. However, the high-lift maneuver capability of the aircraft is limited by severe induced-drag penalties attributed to loss of potential flow leading-edge suction. This drag increase may be alleviated through leading-edge flow control to recover lost aerodynamic thrust through either retention of attached leading-edge flow to higher angles of attack or exploitation of the increased suction potential of separation-induced vortex flow. A low-speed wind-tunnel investigation was undertaken to examine the high-lift devices such as fences, chordwise slots, pylon vortex generators, leading-edge vortex flaps, and sharp leading-edge extensions. The devices were tested individually and in combinations in an attempt to improve high-alpha drag performance with a minimum of low-alpha drag penalty. This report presents an analysis of the force, moment, and static pressure data obtained in angles of attack up to 23 deg, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter, respectively. The results indicate that all the devices produced drag and longitudinal/lateral stability improvements at high lift with, in most cases, minor drag penalties at low angles of attack.

Overview of X-38 Hypersonic Aerothermodynamic Wind Tunnel Data and Comparison with Numerical Results

NASA Technical Reports Server (NTRS)

Campbell, C.; Caram, J.; Berry, S.; Horvath, T.; Merski, N.; Loomis, M.; Venkatapathy, E.

2004-01-01

A NASA team of engineers has been organized to design a crew return vehicle for returning International Space Station crew members from orbit. The hypersonic aerothermodynamic characteristics of the X-23/X-24A derived X-38 crew return vehicle are being evaluated in various wind tunnels in support of this effort. Aerothermodynamic data from two NASA hypersonic tunnels at Mach 6 and Mach 10 has been obtained with cast ceramic models and a thermographic phosphorus digital imaging system. General windward surface heating features are described based on experimental surface heating images and surface oil flow patterns for the nominal hypersonic aerodynamic orientation. Body flap reattachment heating levels are examined. Computational Fluid Dynamics tools have been applied at the appropriate wind tunnel conditions to make comparisons with this data.

Detail view of the leading and top edge of the ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of the leading and top edge of the vertical stabilizer of the Orbiter Discovery showing the thermal protection system components with the white Advanced Flexible Reusable Surface Insulation (AFRSI) blanket and the black High-temperature Reusable Surface Insulation (HRSI) tiles along the outer edges. The marks seen on the HRSI tiles are injection point marks and holes for the application of waterproofing material. This view was taken from a service platform in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

NASA Technical Reports Server (NTRS)

Ghaffari, F.; Chaturvedi, S. K.

1984-01-01

An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

Simulated big sagebrush regeneration supports predicted changes at the trailing and leading edges of distribution shifts

USGS Publications Warehouse

Schlaepfer, Daniel R.; Taylor, Kyle A.; Pennington, Victoria E.; Nelson, Kellen N.; Martin, Trace E.; Rottler, Caitlin M.; Lauenroth, William K.; Bradford, John B.

2015-01-01

Many semi-arid plant communities in western North America are dominated by big sagebrush. These ecosystems are being reduced in extent and quality due to economic development, invasive species, and climate change. These pervasive modifications have generated concern about the long-term viability of sagebrush habitat and sagebrush-obligate wildlife species (notably greater sage-grouse), highlighting the need for better understanding of the future big sagebrush distribution, particularly at the species' range margins. These leading and trailing edges of potential climate-driven sagebrush distribution shifts are likely to be areas most sensitive to climate change. We used a process-based regeneration model for big sagebrush, which simulates potential germination and seedling survival in response to climatic and edaphic conditions and tested expectations about current and future regeneration responses at trailing and leading edges that were previously identified using traditional species distribution models. Our results confirmed expectations of increased probability of regeneration at the leading edge and decreased probability of regeneration at the trailing edge below current levels. Our simulations indicated that soil water dynamics at the leading edge became more similar to the typical seasonal ecohydrological conditions observed within the current range of big sagebrush ecosystems. At the trailing edge, an increased winter and spring dryness represented a departure from conditions typically supportive of big sagebrush. Our results highlighted that minimum and maximum daily temperatures as well as soil water recharge and summer dry periods are important constraints for big sagebrush regeneration. Overall, our results confirmed previous predictions, i.e., we see consistent changes in areas identified as trailing and leading edges; however, we also identified potential local refugia within the trailing edge, mostly at sites at higher elevation. Decreasing

Flexible Plug Repair for Shuttle Wing Leading Edge

NASA Technical Reports Server (NTRS)

Camarda, Charles J.; Sikora, Joseph; Smith, Russel; Rivers, H.; Scotti, Stephen J.; Fuller, Alan M.; Klacka, Robert; Reinders, Martin; Schwind, Francis; Sullivan, Brian;

Increased heat transfer to elliptical leading edges due to spanwise variations in the freestream momentum: Numerical and experimental results

NASA Technical Reports Server (NTRS)

Rigby, D. L.; Vanfossen, G. J.

1992-01-01

A study of the effect of spanwise variation in momentum on leading edge heat transfer is discussed. Numerical and experimental results are presented for both a circular leading edge and a 3:1 elliptical leading edge. Reynolds numbers in the range of 10,000 to 240,000 based on leading edge diameter are investigated. The surface of the body is held at a constant uniform temperature. Numerical and experimental results with and without spanwise variations are presented. Direct comparison of the two-dimensional results, that is, with no spanwise variations, to the analytical results of Frossling is very good. The numerical calculation, which uses the PARC3D code, solves the three-dimensional Navier-Stokes equations, assuming steady laminar flow on the leading edge region. Experimentally, increases in the spanwise-averaged heat transfer coefficient as high as 50 percent above the two-dimensional value were observed. Numerically, the heat transfer coefficient was seen to increase by as much as 25 percent. In general, under the same flow conditions, the circular leading edge produced a higher heat transfer rate than the elliptical leading edge. As a percentage of the respective two-dimensional values, the circular and elliptical leading edges showed similar sensitivity to span wise variations in momentum. By equating the root mean square of the amplitude of the spanwise variation in momentum to the turbulence intensity, a qualitative comparison between the present work and turbulent results was possible. It is shown that increases in leading edge heat transfer due to spanwise variations in freestream momentum are comparable to those due to freestream turbulence.

Reynolds Number and Leading-Edge Bluntness Effects on a 65 Deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2002-01-01

A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2002-01-01

A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

Aerothermodynamics research at NASA Ames Research Center

NASA Technical Reports Server (NTRS)

Deiwert, George S.

1987-01-01

Research activity in the aerothermodynamics branch at the NASA Ames Research Center is reviewed. Advanced concepts and mission studies relating to the next generation aerospace transportation systems are summarized and directions for continued research identified. Theoretical and computational studies directed at determining flow fields and radiative and convective heating loads in real gases are described. Included are Navier-Stokes codes for equilibrium and thermochemical nonequilibrium air. Experimental studies in the 3.5-ft hypersonic wind tunnel, the ballistic ranges, and the electric arc driven shock tube are described. Tested configurations include generic hypersonic aerospace plane configurations, aeroassisted orbital transfer vehicle shapes and Galileo probe models.

Analysis of airfoil leading edge separation bubbles

NASA Technical Reports Server (NTRS)

Carter, J. E.; Vatsa, V. N.

1982-01-01

A local inviscid-viscous interaction technique was developed for the analysis of low speed airfoil leading edge transitional separation bubbles. In this analysis an inverse boundary layer finite difference analysis is solved iteratively with a Cauchy integral representation of the inviscid flow which is assumed to be a linear perturbation to a known global viscous airfoil analysis. Favorable comparisons with data indicate the overall validity of the present localized interaction approach. In addition numerical tests were performed to test the sensitivity of the computed results to the mesh size, limits on the Cauchy integral, and the location of the transition region.

Development and application of computational aerothermodynamics flowfield computer codes

NASA Technical Reports Server (NTRS)

Venkatapathy, Ethiraj

1994-01-01

Research was performed in the area of computational modeling and application of hypersonic, high-enthalpy, thermo-chemical nonequilibrium flow (Aerothermodynamics) problems. A number of computational fluid dynamic (CFD) codes were developed and applied to simulate high altitude rocket-plume, the Aeroassist Flight Experiment (AFE), hypersonic base flow for planetary probes, the single expansion ramp model (SERN) connected with the National Aerospace Plane, hypersonic drag devices, hypersonic ramp flows, ballistic range models, shock tunnel facility nozzles, transient and steady flows in the shock tunnel facility, arc-jet flows, thermochemical nonequilibrium flows around simple and complex bodies, axisymmetric ionized flows of interest to re-entry, unsteady shock induced combustion phenomena, high enthalpy pulsed facility simulations, and unsteady shock boundary layer interactions in shock tunnels. Computational modeling involved developing appropriate numerical schemes for the flows on interest and developing, applying, and validating appropriate thermochemical processes. As part of improving the accuracy of the numerical predictions, adaptive grid algorithms were explored, and a user-friendly, self-adaptive code (SAGE) was developed. Aerothermodynamic flows of interest included energy transfer due to strong radiation, and a significant level of effort was spent in developing computational codes for calculating radiation and radiation modeling. In addition, computational tools were developed and applied to predict the radiative heat flux and spectra that reach the model surface.

Study on measurement of leading and trailing edges of blades based on optical scanning system

NASA Astrophysics Data System (ADS)

Chao, Bi; Liu, Hongguang; Bao, Longxiang; Li, Di

2017-10-01

In the field of aeronautics, the geometry and dimensional accuracy of the blade edges has a large influence on the aerodynamic performance of aero engine. Therefore, a non-contact optical scanning system is established to realize the measurement of leading and trailing edges of blades in a rapid, precise and efficient manner in the paper. Based on the mechanical framework of a traditional CMM, the system is equipped with a specified sensing device as the scanning probe, which is made up by two new-style laser scanning sensors installed at a certain angle to each other by a holder. In the measuring procedure, the geometric dimensions of the measured blade edges on every contour plane are determined by the contour information on five transversals at the leading or trailing edges, which can be used to determine the machining allowance of the blades. In order to verify the effectiveness and practicality of the system set up, a precision forging blade after grinded is adopted as the measured object and its leading and trailing edges are measured by the system respectively. In the experiment, the thickness of blade edges on three contour planes is measured by the optical scanning system several times. As the experiment results show, the repeatability accuracy of the system can meet its design requirements and the inspecting demands of the blade edges. As a result, the optical scanning system could serve as a component of the intelligent manufacturing system of blades to improve the machining quality of the blade edges.

An analytically-based method for predicting the noise generated by the interaction between turbulence and a serrated leading edge

NASA Astrophysics Data System (ADS)

Mathews, J. R.; Peake, N.

2018-05-01

This paper considers the interaction of turbulence with a serrated leading edge. We investigate the noise produced by an aerofoil moving through a turbulent perturbation to uniform flow by considering the scattered pressure from the leading edge. We model the aerofoil as an infinite half plane with a leading edge serration, and develop an analytical model using a Green's function based upon the work of Howe. This allows us to consider both deterministic eddies and synthetic turbulence interacting with the leading edge. We show that it is possible to reduce the noise by using a serrated leading edge compared with a straight edge, but the optimal noise-reducing choice of serration is hard to predict due to the complex interaction. We also consider the effect of angle of attack, and find that in general the serrations are less effective at higher angles of attack.

Effect of spanwise blowing on leading-edge vortex bursting of a highly swept aspect ratio 1.18 delta wing

NASA Technical Reports Server (NTRS)

Scantling, W. L.; Gloss, B. B.

1974-01-01

An investigation was conducted in the Langley 1/8-scale V/STOL model tunnel on a semispan delta wing with a leading-edge sweep of 74 deg, to determine the effectiveness of various locations of upper surface and reflection plane blowing on leading-edge vortex bursting. Constant area nozzles were located on the wing upper surface along a ray swept 79 deg, which was beneath the leading-edge vortex core. The bursting and reformation of the leading-edge vortex was viewed by injecting helium into the vortex core, and employing a schlieren system.

Thermostructural applications of heat pipes for cooling leading edges of high-speed aerospace vehicles

NASA Technical Reports Server (NTRS)

Camarda, Charles J.; Glass, David E.

1992-01-01

Heat pipes have been considered for use on wing leading edge for over 20 years. Early concepts envisioned metal heat pipes cooling a metallic leading edge. Several superalloy/sodium heat pipes were fabricated and successfully tested for wing leading edge cooling. Results of radiant heat and aerothermal testing indicate the feasibility of using heat pipes to cool the stagnation region of shuttle-type space transportation systems. The test model withstood a total seven radiant heating tests, eight aerothermal tests, and twenty-seven supplemental radiant heating tests. Cold-wall heating rates ranged from 21 to 57 Btu/sq ft-s and maximum operating temperatures ranged from 1090 to 1520 F. Follow-on studies investigated the application of heat pipes to cool the stagnation regions of single-stage-to-orbit and advanced shuttle vehicles. Results of those studies indicate that a 'D-shaped' structural design can reduce the mass of the heat-pipe concept by over 44 percent compared to a circular heat-pipe geometry. Simple analytical models for heat-pipe startup from the frozen state (working fluid initially frozen) were adequate to approximate transient, startup, and steady-state heat-pipe performance. Improvement in analysis methods has resulted in the development of a finite-element analysis technique to predict heat-pipe startup from the frozen state. However, current requirements of light-weight design and reliability suggest that metallic heat pipes embedded in a refractory composite material should be used. This concept is the concept presently being evaluated for NASP. A refractory-composite/heat-pipe-cooled wing leading edge is currently being considered for the National Aero-Space Plane (NASP). This concept uses high-temperature refractory-metal/lithium heat pipes embedded within a refractory-composite structure and is significantly lighter than an actively cooled wing leading edge because it eliminates the need for active cooling during ascent and descent. Since the

Thermostructural applications of heat pipes for cooling leading edges of high-speed aerospace vehicles

NASA Astrophysics Data System (ADS)

Camarda, Charles J.; Glass, David E.

1992-10-01

Heat pipes have been considered for use on wing leading edge for over 20 years. Early concepts envisioned metal heat pipes cooling a metallic leading edge. Several superalloy/sodium heat pipes were fabricated and successfully tested for wing leading edge cooling. Results of radiant heat and aerothermal testing indicate the feasibility of using heat pipes to cool the stagnation region of shuttle-type space transportation systems. The test model withstood a total seven radiant heating tests, eight aerothermal tests, and twenty-seven supplemental radiant heating tests. Cold-wall heating rates ranged from 21 to 57 Btu/sq ft-s and maximum operating temperatures ranged from 1090 to 1520 F. Follow-on studies investigated the application of heat pipes to cool the stagnation regions of single-stage-to-orbit and advanced shuttle vehicles. Results of those studies indicate that a 'D-shaped' structural design can reduce the mass of the heat-pipe concept by over 44 percent compared to a circular heat-pipe geometry. Simple analytical models for heat-pipe startup from the frozen state (working fluid initially frozen) were adequate to approximate transient, startup, and steady-state heat-pipe performance. Improvement in analysis methods has resulted in the development of a finite-element analysis technique to predict heat-pipe startup from the frozen state. However, current requirements of light-weight design and reliability suggest that metallic heat pipes embedded in a refractory composite material should be used. This concept is the concept presently being evaluated for NASP. A refractory-composite/heat-pipe-cooled wing leading edge is currently being considered for the National Aero-Space Plane (NASP). This concept uses high-temperature refractory-metal/lithium heat pipes embedded within a refractory-composite structure and is significantly lighter than an actively cooled wing leading edge because it eliminates the need for active cooling during ascent and descent. Since the

Test and Analysis of a Hyper-X Carbon-Carbon Leading Edge Chine

NASA Technical Reports Server (NTRS)

Smith, Russell W.; Sikora, Joseph G.; Lindell, Michael C.

2005-01-01

During parts production for the X43A Mach 10 hypersonic vehicle nondestructive evaluation (NDE) of a leading edge chine detected on imbedded delamination near the lower surface of the part. An ultimate proof test was conducted to verify the ultimate strength of this leading edge chine part. The ultimate proof test setup used a pressure bladder design to impose a uniform distributed pressure field over the bi-planar surface of the chine test article. A detailed description of the chine test article and experimental test setup is presented. Analysis results from a linear status model of the test article are also presented and discussed. Post-test inspection of the specimen revealed no visible failures or areas of delamination.

Aerothermodynamic flow phenomena of the airframe-integrated supersonic combustion ramjet

NASA Technical Reports Server (NTRS)

Walton, James T.

1992-01-01

The unique component flow phenomena is discussed of the airframe-integrated supersonic combustion ramjet (scramjet) in a format geared towards new players in the arena of hypersonic propulsion. After giving an overview of the scramjet aerothermodynamic cycle, the characteristics are then covered individually of the vehicle forebody, inlet, combustor, and vehicle afterbody/nozzle. Attention is given to phenomena such as inlet speeding, inlet starting, inlet spillage, fuel injection, thermal choking, and combustor-inlet interaction.

Lesson from Tungsten Leading Edge Heat Load Analysis in KSTAR Divertor

NASA Astrophysics Data System (ADS)

Hong, Suk-Ho; Pitts, Richard Anthony; Lee, Hyeong-Ho; Bang, Eunnam; Kang, Chan-Soo; Kim, Kyung-Min; Kim, Hong-Tack; ITER Organization Collaboration; Kstar Team Team

2016-10-01

An important design issue for the ITER tungsten (W) divertor and in fact for all such components using metallic plasma-facing elements and which are exposed to high parallel power fluxes, is the question of surface shaping to avoid melting of leading edges. We have fabricated a series of tungsten blocks with a variety of leading edge heights (0.3, 0.6, 1.0, and 2.0 mm), from the ITER worst case to heights even beyond the extreme value tested on JET. They are mounted into adjacent, inertially cooled graphite tile installed in the central divertor region of KSTAR, within the field of view of an infra-red (IR) thermography system with a spatial resolution to 0.4 mm/pixel. Adjustment of the outer divertor strike point position is used to deposit power on the different blocks in different discharges. The measured power flux density on flat regions of the surrounding graphite tiles is used to obtain the parallel power flux, q|| impinging on the various W blocks. Experiments have been performed in Type I ELMing H-mode with Ip = 600 kA, BT = 2 T, PNBI = 3.5 MW, leading to a hot attached divertor with typical pulse lengths of 10 s. Three dimensional ANSYS simulations using q|| and assuming geometric projection of the heat flux are found to be consistent with the observed edge loading. This research was partially supported by Ministry of Science, ICT, and Future Planning under KSTAR project.

Controlled vortical flow on delta wings through unsteady leading edge blowing

NASA Technical Reports Server (NTRS)

Lee, K. T.; Roberts, Leonard

1990-01-01

The vortical flow over a delta wing contributes an important part of the lift - the so called nonlinear lift. Controlling this vortical flow with its favorable influence would enhance aircraft maneuverability at high angle of attack. Several previous studies have shown that control of the vortical flow field is possible through the use of blowing jets. The present experimental research studies vortical flow control by applying a new blowing scheme to the rounded leading edge of a delta wing; this blowing scheme is called Tangential Leading Edge Blowing (TLEB). Vortical flow response both to steady blowing and to unsteady blowing is investigated. It is found that TLEB can redevelop stable, strong vortices even in the post-stall angle of attack regime. Analysis of the steady data shows that the effect of leading edge blowing can be interpreted as an effective change in angle of attack. The examination of the fundamental time scales for vortical flow re-organization after the application of blowing for different initial states of the flow field is studied. Different time scales for flow re-organization are shown to depend upon the effective angle of attack. A faster response time can be achieved at angles of attack beyond stall by a suitable choice of the initial blowing momentum strength. Consequently, TLEB shows the potential of controlling the vortical flow over a wide range of angles of attack; i.e., in both for pre-stall and post-stall conditions.

Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 4. Boston Arts Academy

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 6. Perspectives Charter School

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 7. TechBoston Academy

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Application of a flush airdata sensing system to a wing leading edge (LE-FADS)

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Moes, Timothy R.; Czerniejewski, Mark W.; Nichols, Douglas A.

1993-01-01

This paper investigates the feasibility of locating a flush air-data sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered. The leading-edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil, and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the air-data calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush air-data systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50 to 50 deg and dynamic pressures from 3.6 to 11.4 lb/sq ft were performed.

Application of a Flush Airdata Sensing System to a Wing Leading Edge (LE-FADS)

NASA Technical Reports Server (NTRS)

Whitmore, Stephen A.; Moes, Timothy R.; Czerniejewski, Mark W.; Nichols, Douglas A.

1993-01-01

The feasibility of locating a flush airdata sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered is investigated. The leading-edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the airdata calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush airdata systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50 deg to 50 deg and dynamic pressures from 3.6 to 11.4 lb/sq ft were performed.

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

Robins, A. W.; Carlson, H. W.; Mack, R. J.

1980-01-01

Experimental/theoretical correlations are presented which show that significant levels of leading edge thrust are possible at supersonic speeds for certain planforms which match the theoretical thrust distribution potential with the supporting airfoil geometry. The analytical process employed spanwise distribution of both it and/or that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.

Pulsed Film Cooling on a Turbine Blade Leading Edge

DTIC Science & Technology

2009-09-01

LEADING EDGE 1. Introduction Gas turbine engines are based on the Brayton cycle in which atmospheric air is compressed, heated via combustion...generation. Because the working fluid is in an open loop, a cooling process is absent from the Brayton cycle. The ideal Brayton cycle (one in which...Technology, Taylor & Francis, 2000. Harrison, K. and Bogard, D., “CFD Predictions of Film Cooling Adiabatic Effectiveness for Cylindrical Holes Embedded

A computer program for calculating aerodynamic characteristics of low aspect-ratio wings with partial leading-edge separation

NASA Technical Reports Server (NTRS)

Mehrotra, S. C.; Lan, C. E.

1978-01-01

The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.

Visualization of the separation and subsequent transition near the leading edge of airfoils

NASA Technical Reports Server (NTRS)

Arena, A. V.; Mueller, T. J.

1978-01-01

A visual study was performed using the low speed smoke wind tunnels with the objective of obtaining a better understanding of the structure of leading edge separation bubbles on airfoils. The location of separation, transition and reattachment for a cylindrical nose constant-thickness airfoil model were obtained from smoke photographs and surface oil flow techniques. These data, together with static pressure distributions along the leading edge and upper surface of the model, produced the influence of Reynolds number, angle of attack, and trailing edge flap angle on the size and characteristics of the bubble. Additional visual insight into the unsteady nature of the separation bubble was provided by high speed 16 mm movies. The 8 mm color movies taken of the surface oil flow supported the findings of the high speed movies and clearly showed the formation of a scalloped spanwise separation line at the higher Reynolds number.

Aerothermodynamic Analysis of Commercial Experiment Transporter (COMET) Reentry Capsule

NASA Technical Reports Server (NTRS)

Wood, William A.; Gnoffo, Peter A.; Rault, Didier F. G.

1996-01-01

An aerothermodynamic analysis of the Commercial Experiment Transporter (COMET) reentry capsule has been performed using the laminar thin-layer Navier-Stokes solver Langley Aerothermodynamic Upwind Relaxation Algorithm. Flowfield solutions were obtained at Mach numbers 1.5, 2, 5, 10, 15, 20, 25, and 27.5. Axisymmetric and 5, 10, and 20 degree angles of attack were considered across the Mach-number range, with the Mach 25 conditions taken to 90 degrees angle of attack and the Mach 27.5 cases taken to 60 degrees angle of attack. Detailed surface heat-transfer rates were computed at Mach 20 and 25, revealing that heating rates on the heat-shield shoulder ,can exceed the stagnation-point heating by 230 percent. Finite-rate chemistry solutions were performed above Mach 10, otherwise perfect gas computations were made. Drag, lift, and pitching moment coefficients are computed and details of a wake flow are presented. The effect of including the wake in the solution domain was investigated and base pressure corrections to forebody drag coefficients were numerically determined for the lower Mach numbers. Pitching moment comparisons are made with direct simulation Monte Carlo results in the more rarefied flow at the highest Mach numbers, showing agreement within two-percent. Thin-layer Navier-Stokes computations of the axial force are found to be 15 percent higher across the speed range than the empirical/Newtonian based results used during the initial trajectory analyses.

Simulation and Optimization of an Airfoil with Leading Edge Slat

NASA Astrophysics Data System (ADS)

Schramm, Matthias; Stoevesandt, Bernhard; Peinke, Joachim

2016-09-01

A gradient-based optimization is used in order to improve the shape of a leading edge slat upstream of a DU 91-W2-250 airfoil. The simulations are performed by solving the Reynolds-Averaged Navier-Stokes equations (RANS) using the open source CFD code OpenFOAM. Gradients are computed via the adjoint approach, which is suitable to deal with many design parameters, but keeping the computational costs low. The implementation is verified by comparing the gradients from the adjoint method with gradients obtained by finite differences for a NACA 0012 airfoil. The simulations of the leading edge slat are validated against measurements from the acoustic wind tunnel of Oldenburg University at a Reynolds number of Re = 6 • 105. The shape of the slat is optimized using the adjoint approach resulting in a drag reduction of 2%. Although the optimization is done for Re = 6 • 105, the improvements also hold for a higher Reynolds number of Re = 7.9 • 106, which is more realistic at modern wind turbines.

Aero-Thermo-Dynamic Mass Analysis

NASA Astrophysics Data System (ADS)

Shiba, Kota; Yoshikawa, Genki

2016-07-01

Each gas molecule has its own molecular weight, while such a microscopic characteristic is generally inaccessible, and thus, it is measured indirectly through e.g. ionization in conventional mass analysis. Here, we present a novel approach to the direct measurement of molecular weight through a nanoarchitectonic combination of aerodynamics, thermodynamics, and mechanics, transducing microscopic events into macroscopic phenomena. It is confirmed that this approach can provide molecular weight of virtually any gas or vaporizable liquid sample in real-time without ionization. Demonstrations through analytical calculations, numerical simulations, and experiments verify the validity and versatility of the novel mass analysis realized by a simple setup with a flexible object (e.g. with a bare cantilever and even with a business card) placed in a laminar jet. Owing to its unique and simple working principle, this aero-thermo-dynamic mass analysis (AMA) can be integrated into various analytical devices, production lines, and consumer mobile platforms, opening new chapters in aerodynamics, thermodynamics, mechanics, and mass analysis.

Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading edge vortex separation.

Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading- edge vortex separation.

Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M=0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

Subsonic balance and pressure investigation of a 60-deg delta wing with leading-edge devices (data report)

NASA Technical Reports Server (NTRS)

Rao, D. M.; Tingas, S. A.

1981-01-01

The drag reduction potential of leading edge devices on a 60 degree delta wing at high lift was examined. Geometric variations of fences, chordwise slots, pylon type vortex generators, leading edge vortex flaps, and sharp leading edge extensions were tested individually and in specific combinations to improve high-alpha drag performance with a minimum of low-alpha drag penalty. The force, moment, and surface static pressure data for angles of attack up to 23 degrees, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter are documented.

Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development

NASA Technical Reports Server (NTRS)

Jilly, L. F. (Editor)

1970-01-01

The analytical effort was directed towards (1) completing the design of the combustor exit instrumentation assembly, (2) analyzing the coolant flow distribution of the cowl leading edge tip section, (3) determining effects of purge gas pressure on AIM performance analysis, and (4) analyzing heat transfer and associated stress problems related to the cowl leading edge tip section and the nozzle shroud assembly for test conditions.

Experimental study of delta wing leading-edge devices for drag reduction at high lift

NASA Technical Reports Server (NTRS)

Johnson, T. D., Jr.; Rao, D. M.

1982-01-01

The drag reduction devices selected for evaluation were the fence, slot, pylon-type vortex generator, and sharp leading-edge extension. These devices were tested on a 60 degree flatplate delta (with blunt leading edges) in the Langley Research Center 7- by 10-foot high-speed tunnel at low speed and to angles of attack of 28 degrees. Balance and static pressure measurements were taken. The results indicate that all the devices had significant drag reduction capability and improved longitudinal stability while a slight loss of lift and increased cruise drag occurred.

SIMS chemical analysis of extended impacts on the leading and trailing edges of LDEF experiment AO187-2

NASA Technical Reports Server (NTRS)

Amari, S.; Foote, J.; Simon, Charles G.; Swan, P.; Walker, R. M.; Zinner, E.; Jessberger, E. K.; Lange, G.; Stadermann, F.

1992-01-01

The Long Duration Exposure Facility (LDEF) Experiment AO187-2 consisted of 237 capture cells, 120 on the leading edge and 117 on the trailing edge. Each cell was made of polished Ge plates covered with 2.5 micron thick mylar foil at 200 microns from the Ge. Although all leading edge cells and 105 trailing edge cells had lost their plastic covers during flight, optical and electron microscope examination revealed extended impacts in bare cells from either edge that apparently were produced by high velocity projectiles while the plastic foils were still in place. Detailed optical scanning yielded 53 extended impacts on 100 bare cells from the trailing edge that were selected for SIMS chemical analysis. Lateral multi-element ion probe profiles were obtained on 40 of these impacts. Material that can be attributed to the incoming projectiles was found in all analyzed extended compact features and most seem to be associated with cosmic dust particles. However, LDEF deposits are systematically enriched in the refractory elements Al, Ca, and Ti relative to Mg and Fe when compared to IDP's collected in the stratosphere and to chondritic compositions. These differences are most likely due to elemental fractionation effects during the high velocity impact but real differences between interplanetary particles captured on LDEF and stratospheric IDP's cannot be excluded. Recently we extended our studies to cells from the leading edge and the covered cells from the trailing edge. The 12 covered cells contain 20 extended impact candidates. Ion probe analysis of 3 yielded results similar to those obtained for impacts on the bare cells from the trailing edge. Optical scanning of the bare leading edge cell also reveals many extended impacts (42 on 22 cells scanned to date), demonstrating that the cover foils remained intact at least for some time. However, SIMS analysis showed elements that can reasonably be attributed to micrometeoroids in only 2 out of 11 impacts. Eight impacts

Exploratory study of the effects of wing-leading-edge modifications on the stall/spin behavior of a light general aviation airplane

NASA Technical Reports Server (NTRS)

1979-01-01

Configurations with full-span and segmented leading-edge flaps and full-span and segmented leading-edge droop were tested. Studies were conducted with wind-tunnel models, with an outdoor radio-controlled model, and with a full-scale airplane. Results show that wing-leading-edge modifications can produce large effects on stall/spin characteristics, particularly on spin resistance. One outboard wing-leading-edge modification tested significantly improved lateral stability at stall, spin resistance, and developed spin characteristics.

A computational study of incipient leading-edge separation on a 65-deg delta wing at M = 1.60

NASA Technical Reports Server (NTRS)

Mcmillin, S. Naomi; Pittman, James L.; Thomas, James L.

1990-01-01

A computational study on a 65-deg delta wing at a freestream Mach number of 1.60 has been conducted by obtaining conical Reynolds-averaged Navier-Stokes solutions on a parametric series of geometries which varied in leading-edge radius and/or circular-arc camber. The computational results showed that increasing leading-edge radius or camber can delay the onset of leading-edge separation on the leeside of a delta wing at a specific angle of attack. Reynolds number was varied from 1 x 10 to the 6th to 5 x 10 to the 6th for a turbulent boundary-layer and was shown to have a minor effect on the effectiveness of leading-edge radius and/or camber in delaying the onset of leading-edge separation. Both laminar and turbulent boundary-layer models were investigated at a Reynolds number of 1 x 10 to the 6th, and the predicted flow pattern was found to change from attached flow for the turbulent boundary-layer model to separated flow for the laminar boundary-layer model. Based upon these results, three wind-tunnel models have been designed to be tested in the Langley Unitary Plan Wind Tunnel.

Assessment Of The Aerodynamic And Aerothermodynamic Performance Of The USV-3 High-Lift Re-Entry Vehicle

NASA Astrophysics Data System (ADS)

Pezzella, Giuseppe; Richiello, Camillo; Russo, Gennaro

2011-05-01

This paper deals with the aerodynamic and aerothermodynamic trade-off analysis carried out with the aim to design a hypersonic flying test bed (FTB), namely USV3. Such vehicle will have to be launched with a small expendable launcher and shall re-enter the Earth atmosphere allowing to perform several experiments on critical re-entry phenomena. The demonstrator under study is a re-entry space glider characterized by a relatively simple vehicle architecture able to validate hypersonic aerothermodynamic design database and passenger experiments, including thermal shield and hot structures. Then, a summary review of the aerodynamic characteristics of two FTB concepts, compliant with a phase-A design level, has been provided hereinafter. Indeed, several design results, based both on engineering approach and computational fluid dynamics, are reported and discussed in the paper.

Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 8. High Tech High School

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Case Studies of Leading Edge Small Urban High Schools. Personalization Strategic Designs: 9. MetWest High School

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Plasma Irregularities on the Leading and Trailing Edges of Polar Cap Patches

NASA Astrophysics Data System (ADS)

Lamarche, L. J.; Varney, R. H.; Gillies, R.; Chartier, A.; Mitchell, C. N.

2017-12-01

Plasma irregularities in the polar cap have often been attributed to the gradient drift instability (GDI). Traditional fluid theories of GDI predicts irregularity growth only on the trailing edge of polar patches, where the plasma density gradient is parallel to the plasma drift velocity, however many observations show irregularities also form on the leading edge of patches. We consider decameter-scale irregularities detected by polar-latitude SuperDARN (Super Dual Auroral Radar Network) radars with any relationship between the background density gradients and drift velocity. Global electron density from the Multi-Instrument Data Analysis System (MIDAS), a GPS tomography routine, is used to provide context for where irregularities are observed relative to polar patches and finer-scale background density gradients are found from 3D imaging from both the North and Canada faces of the Resolute Bay Incoherent Scatter Radars (RISR-N and RISR-C) jointly. Shear-based instabilities are considered as mechanisms by which plasma irregularities could form on the leading edge of patches. Theoretical predictions of instability growth from both GDI and shear instabilities are compared with irregularity observations for the October 13, 2016 storm.

A Reduced-Complexity Investigation of Blunt Leading-Edge Separation Motivated by UCAV Aerodynamics

NASA Technical Reports Server (NTRS)

Luckring, James M.; Boelens, Okko J.

2015-01-01

A reduced complexity investigation for blunt-leading-edge vortical separation has been undertaken. The overall approach is to design the fundamental work in such a way so that it relates to the aerodynamics of a more complex Uninhabited Combat Air Vehicle (UCAV) concept known as SACCON. Some of the challenges associated with both the vehicle-class aerodynamics and the fundamental vortical flows are reviewed, and principles from a hierarchical complexity approach are used to relate flow fundamentals to system-level interests. The work is part of roughly 6-year research program on blunt-leading-edge separation pertinent to UCAVs, and was conducted under the NATO Science and Technology Organization, Applied Vehicle Technology panel.

Studies on laminar boundary-layer receptivity to freestream turbulence near a leading edge

NASA Technical Reports Server (NTRS)

Kendall, James M.

1991-01-01

An experimental study of the generation of Tollmien-Schlichting waves and wave packets in a flat-plate boundary-layer by weak freestream turbulence has been conducted with the intent of clarifying receptivity mechanisms. Emphasis was placed upon the properties of such waves at stations as far forward as the minimum critical Reynolds number. It was found that alteration of the flow about the leading edge, due either to an asymmetry associated with lift, or due to a change of the fineness ratio of the leading edge, altered the T-S wave amplitude at early stations. The subsequent growth of the waves proceeded faster than expected according to certain stability theory results. Speculation regarding receptivity mechanisms is made.

SIMS chemical analysis of extended impacts on the leading and trailing edges of LDEF experiment AO187-2

NASA Technical Reports Server (NTRS)

Amari, S.; Foote, J.; Swan, P.; Walker, R. M.; Zinner, E.; Lange, G.

1993-01-01

Numerous 'extended impacts' found in both leading and trailing edge capture cells were successfully analyzed for the chemical composition of projectile residues by secondary ion mass spectrometry (SIMS). Most data were obtained from the trailing edge cells where 45 of 58 impacts were classified as 'probably natural' and the remainder as 'possibly man-made debris.' This is in striking contrast to leading edge cells where 9 of 11 impacts so far measured are definitely classified as orbital debris. Although all the leading edge cells had lost their plastic entrance foils during flight, the rate of foil failure was similar to that of the trailing edge cells, 10 percent of which were recovered intact. Ultraviolet embrittlement is suspected as the major cause of failure on both leading and trailing edges. The major impediment to the accurate determination of projectile chemistry is the fractionation of volatile and refractory elements in the hypervelocity impact and redeposition processes. This effect had been noted in a simulation experiment but is more pronounced in the LDEF capture cells, probably due to the higher average velocities of the space impacts. Surface contamination of the pure Ge surfaces with a substance rich in Si, but also containing Mg and Al, provides an additional problem for the accurate determination of impactor chemistry. The effect is variable, being much larger on surfaces that were exposed to space than in those cells that remained intact. Future work will concentrate on the analyses of more leading edge impacts and the development of new SIMS techniques for the measurement of elemental abundances in extended impacts.

Test and Analysis Correlation of Form Impact onto Space Shuttle Wing Leading Edge RCC Panel 8

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Lyle, Karen H.; Gabrys, Jonathan; Melis, Matthew; Carney, Kelly

2004-01-01

Soon after the Columbia Accident Investigation Board (CAIB) began their study of the space shuttle Columbia accident, "physics-based" analyses using LS-DYNA were applied to characterize the expected damage to the Reinforced Carbon-Carbon (RCC) leading edge from high-speed foam impacts. Forensic evidence quickly led CAIB investigators to concentrate on the left wing leading edge RCC panels. This paper will concentrate on the test of the left-wing RCC panel 8 conducted at Southwest Research Institute (SwRI) and the correlation with an LS-DYNA analysis. The successful correlation of the LS-DYNA model has resulted in the use of LS-DYNA as a predictive tool for characterizing the threshold of damage for impacts of various debris such as foam, ice, and ablators onto the RCC leading edge for shuttle return-to-flight.

Numerical investigation of rarefaction effects in the vicinity of a sharp leading edge

NASA Astrophysics Data System (ADS)

Pan, Shaowu; Gao, Zhenxun; Lee, Chunhian

2014-12-01

This paper presents a study of rarefaction effect on hypersonic flow over a sharp leading edge. Both continuum approach and kinetic method: a widely spread commercial Computational Fluid Dynamics-Navior-Stokes-Fourier (CFD-NSF) software - Fluent together with a direct simulation Monte Carlo (DSMC) code developed by the authors are employed for simulation of transition regime with Knudsen number ranging from 0.005 to 0.2. It is found that Fluent can predict the wall fluxes in the case of hypersonic argon flow over the sharp leading edge for the lowest Kn case (Kn = 0.005) in current paper while for other cases it also has a good agreement with DSMC except at the location near the sharp leading edge. Among all of the wall fluxes, it is found that coefficient of pressure is the most sensitive to rarefaction while heat transfer is the least one. A parameter based on translational nonequilibrium and a cut-off value of 0.34 is proposed for continuum breakdown in this paper. The structure of entropy and velocity profile in boundary layer is analyzed. Also, it is found that the ratio of heat transfer coefficient to skin friction coefficient remains uniform along the surface for the four cases in this paper.

Numerical investigation of mist/air impingement cooling on ribbed blade leading-edge surface.

PubMed

Bian, Qingfei; Wang, Jin; Chen, Yi-Tung; Wang, Qiuwang; Zeng, Min

2017-12-01

The working gas turbine blades are exposed to the environment of high temperature, especially in the leading-edge region. The mist/air two-phase impingement cooling has been adopted to enhance the heat transfer on blade surfaces and investigate the leading-edge cooling effectiveness. An Euler-Lagrange particle tracking method is used to simulate the two-phase impingement cooling on the blade leading-edge. The mesh dependency test has been carried out and the numerical method is validated based on the available experimental data of mist/air cooling with jet impingement on a concave surface. The cooling effectiveness on three target surfaces is investigated, including the smooth and the ribbed surface with convex/concave columnar ribs. The results show that the cooling effectiveness of the mist/air two-phase flow is better than that of the single-phase flow. When the ribbed surfaces are used, the heat transfer enhancement is significant, the surface cooling effectiveness becomes higher and the convex ribbed surface presents a better performance. With the enhancement of the surface heat transfer, the pressure drop in the impingement zone increases, but the incremental factor of the flow friction is smaller than that of the heat transfer enhancement. Copyright © 2017 Elsevier Ltd. All rights reserved.

Experimental and Computational Aerothermodynamics of a Mars Entry Vehicle

NASA Technical Reports Server (NTRS)

Hollis, Brian R.

1996-01-01

An aerothermodynamic database has been generated through both experimental testing and computational fluid dynamics simulations for a 70 deg sphere-cone configuration based on the NASA Mars Pathfinder entry vehicle. The aerothermodynamics of several related parametric configurations were also investigated. Experimental heat-transfer data were obtained at hypersonic test conditions in both a perfect gas air wind tunnel and in a hypervelocity, high-enthalpy expansion tube in which both air and carbon dioxide were employed as test gases. In these facilities, measurements were made with thin-film temperature-resistance gages on both the entry vehicle models and on the support stings of the models. Computational results for freestream conditions equivalent to those of the test facilities were generated using an axisymmetric/2D laminar Navier-Stokes solver with both perfect-gas and nonequilibrium thermochemical models. Forebody computational and experimental heating distributions agreed to within the experimental uncertainty for both the perfect-gas and high-enthalpy test conditions. In the wake, quantitative differences between experimental and computational heating distributions for the perfect-gas conditions indicated transition of the free shear layer near the reattachment point on the sting. For the high enthalpy cases, agreement to within, or slightly greater than, the experimental uncertainty was achieved in the wake except within the recirculation region, where further grid resolution appeared to be required. Comparisons between the perfect-gas and high-enthalpy results indicated that the wake remained laminar at the high-enthalpy test conditions, for which the Reynolds number was significantly lower than that of the perfect-gas conditions.

The effect of leading edge tubercles on dynamic stall

NASA Astrophysics Data System (ADS)

Hrynuk, John

The effect of the leading edge tubercles of humpback whales has been heavily studied for their static benefits. These studies have shown that tubercles inhibit flow separation, limit spanwise flow, and extend the operating angle of a wing beyond the static stall point while maintaining lift, all while having a comparatively low negative impact on drag. The current study extends the prior work to investigating the effect of tubercles on dynamic stall, a fundamental flow phenomenon that occurs when wings undergo dynamic pitching motions. Flow fields around the wing models tested were studied using Laser Induced Fluorescence (LIF) and Molecular Tagging Velocimetry (MTV).Resulting velocity fields show that the dynamics of the formation and separation of the leading edge vortex were fundamentally different between the straight wing and the tubercled wing. Tracking of the Dynamic Stall Vortex (DSV) and Shear Layer Vortices (SLVs), which may have a significant impact on the overall flow behavior, was done along with calculations of vortex circulation. Proximity to the wing surface and total circulation were used to evaluate potential dynamic lift increases provided by the tubercles. The effects of pitch rate on the formation process and benefits of the tubercles were also studied and were generally consistent with prior dynamic stall studies. However, tubercles were shown to affect the SLV formation and the circulation differently at higher pitch rates.

Hypersonic research engine project. Phase 2: Aerothermodynamic Integration Model (AIM) test report

NASA Technical Reports Server (NTRS)

Andersen, W. L.; Kado, L.

1975-01-01

The Hypersonic Research Engine-Aerothermodynamic Integration Model (HRE-AIM) was designed, fabricated, and tested in the Hypersonic Tunnel Facility. The HRE-AIM is described along with its installation in the wind tunnel facility. Test conditions to which the HRE-AIM was subjected and observations made during the tests are discussed. The overall engine performance, component interaction, and ignition limits for the design are evaluated.

A Method for Computing Leading-Edge Loads

NASA Technical Reports Server (NTRS)

Rhode, Richard V; Pearson, Henry A

1933-01-01

In this report a formula is developed that enables the determination of the proper design load for the portion of the wing forward of the front spar. The formula is inherently rational in concept, as it takes into account the most important variables that affect the leading-edge load, although theoretical rigor has been sacrificed for simplicity and ease of application. Some empirical corrections, based on pressure distribution measurements on the PW-9 and M-3 airplanes have been introduced to provide properly for biplanes. Results from the formula check experimental values in a variety of cases with good accuracy in the critical loading conditions. The use of the method for design purposes is therefore felt to be justified and is recommended.

An Aeroacoustic Study of a Leading Edge Slat Configuration

NASA Technical Reports Server (NTRS)

Mendoza, J. M.; Brooks, T. F.; Humphreys, W. M., Jr.

2002-01-01

Aeroacoustic evaluations of high-lift devices have been carried out in the Quiet Flow Facility of the NASA Langley Research Center. The present paper describes detailed flow and acoustic measurements that have been made in order to better understand the noise generated from airflow over a wing leading edge slat configuration, and to possibly predict and reduce this noise source. The acoustic database is obtained by a moveable Small Aperture Directional Array of microphones designed to electronically steer to different portions of models under study. The slat is shown to be a uniform distributed noise source. The data was processed such that spectra and directivity were determined with respect to a one-foot span of slat. The spectra are normalized in various fashions to demonstrate slat noise character. In order to equate portions of the spectra to different slat noise components, trailing edge noise predictions using measured slat boundary layer parameters as inputs are compared to the measured slat noise spectra.

Low-Speed Aerodynamic Data for an 0.18-Scale Model of an F-16XL with Various Leading-Edge Modifications

NASA Technical Reports Server (NTRS)

Hahne, Daniel E.

1999-01-01

Using the F-16XL as a test-bed, two strategies for improving the low-speed flying characteristics that had minimal impact on high-speed performance were evaluated. In addition to the basic F-16XL configuration several modifications to the baseline configuration were tested in the Langley 30- X 60-Foot Tunnel: 1) the notched area at the wing leading edge and fuselage juncture was removed resulting in a continuous 70 deg leading-edge sweep on the inboard portion of the wing; 2) an integral attached-flow leading-edge flap concept was added to the continuous leading edge; and 3) a deployable vortex flap concept was added to the continuous leading edge. The purpose of this report is simply to document the test configurations, test conditions, and data obtained in this investigation for future reference and analysis. No analysis is presented herein and the data only appear in tabulated format.

Effect of static shape deformation on aerodynamics and aerothermodynamics of hypersonic inflatable aerodynamic decelerator

NASA Astrophysics Data System (ADS)

Guo, Jinghui; Lin, Guiping; Bu, Xueqin; Fu, Shiming; Chao, Yanmeng

2017-07-01

The inflatable aerodynamic decelerator (IAD), which allows heavier and larger payloads and offers flexibility in landing site selection at higher altitudes, possesses potential superiority in next generation space transport system. However, due to the flexibilities of material and structure assembly, IAD inevitably experiences surface deformation during atmospheric entry, which in turn alters the flowfield around the vehicle and leads to the variations of aerodynamics and aerothermodynamics. In the current study, the effect of the static shape deformation on the hypersonic aerodynamics and aerothermodynamics of a stacked tori Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is demonstrated and analyzed in detail by solving compressible Navier-Stokes equations with Menter's shear stress transport (SST) turbulence model. The deformed shape is obtained by structural modeling in the presence of maximum aerodynamic pressure during entry. The numerical results show that the undulating shape deformation makes significant difference to flow structure. In particular, the more curved outboard forebody surface results in local flow separations and reattachments in valleys, which consequently yields remarkable fluctuations of surface conditions with pressure rising in valleys yet dropping on crests while shear stress and heat flux falling in valleys yet rising on crests. Accordingly, compared with the initial (undeformed) shape, the corresponding differences of surface conditions get more striking outboard, with maximum augmentations of 379 pa, 2224 pa, and 19.0 W/cm2, i.e., 9.8%, 305.9%, and 101.6% for the pressure, shear stress and heat flux respectively. Moreover, it is found that, with the increase of angle of attack, the aerodynamic characters and surface heating vary and the aeroheating disparities are evident between the deformed and initial shape. For the deformable HIAD model investigated in this study, the more intense surface conditions and changed flight

Observations on Leading-Edge Vortex Development

NASA Astrophysics Data System (ADS)

Glenn, Michael; Lang, Amy; Wahidi, Redha; Wilroy, Jacob

2016-11-01

Most of an insect's lift comes from the leading edge vortex (LEV) that they produce when flapping their wings. There are many variables that make a LEV either stronger or weaker such as: roughness from the scales on their wings, angle of attack (AoA) of wing, size of the wing, and speed of the wing during flapping motion. Experiments were conducted to study LEV development to gain a better understanding of butterfly flight and the importance of LEV formation. The variables emphasized in this particular experiment were the chord length Reynolds numbers. Two smooth plates of 4 inches and 7 inches were compared in this experiment with Re of 1500 and 3000. Matlab was used to track the LEV location and calculate the vorticity and circulation magnitudes. Differences in LEV vortex strength as a function of chord length will be presented. Funding was provided by NSF REU site Grant EEC 1358991 and CBET Grant 1628600.

Heat transfer and material temperature conditions in the leading edge area of impingement-cooled turbine vanes

NASA Astrophysics Data System (ADS)

Berg, H. P.; Pfaff, K.; Hennecke, D. K.

The resultant effects on the cooling effectiveness at the leading edge area of an impingement-cooled turbine vane by varying certain geometrical parameters is described with reference to local internal heat transfer coefficients determined from experiment and temperature calculations. The local heat transfer on the cooling-air side is determined experimentally with the aid of the analogy between heat- and mass transfer. The impingement cooling is provided from an inserted sheet-metal containing a single row of holes. The Reynolds Number and several of the cooling geometry parameters were varied. The results demonstrate the high local resolution of the method of measurement, which allows improved analytical treatment of the leading-edge cooling configuration. These experiments also point to the necessity of not always performing model tests under idealized conditions. This becomes very clear in the case of the tests performed on an application-oriented impingement-cooling configuration like that often encountered in engine manufacture. In conclusion, as an example, temperature calculations are employed to demonstrate the effect on the cooling effectiveness of varying the distances between insert and inner surface of the leading edge. It shows how the effectiveness of the leading edge cooling can be increased by simple geometrical measures, which results in a considerable improvement in service life.

Aerothermodynamic Environments Definition for the Mars Science Laboratory Entry Capsule

NASA Technical Reports Server (NTRS)

Edquist, Karl T.; Dyakonov, Artem A.; Wright, Michael J.; Tang, Chun Y.

2007-01-01

An overview of the aerothermodynamic environments definition status is presented for the Mars Science Laboratory entry vehicle. The environments are based on Navier-Stokes flowfield simulations on a candidate aeroshell geometry and worst-case entry heating trajectories. Uncertainties for the flowfield predictions are based primarily on available ground data since Mars flight data are scarce. The forebody aerothermodynamics analysis focuses on boundary layer transition and turbulent heating augmentation. Turbulent transition is expected prior to peak heating, a first for Mars entry, resulting in augmented heat flux and shear stress at the same heatshield location. Afterbody computations are also shown with and without interference effects of reaction control system thruster plumes. Including uncertainties, analysis predicts that the heatshield may experience peaks of 225 W/sq cm for turbulent heat flux, 0.32 atm for stagnation pressure, and 400 Pa for turbulent shear stress. The afterbody heat flux without thruster plume interference is predicted to be 7 W/sq cm on the backshell and 10 W/sq cm on the parachute cover. If the reaction control jets are fired near peak dynamic pressure, the heat flux at localized areas could reach as high as 76 W/sq cm on the backshell and 38 W/sq cm on the parachute cover, including uncertainties. The final flight environments used for hardware design will be updated for any changes in the aeroshell configuration, heating design trajectories, or uncertainties.

Influence of Additional Leading-Edge Surface Roughness on Performances in Highly Loaded Compressor Cascade

NASA Astrophysics Data System (ADS)

Chen, Shaowen; Xu, Hao; Sun, Shijun; Zhang, Longxin; Wang, Songtao

2015-05-01

Experimental research has been carried out at low speed to investigate the effect of additional leading-edge surface roughness on a highly-loaded axial compressor cascade. A 5-hole aerodynamic probe has been traversed across one pitch to obtain the distribution of total pressure loss coefficient, secondary flow vector, flow angles and other aerodynamic parameters at the exit section. Meanwhile, ink-trace flow visualization has been used to measure the flow fields on the walls of cascades and a detailed topology structure of the flow on the walls has been obtained. Aerodynamic parameters and flow characteristics are compared by arranging different levels of roughness on various parts of the leading edge. The results show that adding surface roughness at the leading edge and on the suction side obviously influences cascade performance. Aggravated 3-D flow separation significantly increases the loss in cascades, and the loss increases till 60% when the level of emery paper is 80 mm. Even there is the potential to improve cascade performance in local area of cascade passage. The influence of the length of surface roughness on cascade performance is not always adverse, and which depends on the position of surface roughness.

A Three-Dimensional Solution of Flows over Wings with Leading-Edge Vortex Separation. Part 1: Engineering Document

NASA Technical Reports Server (NTRS)

Brune, G. W.; Weber, J. A.; Johnson, F. T.; Lu, P.; Rubbert, P. E.

1975-01-01

A method of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings with leading-edge vortex separation in incompressible flow is presented. The method employs an inviscid flow model in which the wing and the rolled-up vortex sheets are represented by piecewise, continuous quadratic doublet sheet distributions. The Kutta condition is imposed on all wing edges. Computed results are compared with experimental data and with the predictions of the leading-edge suction analogy for a selected number of wing planforms over a wide range of angle of attack. These comparisons show the method to be very promising, capable of producing not only force predictions, but also accurate predictions of detailed surface pressure distributions, loads, and moments.

Effect of leading-edge roughness on stability and transition of wind turbine blades

NASA Astrophysics Data System (ADS)

Kutz, Douglas; Freels, Justin; Hidore, John; White, Edward

2011-11-01

Over time, wind turbine blades erode due to impacts with sand and other debris. The resulting surface roughness degrades the blades' aerodynamic performance. Experimental studies conducted at the Texas A&M University Low-Speed Wind Tunnel examine roughness effects using a 2D NACA 63-418 airfoil with interchangeable leading edges of varying roughness at chord Reynolds numbers up to 3 . 0 ×106 . These data reveal decreased CL , max and increased CD , min as roughness increases. At very high roughness levels, even the lift curve slope is reduced. To better understand these findings and improve modeling of roughness effects, extensive boundary layer measurements including surface-mounted hotfilms and boundary-layer velocity profiles are used to assess how laminar-to-turbulent transition is promoted by roughness. As expected, roughness accelerates transition. Tollmien-Schlichting (TS) transition is observed only for a smooth leading edge while bypass transition is observed for the moderate and high roughness levels. Results are compared to N-factor transition predictions generated with software used by the wind industry. Predictions are successful for the smooth leading edge but even the low roughness level prevents correct transition prediction using TS-based methods. Support for this work by Vestas Technology Americas, Inc., is gratefully acknowledged as is the support of the wind-energy research group and the Low-Speed Wind Tunnel staff.

Mechanisms of collective cell movement lacking a leading or free front edge in vivo.

PubMed

Uechi, Hiroyuki; Kuranaga, Erina

2017-08-01

Collective cell movement is one of the strategies for achieving the complex shapes of tissues and organs. In this process, multiple cells within a group held together by cell-cell adhesion acquire mobility and move together in the same direction. In some well-studied models of collective cell movement, the mobility depends strongly on traction generated at the leading edge by cells located at the front. However, recent advances in live-imaging techniques have led to the discovery of other types of collective cell movement lacking a leading edge or even a free edge at the front, in a diverse array of morphological events, including tubule elongation, epithelial sheet extension, and tissue rotation. We herein review some of the developmental events that are organized by collective cell movement and attempt to elucidate the underlying cellular and molecular mechanisms, which include membrane protrusions, guidance cues, cell intercalation, and planer cell polarity, or chirality pathways.

Case Studies of Leading Edge Small Urban High Schools. Core Academic Strategic Designs: 3. University Park Campus School

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

SSTAC/ARTS review of the draft Integrated Technology Plan (ITP). Volume 8: Aerothermodynamics Automation and Robotics (A/R) systems sensors, high-temperature superconductivity

DOE Office of Scientific and Technical Information (OSTI.GOV)

Not Available

Viewgraphs of briefings presented at the SSTAC/ARTS review of the draft Integrated Technology Plan (ITP) on aerothermodynamics, automation and robotics systems, sensors, and high-temperature superconductivity are included. Topics covered include: aerothermodynamics; aerobraking; aeroassist flight experiment; entry technology for probes and penetrators; automation and robotics; artificial intelligence; NASA telerobotics program; planetary rover program; science sensor technology; direct detector; submillimeter sensors; laser sensors; passive microwave sensing; active microwave sensing; sensor electronics; sensor optics; coolers and cryogenics; and high temperature superconductivity.

SSTAC/ARTS review of the draft Integrated Technology Plan (ITP). Volume 8: Aerothermodynamics Automation and Robotics (A/R) systems sensors, high-temperature superconductivity

NASA Technical Reports Server (NTRS)

1991-01-01

Viewgraphs of briefings presented at the SSTAC/ARTS review of the draft Integrated Technology Plan (ITP) on aerothermodynamics, automation and robotics systems, sensors, and high-temperature superconductivity are included. Topics covered include: aerothermodynamics; aerobraking; aeroassist flight experiment; entry technology for probes and penetrators; automation and robotics; artificial intelligence; NASA telerobotics program; planetary rover program; science sensor technology; direct detector; submillimeter sensors; laser sensors; passive microwave sensing; active microwave sensing; sensor electronics; sensor optics; coolers and cryogenics; and high temperature superconductivity.

Localized Ras signaling at the leading edge regulates PI3K, cell polarity, and directional cell movement

PubMed Central

Sasaki, Atsuo T.; Chun, Cheryl; Takeda, Kosuke; Firtel, Richard A.

2004-01-01

During chemotaxis, receptors and heterotrimeric G-protein subunits are distributed and activated almost uniformly along the cell membrane, whereas PI(3,4,5)P3, the product of phosphatidylinositol 3-kinase (PI3K), accumulates locally at the leading edge. The key intermediate event that creates this strong PI(3,4,5)P3 asymmetry remains unclear. Here, we show that Ras is rapidly and transiently activated in response to chemoattractant stimulation and regulates PI3K activity. Ras activation occurs at the leading edge of chemotaxing cells, and this local activation is independent of the F-actin cytoskeleton, whereas PI3K localization is dependent on F-actin polymerization. Inhibition of Ras results in severe defects in directional movement, indicating that Ras is an upstream component of the cell's compass. These results support a mechanism by which localized Ras activation mediates leading edge formation through activation of basal PI3K present on the plasma membrane and other Ras effectors required for chemotaxis. A feedback loop, mediated through localized F-actin polymerization, recruits cytosolic PI3K to the leading edge to amplify the signal. PMID:15534002

Subsonic Investigation of Leading-Edge Flaps Designed for Vortex- and Attached-Flow on a High-Speed Civil Transport Configuration

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Kemmerly, Guy T.; Kjerstad, Kevin J.; Lessard, Victor R.

1999-01-01

A wind tunnel investigation of two separate leading-edge flaps, designed for vortex and attached-flow, respectively, were conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.12 to 0.27, with corresponding chord Reynolds numbers of 2.50 x 10 (sup 6) to 5.50 x 10 (sup 6). Variations of the leading-edge flap deflection angle were tested with outboard leading-edge flaps deflected 0 deg. and 26.4 deg. Trailing-edge flaps were deflected 0 deg., 10 deg., 12.9 deg., and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein. The data associated with each deflected leading-edge flap indicate L/D improvements over the undeflected configuration. These improvements may be instrumental in providing the necessary lift augmentation required by an actual HSCT during the climb-out and landing phases of the flight envelope. However, further tests will have to be done to assess their full potential.

Navier-Stokes analysis of airfoils with leading edge ice accretions

NASA Technical Reports Server (NTRS)

Potapczuk, Mark G.

1993-01-01

A numerical analysis of the flowfield characteristics and the performance degradation of an airfoil with leading edge ice accretions was performed. The important fluid dynamic processes were identified and calculated. Among these were the leading edge separation bubble at low angles of attack, complete separation on the low pressure surface resulting in premature shell, drag rise due to the ice shape, and the effects of angle of attack on the separated flow field. Comparisons to experimental results were conducted to confirm these calculations. A computer code which solves the Navier-Stokes equations in two dimensions, ARC2D, was used to perform the calculations. A Modified Mixing Length turbulence model was developed to produce grids for several ice shape and airfoil combinations. Results indicate that the ability to predict overall performance characteristics, such as lift and drag, at low angles of attack is excellent. Transition location is important for accurately determining separation bubble shape. Details of the flowfield in and downstream of the separated regions requires some modifications. Calculations for the stalled airfoil indicate periodic shedding of vorticity that was generated aft of the ice accretion. Time averaged pressure values produce results which compare favorably with experimental information. A turbulence model which accounts for the history effects in the flow may be justified.

Influence of the Angle of Attack on the Aerothermodynamics of the Mars Science Laboratory

NASA Technical Reports Server (NTRS)

Dyakonov, Artem A.; Edquist, Karl T.; Schoenenberger, Mark

2006-01-01

An investigation of the effects of the incidence angle on the aerothermodynamic environments of the Mars Science Laboratory has been conducted. Flight conditions of peak heating, peak deceleration and chute deploy are selected and the effects of the angle of attack on the aerodynamics and aerothermodynamics are analyzed. The investigation found that static aerodynamics are well behaved within the considered range of incidence angles. Leeside laminar and turbulent computed heating rates decrease with incidence, despite the increase in the leeside running length. Stagnation point was found to stay on the conical flank at all angles of attack, and this is linked to the rapid flow expansion around the shoulder. Hypersonic lift to drag ratio is limited by the heating rates in the region of the windside shoulder. The effects of the high angle of incidence on the dynamic aero at low Mach remains to be determined. Influence of the angle of attack on the smooth-wall transition parameter indicates, that higher angle of attack flight may result in delayed turbulence onset, however, a coupled analysis, involving flight trajectory simulation is necessary.

Formation and Development of the Dynamic Stall Vortex on a Wing with Leading Edge Tubercles

NASA Astrophysics Data System (ADS)

Hrynuk, John; Bohl, Douglas

2015-11-01

Humpback whales are unique in that their flippers have leading edge ``bumps'' or tubercles. Past work on airfoils inspired by whale flippers has centered on the static aerodynamic characteristics of these airfoils. The current study uses Molecular Tagging Velocimetry (MTV) to investigate the effects of tubercles on dynamically pitching NACA 0012 airfoils. A baseline (i.e. straight leading edge) wing and one modified with leading edge tubercles are investigated. Tracking of the Dynamic Stall Vortex (DSV) is performed to quantitatively compare the DSV formation location, path, and convective velocity for tubercled and baseline wings. The results show that there is a spanwise variation in the initial formation location and motion of the DSV on the modified wing. Once formed, the DSV aligns into a more uniform spanwise structure. As the pitching motion progresses, the DSV on the modified wing convects away from the airfoil surface later and slower than is observed for the baseline airfoil. The results indicate that the tubercles may delay stall when compared to the baseline airfoil. This work was supported by NSF Grant # 0845882.

Guidance signalling regulates leading edge behaviour during collective cell migration of cardiac cells in Drosophila.

PubMed

Raza, Qanber; Jacobs, J Roger

2016-11-15

Collective cell migration is the coordinated movement of cells, which organize tissues during morphogenesis, repair and some cancers. The motile cell membrane of the advancing front in collective cell migration is termed the Leading Edge. The embryonic development of the vertebrate and Drosophila hearts are both characterized by the coordinated medial migration of a bilateral cluster of mesodermal cells. In Drosophila, the cardioblasts form cohesive bilateral rows that migrate collectively as a unit towards the dorsal midline to form the dorsal vessel. We have characterized the collective cell migration of cardioblasts as an in vivo quantitative model to study the behaviour of the Leading Edge. We investigated whether guidance signalling through Slit and Netrin pathways plays a role in cell migration during heart development. Through time-lapse imaging and quantitative assessment of migratory behaviour of the cardioblasts in loss-of-function mutants, we demonstrate that both Slit and Netrin mediated signals are autonomously and concomitantly required to maximize migration velocity, filopodial and lamellipodial activities. Additionally, we show that another Slit and Netrin receptor, Dscam1, the role of which during heart development was previously unknown, is required for both normal migration of cardioblasts and luminal expansion. Leading edge behaviour analysis revealed a dosage dependent genetic interaction between Slit and Netrin receptors suggesting that downstream signalling through these receptors converge on a common output that increases leading edge activity of the cardioblasts. Finally, we found that guidance signalling maintains the balance between epithelial and mesenchymal characteristics of the migrating cardioblasts. Copyright © 2016 Elsevier Inc. All rights reserved.

Heat transfer characteristics of hypersonic waveriders with an emphasis on the leading edge effects. M.S. Thesis, 1991

NASA Technical Reports Server (NTRS)

Vanmol, Denis O.; Anderson, John D., Jr.

1992-01-01

The heat transfer characteristics in surface radiative equilibrium and the aerodynamic performance of blunted hypersonic waveriders are studied along two constant dynamic pressure trajectories for four different Mach numbers. The inviscid leading edge drag was found to be a small (4 to 8 percent) but not negligible fraction of the inviscid drag of the vehicle. Although the viscous drag at the leading edge can be neglected, the presence of the leading edge will influence the transition pattern of the upper and the lower surfaces and therefore affect the viscous drag of the entire vehicle. For an application similar to the National Aerospace Plane (NASP), the present study demonstrates that the waverider remains a valuable concept at high Mach numbers if a state-of-the-art active cooling device is used along the leading edge. At low Mach number (less than 5), the study shows the surface radiative cooling might be sufficient. In all cases, radiative cooling is sufficient for the upper and lower surfaces of the vehicle if ceramic composites are used as thermal protection.

Metallic Concepts for Repair of Reinforced Carbon-Carbon Space Shuttle Leading Edges

NASA Technical Reports Server (NTRS)

Ritzert, Frank; Nesbitt, James

2007-01-01

The Columbia accident has focused attention on the critical need for on-orbit repair concepts for wing leading edges in the event that potentially catastrophic damage is incurred during Space Shuttle Orbiter flight. The leading edge of the space shuttle wings consists of a series of eleven panels on each side of the orbiter. These panels are fabricated from reinforced carbon-carbon (RCC) which is a light weight composite with attractive strength at very high temperatures. The damage that was responsible for the loss of the Colombia space shuttle was deemed due to formation of a large hole in one these RCC leading edge panels produced by the impact of a large piece of foam. However, even small cracks in the RCC are considered as potentially catastrophic because of the high temperature re-entry environment. After the Columbia accident, NASA has explored various means to perform on-orbit repairs in the event that damage is sustained in future shuttle flights. Although large areas of damage, such as that which doomed Columbia, are not anticipated to re-occur due to various improvements to the shuttle, especially the foam attachment, NASA has also explored various options for both small and large area repair. This paper reports one large area repair concept referred to as the "metallic over-wrap." Environmental conditions during re-entry of the orbiter impose extreme requirements on the RCC leading edges as well as on any repair concepts. These requirements include temperatures up to 3000 F (1650 C) for up to 15 minutes in the presence of an extremely oxidizing plasma environment. Figure 1 shows the temperature profile across one panel (#9) which is subject to the highest temperatures during re-entry. Although the RCC possesses adequate mechanical strength at these temperatures, it lacks oxidation resistance. Oxidation protection is afforded by converting the outer layers of the RCC to SiC by chemical vapor deposition (CVD). At high temperatures in an oxidizing

Summary of past experience in natural laminar flow and experimental program for resilient leading edge

NASA Technical Reports Server (NTRS)

Carmichael, B. H.

1979-01-01

The potential of natural laminar flow for significant drag reduction and improved efficiency for aircraft is assessed. Past experience with natural laminar flow as reported in published and unpublished data and personal observations of various researchers is summarized. Aspects discussed include surface contour, waviness, and smoothness requirements; noise and vibration effects on boundary layer transition, boundary layer stability criteria; flight experience with natural laminar flow and suction stabilized boundary layers; and propeller slipstream, rain, frost, ice and insect contamination effects on boundary layer transition. The resilient leading edge appears to be a very promising method to prevent leading edge insect contamination.

LFC leading edge glove flight: Aircraft modification design, test article development and systems integration

NASA Technical Reports Server (NTRS)

Etchberger, F. R.

1983-01-01

Reduction of skin friction drag by suction of boundary layer air to maintain laminar flow has been known since Prandtl's published work in 1904. The dramatic increases in fuel costs and the potential for periods of limited fuel availability provided the impetus to explore technologies to reduce transport aircraft fuel consumption. NASA sponsored the Aircraft Energy Efficiency (ACEE) program in 1976 to develop technologies to improve fuel efficiency. This report documents the Lockheed-Georgia Company accomplishments in designing and fabricating a leading-edge flight test article incorporating boundary layer suction slots to be flown by NASA on their modified JetStar aircraft. Lockheed-Georgia Company performed as the integration contractor to design the JetStar aircraft modification to accept both a Lockheed and a McDonnell Douglas flight test article. McDonnell Douglas uses a porous skin concept. The report describes aerodynamic analyses, fabrication techniques, JetStar modifications, instrumentation requirements, and structural analyses and testing for the Lockheed test article. NASA will flight test the two LFC leading-edge test articles in a simulated commercial environment over a 6 to 8 month period in 1984. The objective of the flight test program is to evaluate the effectiveness of LFC leading-edge systems in reducing skin friction drag and consequently improving fuel efficiency.

A Factor Affecting Transonic Leading-edge Flow Separation

NASA Technical Reports Server (NTRS)

Wood, George P; Gooderum, Paul B

1956-01-01

A change in flow pattern that was observed as the free-stream Mach number was increased in the vicinity of 0.8 was described in NACA Technical Note 1211 by Lindsey, Daley, and Humphreys. The flow on the upper surface behind the leading edge of an airfoil at an angle of attack changed abruptly from detached flow with an extensive region of separation to attached supersonic flow terminated by a shock wave. In the present paper, the consequences of shock-wave - boundary layer interaction are proposed as a factor that may be important in determining the conditions under which the change in flow pattern occurs. Some experimental evidence in support of the importance of this factor is presented.

Numerical evaluations of the effect of leading-edge protuberances on the static and dynamic stall characteristics of an airfoil

NASA Astrophysics Data System (ADS)

Cai, C.; Zuo, Z. G.; Liu, S. H.; Wu, Y. L.; Wang, F. B.

2013-12-01

Wavy leading edge modifications of airfoils through imitating humpback whale flippers has been considered as a viable passive way to control flow separation. In this paper, flows around a baseline 634-021 airfoil and one with leading-edge sinusoidal protuberances were simulated using S-A turbulence model. When studying the static stall characteristics, it is found that the modified airfoil does not stall in the traditional manner, with increasing poststall lift coefficients. At high angles of attack, the flows past the wavy leading edge stayed attached for a distance, while the baseline foil is in a totally separated flow condition. On this basis, the simulations of pitch characteristic were carried out for both foils. At high angles of attack mild variations in lift and drag coefficients of the modified foil can be found, leading to a smaller area of hysteresis loop. The special structure of wavy leading edge can help maintain high consistency of the flow field in dynamic pitching station within a particular range of angles of attack.

A Mesh Refinement Study on the Impact Response of a Shuttle Leading-Edge Panel Finite Element Simulation

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.; Lyle, Karen H.; Spellman, Regina L.

2006-01-01

A study was performed to examine the influence of varying mesh density on an LS-DYNA simulation of a rectangular-shaped foam projectile impacting the space shuttle leading edge Panel 6. The shuttle leading-edge panels are fabricated of reinforced carbon-carbon (RCC) material. During the study, nine cases were executed with all possible combinations of coarse, baseline, and fine meshes of the foam and panel. For each simulation, the same material properties and impact conditions were specified and only the mesh density was varied. In the baseline model, the shell elements representing the RCC panel are approximately 0.2-in. on edge, whereas the foam elements are about 0.5-in. on edge. The element nominal edge-length for the baseline panel was halved to create a fine panel (0.1-in. edge length) mesh and doubled to create a coarse panel (0.4-in. edge length) mesh. In addition, the element nominal edge-length of the baseline foam projectile was halved (0.25-in. edge length) to create a fine foam mesh and doubled (1.0-in. edge length) to create a coarse foam mesh. The initial impact velocity of the foam was 775 ft/s. The simulations were executed in LS-DYNA for 6 ms of simulation time. Contour plots of resultant panel displacement and effective stress in the foam were compared at four discrete time intervals. Also, time-history responses of internal and kinetic energy of the panel, kinetic and hourglass energy of the foam, and resultant contact force were plotted to determine the influence of mesh density.

Flow Field Characteristics of Finite-span Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid; Office of Naval Research Collaboration

2011-11-01

Past work has shown that humpback whale-like leading edge protuberances can significantly alter the load characteristics of both 2D and finite-span hydrofoils. To understand the mechanisms responsible for observed performance changes, the flow field characteristics of a baseline hydrofoil and models with leading edge protuberances were examined using the Stereo Particle Image Velocimetry (SPIV) technique. The near surface flow field on the hydrofoils was measured along with the tip vortex flow field on finite-span hydrofoils. Angles of attack ranging from 6 to 24 degrees were examined at freestream velocities of 1.8 m/s and 4.5 m/s, corresponding to Reynolds numbers of 180 and 450 thousand, respectively. While Reynolds number does not play a major role in establishing the flow field trends, both the protuberance geometry and spatial proximity to protuberances affect the velocity and vorticity characteristics near the foil surface, and in the wake and tip vortex. Near surface measurements reveal counter-rotating vortices on protuberance shoulders, while tip vortex measurements show that streamwise vorticity can be strongly affected by the presence of protuberances. The observed flow field characteristics will be presented. Sponsored by the ONR-ULI program.

User's Manual for the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA)

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Cheatwood, F. McNeil

1996-01-01

This user's manual provides detailed instructions for the installation and the application of version 4.1 of the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA). Also provides simulation of flow field in thermochemical nonequilibrium around vehicles traveling at hypersonic velocities through the atmosphere. Earlier versions of LAURA were predominantly research codes, and they had minimal (or no) documentation. This manual describes UNIX-based utilities for customizing the code for special applications that also minimize system resource requirements. The algorithm is reviewed, and the various program options are related to specific equations and variables in the theoretical development.

Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

NASA Technical Reports Server (NTRS)

Hernandez, Gloria; Wood, Richard M.; Covell, Peter F.

1994-01-01

An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Low-speed cascade investigation of loaded leading-edge compressor blades

NASA Technical Reports Server (NTRS)

Emery, James C

1956-01-01

Six percent thick NACA 63-series compressor-blade sections having a loaded leading-edge A4K6 mean line have been investigated systematically in a two-dimensional porous-wall cascade over a range of Reynolds numbers from 160,000 to 385,000. Blades cambered to have isolated-airfoil lift coefficients of 0.6, 1.2, 1.8, and 2.4 were tested over the usable angle-of-attack range at inlet-air angles of 30 degrees, 45 degrees, and 60 degrees and solidities of 1.0 and 1.5. A comparison with data of NACA RM L51G31, shows that the angle-of-attack operating range is 2 degrees to 4 degrees less than the range for the uniformly loaded section; however, the wake losses near design angle of attack are slightly lower than those for the uniformly loaded section. Except for highly cambered blades at high inlet angles, the 63-(C s oA4K6)06 compressor-blade sections are capable of more efficient operation for moderate-speed subsonic compressors at design angle of attack than are the 65-(C s oa10)10 or the 65-(c s oA2I8b)10 compressor-blade sections. In contrast to the other sections, the loaded leading-edge sections are capable of operating efficiently at the lower Reynolds numbers.

Pressure distributions from subsonic tests of an advanced laminar-flow-control wing with leading- and trailing-edge flaps

NASA Technical Reports Server (NTRS)

Applin, Zachary T.; Gentry, Garl L., Jr.

1988-01-01

An unswept, semispan wing model equipped with full-span leading- and trailing-edge flaps was tested in the Langley 14- by 22-Foot Subsonic Tunnel to determine the effect of high-lift components on the aerodynamics of an advanced laminar-flow-control (LFC) airfoil section. Chordwise pressure distributions near the midsemispan were measured for four configurations: cruise, trailing-edge flap only, and trailing-edge flap with a leading-edge Krueger flap of either 0.10 or 0.12 chord. Part 1 of this report (under separate cover) presents a representative sample of the plotted pressure distribution data for each configuration tested. Part 2 presents the entire set of plotted and tabulated pressure distribution data. The data are presented without analysis.

Case Studies of Leading Edge Small Urban High Schools. Core Academic Strategic Designs: 1. Academy of the Pacific Rim

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Case Studies of Leading Edge Small Urban High Schools. Core Academic Strategic Designs: 2. Noble Street Charter High School

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 5. Life Academy of Health and Bioscience

ERIC Educational Resources Information Center

Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

2008-01-01

This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high…

Mass loss of TEOS-coated RCC subjected to the environment at the shuttle wing leading edge

NASA Technical Reports Server (NTRS)

Stroud, C. W.; Rummler, D. R.

1981-01-01

Coated, reinforced carbon-carbon (RCC) is used for the leading edges of the Space Shuttle. The mass loss characteristics of RCC specimens coated with tetra-ethyl-ortho-silicate (TEOS) were determined for conditions which simulated the entry environment expected at the stagnation area of the wing leading edge. Maximum specimen temperature was 1632 K. Specimens were exposed for up to 100 missions. Stress levels up to 8.274 MPa caused an average increase in oxidation of 6 percent over unstressed specimens. Experimentally determined mass losses were compared with those predicted by an existing empirical analysis.

Characterization of Lift and Drag on Two Dimensional Airfoils with and without Sinusoidal Leading Edges

NASA Astrophysics Data System (ADS)

Acosta, Gregorio I.

An experimental investigation was taken on a 63-021 NACA airfoil, to characterize lift and drag and how the effects of sinusoidal leading edges affect the aerodynamic properties. A theoretical model is also purposed by implementing a perturbation on thin-airfoil theory. Two sets of airfoils were machined and tested inside a low-speed open circuit wind tunnel. Data from a pressure scanner and particle image velocity will give an insight of how the modified leading edges affect the aerodynamic properties. A Fourier series expansion was used to solve for the lifting-line model, by use of thin-airfoil theory and complex number theory.

Cooling Strategies for Vane Leading Edges in a Syngas Environment Including Effects of Deposition and Turbulence

DOE Office of Scientific and Technical Information (OSTI.GOV)

Ames, Forrest; Bons, Jeffrey

2014-09-30

The Department of Energy has goals to move land based gas turbine systems to alternate fuels including coal derived synthetic gas and hydrogen. Coal is the most abundant energy resource in the US and in the world and it is economically advantageous to develop power systems which can use coal. Integrated gasification combined cycles are (IGCC) expected to allow the clean use of coal derived fuels while improving the ability to capture and sequester carbon dioxide. These cycles will need to maintain or increase turbine entry temperatures to develop competitive efficiencies. The use of coal derived syngas introduces a rangemore » of potential contaminants into the hot section of the gas turbine including sulfur, iron, calcium, and various alkali metals. Depending on the effectiveness of the gas clean up processes, there exists significant likelihood that the remaining materials will become molten in the combustion process and potentially deposit on downstream turbine surfaces. Past evidence suggests that deposition will be a strong function of increasing temperature. Currently, even with the best gas cleanup processes a small level of particulate matter in the syngas is expected. Consequently, particulate deposition is expected to be an important consideration in the design of turbine components. The leading edge region of first stage vanes most often have higher deposition rates than other areas due to strong fluid acceleration and streamline curvature in the vicinity of the surface. This region remains one of the most difficult areas in a turbine nozzle to cool due to high inlet temperatures and only a small pressure ratio for cooling. The leading edge of a vane often has relatively high heat transfer coefficients and is often cooled using showerhead film cooling arrays. The throat of the first stage nozzle is another area where deposition potentially has a strongly adverse effect on turbine performance as this region meters the turbine inlet flow. Based on

A review of models of fluctuating protrusion and retraction patterns at the leading edge of motile cells.

PubMed

Ryan, Gillian L; Watanabe, Naoki; Vavylonis, Dimitrios

2012-04-01

A characteristic feature of motile cells as they undergo a change in motile behavior is the development of fluctuating exploratory motions of the leading edge, driven by actin polymerization. We review quantitative models of these protrusion and retraction phenomena. Theoretical studies have been motivated by advances in experimental and computational methods that allow controlled perturbations, single molecule imaging, and analysis of spatiotemporal correlations in microscopic images. To explain oscillations and waves of the leading edge, most theoretical models propose nonlinear interactions and feedback mechanisms among different components of the actin cytoskeleton system. These mechanisms include curvature-sensing membrane proteins, myosin contraction, and autocatalytic biochemical reaction kinetics. We discuss how the combination of experimental studies with modeling promises to quantify the relative importance of these biochemical and biophysical processes at the leading edge and to evaluate their generality across cell types and extracellular environments. Copyright © 2012 Wiley Periodicals, Inc.

Leading edge vortex control on a delta wing with dielectric barrier discharge plasma actuators

NASA Astrophysics Data System (ADS)

Shen, Lu; Wen, Chih-yung

2017-06-01

This paper presents an experimental investigation of the application of dielectric barrier discharge (DBD) plasma actuators on a slender delta wing to control the leading edge vortices (LEVs). The experiments are conducted in a wind tunnel with a Reynolds number of 50 000 based on the chord length. The smoke flow visualization reveals that the DBD plasma actuators at the leading edges significantly modify the vortical flow structure over the delta wing. It is noted that symmetric control at both semi-spans and asymmetric control at a single semi-span leads to opposite effects on the local LEVs. Particle image velocimetry (PIV) indicates that the shear layer is deformed by the actuators. Therefore, both the strength and the shape of the LEV cores are deeply affected. The six-component force measurement shows that the DBD plasma actuators have a limited effect on lift and drag while inducing relatively large moments. This suggests that the DBD plasma actuator is a promising technique for delta wing maneuvering.

A comparison of experimental and calculated thin-shell leading-edge buckling due to thermal stresses

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.

1988-01-01

High-temperature thin-shell leading-edge buckling test data are analyzed using NASA structural analysis (NASTRAN) as a finite element tool for predicting thermal buckling characteristics. Buckling points are predicted for several combinations of edge boundary conditions. The problem of relating the appropriate plate area to the edge stress distribution and the stress gradient is addressed in terms of analysis assumptions. Local plasticity was found to occur on the specimen analyzed, and this tended to simplify the basic problem since it effectively equalized the stress gradient from loaded edge to loaded edge. The initial loading was found to be difficult to select for the buckling analysis because of the transient nature of thermal stress. Multiple initial model loadings are likely required for complicated thermal stress time histories before a pertinent finite element buckling analysis can be achieved. The basic mode shapes determined from experimentation were correctly identified from computation.

Increased heat transfer to a cylindrical leading edge due to spanwise variations in the freestream velocity

NASA Technical Reports Server (NTRS)

Rigby, D. L.; Vanfossen, G. J.

1991-01-01

The present study numerically demonstrates how small spanwise variations in velocity upstream of a body can cause relatively large increases in the spanwise-averaged heat transfer to the leading edge. Vorticity introduced by spanwise variations, first decays as it drifts downstream, then amplifies in the stagnation region as a result of vortex stretching. This amplification can cause a periodic array of 3 D structures, similar to horseshoe vortices, to form. The numerical results indicate that, for the given wavelength, there is an amplitude threshold below which a structure does not form. A one-dimensional analysis, to predict the decay of vorticity in the absence of the body, in conjunction with the full numerical results indicated that the threshold is more accurately stated as minimum level of vorticity required in the leading edge region for a structure to form. It is possible, using the one-dimensional analysis, to compute an optimum wavelength in terms of the maximum vorticity reaching the leading edge region for given amplitude. A discussion is presented which relates experimentally observed trends to the trends of the present phenomena.

Velocity and Drag Evolution From the Leading Edge of a Model Mangrove Forest

NASA Astrophysics Data System (ADS)

Maza, Maria; Adler, Katherine; Ramos, Diogo; Garcia, Adrian Mikhail; Nepf, Heidi

2017-11-01

An experimental study of unidirectional flow through a model mangrove forest measured both velocity and forces on individual trees. The individual trees were 1/12th scale models of mature Rhizophora, including 24 prop roots distributed in a three-dimensional layout. Thirty-two model trees were distributed in a staggered array producing a 2.5 m long forest. The velocity evolved from a boundary layer profile at the forest leading edge to a vertical profile determined by the vertical distribution of frontal area, with significantly higher velocity above the prop roots. Fully developed conditions were reached at the fifth tree row from the leading edge. Within the root zone the velocity was reduced by up to 50% and the TKE was increased by as much as fivefold, relative to the upstream conditions. TKE in the root zone was mainly produced by root and trunk wakes, and it agreed in magnitude with the estimation obtained using the Tanino and Nepf (2008) formulation. Maximum TKE occurred at the top of the roots, where a strong shear region was associated with the change in frontal area. The drag measured on individual trees decreased from the leading edge and reached a constant value at the fifth row and beyond, i.e., in the fully developed region. The drag exhibited a quadratic dependence on velocity, which justified the definition of a quadratic drag coefficient. Once the correct drag length-scale was defined, the measured drag coefficients collapsed to a single function of Reynolds number.

The Effect of Leading-Edge Sweep and Surface Inclination on the Hypersonic Flow Field Over a Blunt Flat Plate

NASA Technical Reports Server (NTRS)

Creager, Marcus O.

1959-01-01

An investigation of the effects of variation of leading-edge sweep and surface inclination on the flow over blunt flat plates was conducted at Mach numbers of 4 and 5.7 at free-stream Reynolds numbers per inch of 6,600 and 20,000, respectively. Surface pressures were measured on a flat plate blunted by a semicylindrical leading edge over a range of sweep angles from 0 deg to 60 deg and a range of surface inclinations from -10 deg to +10 deg. The surface pressures were predicted within an average error of +/- 8 percent by a combination of blast-wave and boundary-layer theory extended herein to include effects of sweep and surface inclination. This combination applied equally well to similar data of other investigations. The local Reynolds number per inch was found to be lower than the free-stream Reynolds number per inch. The reduction in local Reynolds number was mitigated by increasing the sweep of the leading edge. Boundary-layer thickness and shock-wave shape were changed little by the sweep of the leading edge.

A PKC-MARCKS-PI3K regulatory module links Ca2+ and PIP3 signals at the leading edge of polarized macrophages.

PubMed

Ziemba, Brian P; Falke, Joseph J

2018-01-01

The leukocyte chemosensory pathway detects attractant gradients and directs cell migration to sites of inflammation, infection, tissue damage, and carcinogenesis. Previous studies have revealed that local Ca2+ and PIP3 signals at the leading edge of polarized leukocytes play central roles in positive feedback loop essential to cell polarization and chemotaxis. These prior studies showed that stimulation of the leading edge Ca2+ signal can strongly activate PI3K, thereby triggering a larger PIP3 signal, but did not elucidate the mechanistic link between Ca2+ and PIP3 signaling. A hypothesis explaining this link emerged, postulating that Ca2+-activated PKC displaces the MARCKS protein from plasma membrane PIP2, thereby releasing sequestered PIP2 to serve as the target and substrate lipid of PI3K in PIP3 production. In vitro single molecule studies of the reconstituted pathway on lipid bilayers demonstrated the feasibility of this PKC-MARCKS-PI3K regulatory module linking Ca2+ and PIP3 signals in the reconstituted system. The present study tests the model predictions in live macrophages by quantifying the effects of: (a) two pathway activators-PDGF and ATP that stimulate chemoreceptors and Ca2+ influx, respectively; and (b) three pathway inhibitors-wortmannin, EGTA, and Go6976 that inhibit PI3K, Ca2+ influx, and PKC, respectively; on (c) four leading edge activity sensors-AKT-PH-mRFP, CKAR, MARCKSp-mRFP, and leading edge area that report on PIP3 density, PKC activity, MARCKS membrane binding, and leading edge expansion/contraction, respectively. The results provide additional evidence that PKC and PI3K are both essential elements of the leading edge positive feedback loop, and strongly support the existence of a PKC-MARCKS-PI3K regulatory module linking the leading edge Ca2+ and PIP3 signals. As predicted, activators stimulate leading edge PKC activity, displacement of MARCKS from the leading edge membrane and increased leading edge PIP3 levels, while inhibitors

A PKC-MARCKS-PI3K regulatory module links Ca2+ and PIP3 signals at the leading edge of polarized macrophages

PubMed Central

Ziemba, Brian P.

2018-01-01

The leukocyte chemosensory pathway detects attractant gradients and directs cell migration to sites of inflammation, infection, tissue damage, and carcinogenesis. Previous studies have revealed that local Ca2+ and PIP3 signals at the leading edge of polarized leukocytes play central roles in positive feedback loop essential to cell polarization and chemotaxis. These prior studies showed that stimulation of the leading edge Ca2+ signal can strongly activate PI3K, thereby triggering a larger PIP3 signal, but did not elucidate the mechanistic link between Ca2+ and PIP3 signaling. A hypothesis explaining this link emerged, postulating that Ca2+-activated PKC displaces the MARCKS protein from plasma membrane PIP2, thereby releasing sequestered PIP2 to serve as the target and substrate lipid of PI3K in PIP3 production. In vitro single molecule studies of the reconstituted pathway on lipid bilayers demonstrated the feasibility of this PKC-MARCKS-PI3K regulatory module linking Ca2+ and PIP3 signals in the reconstituted system. The present study tests the model predictions in live macrophages by quantifying the effects of: (a) two pathway activators—PDGF and ATP that stimulate chemoreceptors and Ca2+ influx, respectively; and (b) three pathway inhibitors—wortmannin, EGTA, and Go6976 that inhibit PI3K, Ca2+ influx, and PKC, respectively; on (c) four leading edge activity sensors—AKT-PH-mRFP, CKAR, MARCKSp-mRFP, and leading edge area that report on PIP3 density, PKC activity, MARCKS membrane binding, and leading edge expansion/contraction, respectively. The results provide additional evidence that PKC and PI3K are both essential elements of the leading edge positive feedback loop, and strongly support the existence of a PKC-MARCKS-PI3K regulatory module linking the leading edge Ca2+ and PIP3 signals. As predicted, activators stimulate leading edge PKC activity, displacement of MARCKS from the leading edge membrane and increased leading edge PIP3 levels, while inhibitors

Leading-edge flow reattachment and the lateral static stability of low-aspect-ratio rectangular wings

NASA Astrophysics Data System (ADS)

Linehan, Thomas; Mohseni, Kamran

2017-11-01

The relationship between lateral static stability derivative, Clβ,lift coefficient, CL, and angle of attack was investigated for rectangular wings of aspect ratio A R =0.75 ,1 ,1.5 , and 3 using Stereo-Digital Particle Image Velocimetry (S-DPIV) and direct force and moment measurements. When the product Cl βA R is plotted with respect to CL, the lateral stability curves of each wing collapse to a single line for CL<0.7 . For CL>0.7 , the linearity and scaling of Clβwith respect to CL is lost. S-DPIV is used to elucidate the flow physics in this nonlinear regime. At α =10∘ , the leading-edge separation region emerges on the leeward portion of the sideslipped wing by means of vortex shedding. For the A R ≤1.5 wings at α >15∘ , the tip vortex downwash is sufficient to restrict the shedding of leading-edge vorticity thereby sustaining the lift of the leading-edge separation region at high angles of attack. Concurrently, the windward tip vortex grows in size and strength with increasing angle of attack, displacing the leading-edge separation region further toward the leeward wing. This reorganization of lift-generating vorticity results in the initial nonlinearities between Cl β and CL at angles of attack for which CL is still increasing. At angles of attack near that of maximum lift for the A R ≤1 wings, the windward tip vortex lifts off the wing, decreasing the lateral static stability of the wing prior to lift stall. For the A R =3 wing at α >10∘ , nonlinear trends in Cl β versus CL occur due to the spanwise evolution of stalled flow.

Textbook Multigrid Efficiency for Leading Edge Stagnation

NASA Technical Reports Server (NTRS)

Diskin, Boris; Thomas, James L.; Mineck, Raymond E.

2004-01-01

A multigrid solver is defined as having textbook multigrid efficiency (TME) if the solutions to the governing system of equations are attained in a computational work which is a small (less than 10) multiple of the operation count in evaluating the discrete residuals. TME in solving the incompressible inviscid fluid equations is demonstrated for leading-edge stagnation flows. The contributions of this paper include (1) a special formulation of the boundary conditions near stagnation allowing convergence of the Newton iterations on coarse grids, (2) the boundary relaxation technique to facilitate relaxation and residual restriction near the boundaries, (3) a modified relaxation scheme to prevent initial error amplification, and (4) new general analysis techniques for multigrid solvers. Convergence of algebraic errors below the level of discretization errors is attained by a full multigrid (FMG) solver with one full approximation scheme (FAS) cycle per grid. Asymptotic convergence rates of the FAS cycles for the full system of flow equations are very fast, approaching those for scalar elliptic equations.

An empirically-based model for the lift coefficients of twisted airfoils with leading-edge tubercles

NASA Astrophysics Data System (ADS)

Ni, Zao; Su, Tsung-chow; Dhanak, Manhar

2018-04-01

Experimental data for untwisted airfoils are utilized to propose a model for predicting the lift coefficients of twisted airfoils with leading-edge tubercles. The effectiveness of the empirical model is verified through comparison with results of a corresponding computational fluid-dynamic (CFD) study. The CFD study is carried out for both twisted and untwisted airfoils with tubercles, the latter shown to compare well with available experimental data. Lift coefficients of twisted airfoils predicted from the proposed empirically-based model match well with the corresponding coefficients determined using the verified CFD study. Flow details obtained from the latter provide better insight into the underlying mechanism and behavior at stall of twisted airfoils with leading edge tubercles.

Experimental investigation of wavy leading edges on rod-aerofoil interaction noise

NASA Astrophysics Data System (ADS)

Chen, Weijie; Qiao, Weiyang; Tong, Fan; Wang, Liangfeng; Wang, Xunnian

2018-05-01

Experimental studies are performed to investigate the effect of wavy leading edges on rod-aerofoil interaction noise in an open-jet anechoic wind tunnel. NACA 0012 aerofoils with straight and wavy leading edges (denoted by SLE and WLE, respectively) are embedded in the wake of a circular rod. The WLEs are in the form of sinusoidal profiles of amplitude, A, and wavelength, W. Parametric studies of the amplitude and wavelength characteristics are conducted to understand the effect of WLEs on noise reduction. It is observed that the sound power reduction level is sensitive to both the amplitude and wavelength of the WLEs. The WLE with the largest amplitude and smallest wavelength can achieve the most considerable noise reduction effect of up to 4 dB. The influences of rod diameter, d, and free-stream velocity, U0, on the noise reduction effect of the WLEs are also investigated. In addition, a parametric study of the influence of separating rod-aerofoil distance on the acoustic radiation of the SLE case and on the sound power reduction level of the WLE cases is performed. It is found that a critical spacing exists where the acoustic radiation and noise reduction can be divided into two different "modes".

Compressible Navier-Stokes equations: A study of leading edge effects

NASA Technical Reports Server (NTRS)

Hariharan, S. I.; Karbhari, P. R.

1987-01-01

A computational method is developed that allows numerical calculations of the time dependent compressible Navier-Stokes equations.The current results concern a study of flow past a semi-infinite flat plate.Flow develops from given inflow conditions upstream and passes over the flat plate to leave the computational domain without reflecting at the downstream boundary. Leading edge effects are included in this paper. In addition, specification of a heated region which gets convected with the flow is considered. The time history of this convection is obtained, and it exhibits a wave phenomena.

High-Energy Atmospheric Reentry Test Aerothermodynamic Analysis

NASA Technical Reports Server (NTRS)

Mazaheri, Alireza

2013-01-01

This paper presents an assessment of the aerothermodynamic environment around an 8.3 meter High Energy Atmospheric Reentry Test (HEART) vehicle. This study generated twelve nose shape configurations and compared their responses at the peak heating trajectory point against the baseline nose shape. The heat flux sensitivity to the angle of attack variations are also discussed. The possibility of a two-piece Thermal Protection System (TPS) design at the nose is also considered, as are the surface catalytic affects of the aeroheating environment of such configuration. Based on these analyses, an optimum nose shape is proposed to minimize the surface heating. A recommendation is also made for a two-piece TPS design, for which the surface catalytic uncertainty associated with the jump in heating at the nose-IAD juncture is reduced by a minimum of 93%. In this paper, the aeroshell is assumed to be rigid and the inflatable fluid interaction effect is left for future investigations

Direct Numerical Simulation of an Airfoil with Sand Grain Roughness on the Leading Edge

NASA Technical Reports Server (NTRS)

Ribeiro, Andre F. P.; Casalino, Damiano; Fares, Ehab; Choudhari, Meelan

2016-01-01

As part of a computational study of acoustic radiation due to the passage of turbulent boundary layer eddies over the trailing edge of an airfoil, the Lattice-Boltzmann method is used to perform direct numerical simulations of compressible, low Mach number flow past an NACA 0012 airfoil at zero degrees angle of attack. The chord Reynolds number of approximately 0.657 million models one of the test conditions from a previous experiment by Brooks, Pope, and Marcolini at NASA Langley Research Center. A unique feature of these simulations involves direct modeling of the sand grain roughness on the leading edge, which was used in the abovementioned experiment to trip the boundary layer to fully turbulent flow. This report documents the findings of preliminary, proof-of-concept simulations based on a narrow spanwise domain and a limited time interval. The inclusion of fully-resolved leading edge roughness in this simulation leads to significantly earlier transition than that in the absence of any roughness. The simulation data is used in conjunction with both the Ffowcs Williams-Hawkings acoustic analogy and a semi-analytical model by Roger and Moreau to predict the farfield noise. The encouraging agreement between the computed noise spectrum and that measured in the experiment indicates the potential payoff from a full-fledged numerical investigation based on the current approach. Analysis of the computed data is used to identify the required improvements to the preliminary simulations described herein.

Applications of the ram accelerator to hypervelocity aerothermodynamic testing

NASA Technical Reports Server (NTRS)

Bruckner, A. P.; Knowlen, C.; Hertzberg, A.

1992-01-01

A ram accelerator used as a hypervelocity launcher for large-scale aeroballistic range applications in hypersonics and aerodynamics research is presented. It is an in-bore ramjet device in which a projectile shaped like the centerbody of a supersonic ramjet is propelled down a stationary tube filled with a tailored combustible gas mixture. Ram accelerator operation has been demonstrated at 39 mm and 90 mm bores, supporting the proposition that this launcher concept can be scaled up to very large bore diameters of the order of 30-60 cm. It is concluded that high quality data obtained from the tube wall and projectile during the aceleration process itself are very useful for understanding aerothermodynamics of hypersonic flow in general, and for providing important CFD validation benchmarks.

Distributed-Memory Computing With the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA)

NASA Technical Reports Server (NTRS)

Riley, Christopher J.; Cheatwood, F. McNeil

1997-01-01

The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA), a Navier-Stokes solver, has been modified for use in a parallel, distributed-memory environment using the Message-Passing Interface (MPI) standard. A standard domain decomposition strategy is used in which the computational domain is divided into subdomains with each subdomain assigned to a processor. Performance is examined on dedicated parallel machines and a network of desktop workstations. The effect of domain decomposition and frequency of boundary updates on performance and convergence is also examined for several realistic configurations and conditions typical of large-scale computational fluid dynamic analysis.

The Leading Edge: A Career Development Workshop Series for Young Adults. Facilitator Guide.

ERIC Educational Resources Information Center

Canadian Career Development Foundation, Ottawa (Ontario).

This booklet is designed to be used by facilitators of the Canadian Career Development Foundation's "The Leading Edge: A Career Development Workshop Series for Young Adults." The guide provides information, including objectives of the workshops and lists of required materials, needed in order to facilitate an introductory session as well…

Dynamics Impact Tolerance of Shuttle RCC Leading Edge Panels Using LS-DYNA

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.; Lyle, Karen H.; Jones, Lisa E.; Hardy, Robin C.; Spellman, Regina L.; Carney, Kelly S.; Melis, Matthew E.; Stockwell, Alan E.

2005-01-01

This paper describes a research program conducted to enable accurate prediction of the impact tolerance of the shuttle Orbiter leading-edge wing panels using physics-based codes such as LS-DYNA, a nonlinear, explicit transient dynamic finite element code. The shuttle leading-edge panels are constructed of Reinforced-Carbon-Carbon (RCC) composite material, which is used because of its thermal properties to protect the shuttle during reentry into the Earth's atmosphere. Accurate predictions of impact damage from insulating foam and other debris strikes that occur during launch required materials characterization of expected debris, including strain-rate effects. First, analytical models of individual foam and RCC materials were validated. Next, analytical models of foam cylinders impacting 6- in. x 6-in. RCC flat plates were developed and validated. LS-DYNA pre-test models of the RCC flat plate specimens established the impact velocity of the test for three damage levels: no-detectable damage, non-destructive evaluation (NDE) detectable damage, or visible damage such as a through crack or hole. Finally, the threshold of impact damage for RCC on representative Orbiter wing panels was predicted for both a small through crack and for NDE-detectable damage.

Dynamic Impact Tolerance of Shuttle RCC Leading Edge Panels using LS-DYNA

NASA Technical Reports Server (NTRS)

Fasanella, Edwin; Jackson, Karen E.; Lyle, Karen H.; Jones, Lisa E.; Hardy, Robin C.; Spellman, Regina L.; Carney, Kelly S.; Melis, Matthew E.; Stockwell, Alan E.

2008-01-01

This paper describes a research program conducted to enable accurate prediction of the impact tolerance of the shuttle Orbiter leading-edge wing panels using 'physics-based- codes such as LS-DYNA, a nonlinear, explicit transient dynamic finite element code. The shuttle leading-edge panels are constructed of Reinforced-Carbon-Carbon (RCC) composite material, which issued because of its thermal properties to protect the shuttle during re-entry into the Earth's atmosphere. Accurate predictions of impact damage from insulating foam and other debris strikes that occur during launch required materials characterization of expected debris, including strain-rate effects. First, analytical models of individual foam and RCC materials were validated. Next, analytical models of individual foam cylinders impacting 6-in. x 6-in. RCC flat plates were developed and validated. LS-DYNA pre-test models of the RCC flat plate specimens established the impact velocity of the test for three damage levels: no-detectable damage, non-destructive evaluation (NDE) detectable damage, or visible damage such as a through crack or hole. Finally, the threshold of impact damage for RCC on representative Orbiter wing panels was predicted for both a small through crack and for NDE-detectable damage.

The effect of undulating leading-edge modifications on NACA 0021 airfoil characteristics

NASA Astrophysics Data System (ADS)

Rostamzadeh, N.; Kelso, R. M.; Dally, B. B.; Hansen, K. L.

2013-11-01

In spite of its mammoth physical size, the humpback whale's manoeuvrability in hunting has captured the attention of biologists as well as fluid mechanists. It has now been established that the protrusions on the leading-edges of the humpback's pectoral flippers, known as tubercles, account for this species' agility and manoeuvrability. In the present work, Prandtl's nonlinear lifting-line theory was employed to propose a hypothesis that the favourable traits observed in the performance of tubercled lifting bodies are not exclusive to this form of leading-edge configuration. Accordingly, a novel alternative to tubercles was introduced and incorporated into the design of four airfoils that underwent wind tunnel force and pressure measurement tests in the transitional flow regime. In addition, a Computation Fluid Dynamics study was performed using the Shear Stress Transport transitional model in the context of unsteady Reynolds-Averaged Navier-Stokes at several attack angles. The results from the numerical investigation are in reasonable agreement with those of the experiments, and suggest the presence of features that are also observed in flows over tubercled foils, most notably a distinct pair of streamwise vortices for each wavelength of the tubercle-like feature.

International Union of Theoretical and Applied Mechanics (IUTAM) Symposium on Aerothermodynamics in Combustors

NASA Astrophysics Data System (ADS)

Lee, Richard S.; Whitelaw, J. H.; Wung, T.-S.

1991-06-01

The subject of aerothermodynamics is playing an ever increasingly critical role in a variety of important industrial and technical problems in the design of combustors. In recent years, it has become the focus of attention among investigators from research laboratories and industries around the world resulting in a large number of meetings on its various aspects every year. However, most of these meetings deal with a certain problem area, for instance that of the global combustion of fuel droplets in a flow. Because of the inherent complexities involved in such flows, the analytical effort has been mostly confined to over-simplified and over-idealized flow systems while the experimental effort has been mostly directed towards global measurements of flows found in industrial applications. With the rapid and phenomenal developments of key research tools mostly in the last two decades, in particular those of modern digital computers, laser optics, and electronics, many of the previously unthinkable, rigorous investigations in real-life flows have gradually become feasible. It is against this background that this international conference on the aerothermodynamics in combustors is being held at this point in time. This symposium involves the presentation of papers concerned with flow and thermodynamic characteristics of combustors, with emphasis on gas-turbine combustors and including information relevant to rocket motors, internal combustion engines and furnaces.

An inlet analysis for the NASA hypersonic research engine aerothermodynamic integration model

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Russell, J. W.; Mackley, E. A.; Simmonds, A. L.

1974-01-01

A theoretical analysis for the inlet of the NASA Hypersonic Research Engine (HRE) Aerothermodynamic Integration Model (AIM) has been undertaken by use of a method-of-characteristics computer program. The purpose of the analysis was to obtain pretest information on the full-scale HRE inlet in support of the experimental AIM program (completed May 1974). Mass-flow-ratio and additive-drag-coefficient schedules were obtained that well defined the range effected in the AIM tests. Mass-weighted average inlet total-pressure recovery, kinetic energy efficiency, and throat Mach numbers were obtained.

Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration

NASA Technical Reports Server (NTRS)

Maki, Ralph L.

1959-01-01

Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

Aerothermodynamic optimization of Earth entry blunt body heat shields for Lunar and Mars return

NASA Astrophysics Data System (ADS)

Johnson, Joshua E.

A differential evolutionary algorithm has been executed to optimize the hypersonic aerodynamic and stagnation-point heat transfer performance of Earth entry heat shields for Lunar and Mars return manned missions with entry velocities of 11 and 12.5 km/s respectively. The aerothermodynamic performance of heat shield geometries with lift-to-drag ratios up to 1.0 is studied. Each considered heat shield geometry is composed of an axial profile tailored to fit a base cross section. Axial profiles consist of spherical segments, spherically blunted cones, and power laws. Heat shield cross sections include oblate and prolate ellipses, rounded-edge parallelograms, and blendings of the two. Aerothermodynamic models are based on modified Newtonian impact theory with semi-empirical correlations for convection and radiation. Multi-objective function optimization is performed to determine optimal trade-offs between performance parameters. Objective functions consist of minimizing heat load and heat flux and maximizing down range and cross range. Results indicate that skipping trajectories allow for vehicles with L/D = 0.3, 0.5, and 1.0 at lunar return flight conditions to produce maximum cross ranges of 950, 1500, and 3000 km respectively before Qs,tot increases dramatically. Maximum cross range increases by ˜20% with an increase in entry velocity from 11 to 12.5 km/s. Optimal configurations for all three lift-to-drag ratios produce down ranges up to approximately 26,000 km for both lunar and Mars return. Assuming a 10,000 kg mass and L/D = 0.27, the current Orion configuration is projected to experience a heat load of approximately 68 kJ/cm2 for Mars return flight conditions. For both L/D = 0.3 and 0.5, a 30% increase in entry vehicle mass from 10,000 kg produces a 20-30% increase in Qs,tot. For a given L/D, highly-eccentric heat shields do not produce greater cross range or down range. With a 5 g deceleration limit and L/D = 0.3, a highly oblate cross section with an

Structural characteristics of the shock-induced boundary layer separation extended to the leading edge

NASA Astrophysics Data System (ADS)

Tao, Y.; Liu, W. D.; Fan, X. Q.; Zhao, Y. L.

2017-07-01

For a better understanding of the local unstart of supersonic/hypersonic inlet, a series of experiments has been conducted to investigate the shock-induced boundary layer separation extended to the leading edge. Using the nanoparticle-based planar laser scattering, we recorded the fine structures of these interactions under different conditions and paid more attention to their structural characteristics. According to their features, these interactions could be divided into four types. Specifically, Type A wave pattern is similar to the classic shock wave/turbulent boundary layer interaction, and Type B wave configuration consists of an overall Mach reflection above the large scale separation bubble. Due to the gradual decrease in the size of the separation bubble, the separation bubble was replaced by several vortices (Type C wave pattern). Besides, for Type D wave configuration which exists in the local unstart inlet, there appears to be some flow spillage around the leading edge.

Planform curvature effects on flutter characteristics of a wing with 56 deg leading-edge sweep and panel aspect ratio of 1.14

NASA Technical Reports Server (NTRS)

Keller, Donald F.; Sandford, Maynard C.; Pinkerton, Theresa L.

1991-01-01

An experimental and analytical investigation was initiated to determine the effects of planform curvature (curving the leading and trailing edges of a wing in the X-Y plane) on the transonic flutter characteristics of a series of three moderately swept wing models. Experimental flutter results were obtained in the Langley Transonic Dynamics Tunnel for Mach numbers from 0.60-1.00, with air as the test medium. The models were semispan cantilevered wings with a 3 percent biconvex airfoil and a panel aspect ratio of 1.14. The baseline model had straight leading and trailing edges (i.e., no planform curvature). The radii of curvature of the leading edges for these two models were 200 and 80 inches. The radii of curvature of the leading edges of the other two models were determined so that the root and tip chords were identical for all three models. Experimental results showed that flutter-speed index and flutter frequency ratio increased as planform curvature increase (radius of curvature of the leading edge was decreased) over the test range of Mach numbers. Analytical flutter results were calculated with a subsonic flutter-prediction program, and they agreed well with the experimental results.

Subsonic Investigation of a Leading-Edge Boundary Layer Control Suction System on a High-Speed Civil Transport Configuration

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Applin, Zachary T.; Kemmerly, Guy T.; Coe, Paul L., Jr.; Owens, D. Bruce; Gile, Brenda E.; Parikh, Pradip G.; Smith, Don

1999-01-01

A wind tunnel investigation of a leading edge boundary layer control system was conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.08 to 0.27, with corresponding chord Reynolds numbers of 1.79 x 10(exp 6) to 5.76 x 10(exp 6). Variations in the amount of suction, as well as the size and location of the suction area, were tested with outboard leading edge flaps deflected 0 and 30 deg and trailing-edge flaps deflected 0 and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein.

Textbook Multigrid Efficiency for Leading Edge Stagnation

NASA Technical Reports Server (NTRS)

Diskin, Boris; Thomas, James L.; Mineck, Raymond E.

2004-01-01

A multigrid solver is defined as having textbook multigrid efficiency (TME) if the solutions to the governing system of equations are attained in a computational work which is a small (less than 10) multiple of the operation count in evaluating the discrete residuals. TME in solving the incompressible inviscid fluid equations is demonstrated for leading- edge stagnation flows. The contributions of this paper include (1) a special formulation of the boundary conditions near stagnation allowing convergence of the Newton iterations on coarse grids, (2) the boundary relaxation technique to facilitate relaxation and residual restriction near the boundaries, (3) a modified relaxation scheme to prevent initial error amplification, and (4) new general analysis techniques for multigrid solvers. Convergence of algebraic errors below the level of discretization errors is attained by a full multigrid (FMG) solver with one full approximation scheme (F.4S) cycle per grid. Asymptotic convergence rates of the F.4S cycles for the full system of flow equations are very fast, approaching those for scalar elliptic equations.

Blunt-Body Entry Vehicle Aerothermodynamics: Transition and Turbulence on the CEV and MSL Configurations

NASA Technical Reports Server (NTRS)

Hollis, Brian R.

2010-01-01

Recent, current, and planned NASA missions that employ blunt-body entry vehicles pose aerothermodynamic problems that challenge the state-of-the art of experimental and computational methods. The issues of boundary-layer transition and turbulent heating on the heat shield have become important in the designs of both the Mars Science Laboratory and Crew Exploration Vehicle. While considerable experience in these general areas exists, that experience is mainly derived from simple geometries; e.g. sharp-cones and flat-plates, or from lifting bodies such as the Space Shuttle Orbiter. For blunt-body vehicles, application of existing data, correlations, and comparisons is questionable because an all, or mostly, subsonic flow field is produced behind the bow shock, as compared to the supersonic (or even hypersonic) flow of other configurations. Because of the need for design and validation data for projects such as MSL and CEV, many new experimental studies have been conducted in the last decade to obtain detailed boundary-layer transition and turbulent heating data on this class of vehicle. In this paper, details of several of the test programs are reviewed. The laminar and turbulent data from these various test are shown to correlate in terms of edge-based Stanton and Reynolds number functions. Correlations are developed from the data for transition onset and turbulent heating augmentation as functions of momentum thickness Reynolds number. These correlation can be employed as engineering-level design and analysis tools.

Fabrication of lithium/C-103 alloy heat pipes for sharp leading edge cooling

NASA Astrophysics Data System (ADS)

Ai, Bangcheng; Chen, Siyuan; Yu, Jijun; Lu, Qin; Han, Hantao; Hu, Longfei

2018-05-01

In this study, lithium/C-103 alloys heat pipes are proposed for sharp leading edge cooling. Three models of lithium/C-103 alloy heat pipes were fabricated. And their startup properties were tested by radiant heat tests and aerothermal tests. It is found that the startup temperature of lithium heat pipe was about 860 °C. At 1000 °C radiant heat tests, the operating temperature of lithium/C-103 alloy heat pipe is lower than 860 °C. Thus, startup failure occurs. At 1100 °C radiant heat tests and aerothermal tests, the operating temperature of lithium/C-103 alloy heat pipe is higher than 860 °C, and the heat pipe starts up successfully. The startup of lithium/C-103 alloy heat pipe decreases the leading edge temperature effectively, which endows itself good ablation resistance. After radiant heat tests and aerothermal tests, all the heat pipe models are severely oxidized because of the C-103 poor oxidation resistance. Therefore, protective coatings are required for further applications of lithium/C-103 alloy heat pipes.

Two Functionally Distinct Sources of Actin Monomers Supply the Leading Edge of Lamellipodia

PubMed Central

Vitriol, Eric A.; McMillen, Laura M.; Kapustina, Maryna; Gomez, Shawn M.; Vavylonis, Dimitrios; Zheng, James Q.

2015-01-01

Summary Lamellipodia, the sheet-like protrusions of motile cells, consist of networks of actin filaments (F-actin) regulated by the ordered assembly from and disassembly into actin monomers (G-actin). Traditionally, G-actin is thought to exist as a homogeneous pool. Here, we show that there are two functionally and molecularly distinct sources of G-actin that supply lamellipodial actin networks. G-actin originating from the cytosolic pool requires the monomer binding protein thymosin β4 (Tβ4) for optimal leading edge localization, is targeted to formins, and is responsible for creating an elevated G/F-actin ratio that promotes membrane protrusion. The second source of G-actin comes from recycled lamellipodia F-actin. Recycling occurs independently of Tβ4 and appears to regulate lamellipodia homeostasis. Tβ4-bound G-actin specifically localizes to the leading edge because it doesn’t interact with Arp2/3-mediated polymerization sites found throughout the lamellipodia. These findings demonstrate that actin networks can be constructed from multiple sources of monomers with discrete spatiotemporal functions. PMID:25865895

Space Shuttle Orbiter Wing-Leading-Edge Panel Thermo-Mechanical Analysis for Entry Conditions

NASA Technical Reports Server (NTRS)

Knight, Norman F., Jr.; Song, Kyongchan; Raju, Ivatury S.

2010-01-01

Linear elastic, thermo-mechanical stress analyses of the Space Shuttle Orbiter wing-leading-edge panels is presented for entry heating conditions. The wing-leading-edge panels are made from reinforced carbon-carbon and serve as a part of the overall thermal protection system. Three-dimensional finite element models are described for three configurations: integrated configuration, an independent single-panel configuration, and a local lower-apex joggle segment. Entry temperature conditions are imposed and the through-the-thickness response is examined. From the integrated model, it was concluded that individual panels can be analyzed independently since minimal interaction between adjacent components occurred. From the independent single-panel model, it was concluded that increased through-the-thickness stress levels developed all along the chord of a panel s slip-side joggle region, and hence isolated local joggle sections will exhibit the same trend. From the local joggle models, it was concluded that two-dimensional plane-strain models can be used to study the influence of subsurface defects along the slip-side joggle region of these panels.

Analytical observations on the aerodynamics of a delta wing with leading edge flaps

NASA Technical Reports Server (NTRS)

Oh, S.; Tavella, D.

1986-01-01

The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.

Effects of wing leading-edge radius and Reynolds number on longitudinal aerodynamic characteristics of highly swept wing-body configurations at subsonic speeds

NASA Technical Reports Server (NTRS)

Henderson, W. P.

1976-01-01

An investigation was conducted in the Langley low turbulence pressure tunnel to determine the effects of wing leading edge radius and Reynolds number on the longitudinal aerodynamic characteristics of a series of highly swept wing-body configurations. The tests were conducted at Mach numbers below 0.30, angles of attack up to 16 deg, and Reynolds numbers per meter from 6.57 million to 43.27 million. The wings under study in this investigation had leading edge sweep angles of 61.7 deg, 64.61 deg, and 67.01 deg in combination with trailing edge sweep angles of 0 deg and 40.6 deg. The leading edge radii of each wing planform could be varied from sharp to nearly round.

Intermediate Experimental Vehicle, ESA Program IXV ATDB Tool and Aerothermodynamic Characterization

NASA Astrophysics Data System (ADS)

Mareschi, Vincenzo; Ferrarella, Daniela; Zaccagnino, Elio; Tribot, Jean-Pierre; Vallee, Jean-Jacques; Haya-Ramos, Rodrigo; Rufolo, Giuseppe; Mancuso, Salvatore

2011-05-01

In the complex domain of the space technologies and among the different applications available in Europe, a great interest has been placed since several years in the development of re-entry technologies. Among the different achievements obtained in that field it is to be recalled the experience of the Atmospheric Re-entry Vehicle flight in 1998 and a certain number of important investments per-formed at Agency and national levels like Hermes, MSTP, Festip, X-38, FLPP, TRP, GSTP, HSTS, AREV, Pre-X. IXV (Intermediate eXperimental V ehicle) builds on these past experiences and studies and it is conceived to be the next technological step forward with respect to ARD With respect to previous European ballistic or quasi- ballistic demonstrators, IXV will have an increased in- flight manoeuvrability and the planned mission will allow verifying the performances of the required technologies against a wider re-entry corridor. This will imply from the pure technological aspect to increase the level of engagement on critical technologies and disciplines like aerodynamics/aerothermodynamics, guidance, navigation, control, thermal protection materials and in flight measurements. In order to support the TPS design and the other sub- systems, an AeroThermodynamicDataBase Tool has been developed by Dassault Aviation and integrated by Thales Alenia Space with the Functional Engineering Simulator (used for GNC performances evaluation) in order to characterize the aerothermodynamic behaviour of the vehicle. This paper will describe: - The methodology used to develop the ATDB tool, based on the processing of CFD computations and WTT campaigns results. - The utilization of the ATDB tool, by means of its integration into the System process. - The methodology used for the aerothermal characterization of IXV.

The Leading Edge: A Career Development Workshop Series for Young Adults. Participant Workbook.

ERIC Educational Resources Information Center

Canadian Career Development Foundation, Ottawa (Ontario).

This booklet is designed for participants in "The Leading Edge: A Career Development Workshop Series for Young Adults." It provides the 27 participant handouts for the six workshops in the series. The first in the series, "Setting the Stage: The Changing World of Work," is a workshop to clarify what is occurring in the world of…

Unsteady behavior of leading-edge vortex and diffuser stall in a centrifugal compressor with vaned diffuser

NASA Astrophysics Data System (ADS)

Fujisawa, Nobumichi; Hara, Shotaro; Ohta, Yutaka

2016-02-01

The characteristics of a rotating stall of an impeller and diffuser and the evolution of a vortex generated at the diffuser leading-edge (i.e., the leading-edge vortex (LEV)) in a centrifugal compressor were investigated by experiments and numerical analysis. The results of the experiments revealed that both the impeller and diffuser rotating stalls occurred at 55 and 25 Hz during off-design flow operation. For both, stall cells existed only on the shroud side of the flow passages, which is very close to the source location of the LEV. According to the CFD results, the LEV is made up of multiple vortices. The LEV is a combination of a separated vortex near the leading- edge and a longitudinal vortex generated by the extended tip-leakage flow from the impeller. Therefore, the LEV is generated by the accumulation of vorticity caused by the velocity gradient of the impeller discharge flow. In partial-flow operation, the spanwise extent and the position of the LEV origin are temporarily transmuted. The LEV develops with a drop in the velocity in the diffuser passage and forms a significant blockage within the diffuser passage. Therefore, the LEV may be regarded as being one of the causes of a diffuser stall in a centrifugal compressor.

Thermal/structural analyses of several hydrogen-cooled leading-edge concepts for hypersonic flight vehicles

NASA Technical Reports Server (NTRS)

Gladden, Herbert J.; Melis, Matthew E.; Mockler, Theodore T.; Tong, Mike

1990-01-01

The aerodynamic heating at high flight Mach numbers, when shock interference heating is included, can be extremely high and can exceed the capability of most conventional metallic and potential ceramic materials available. Numerical analyses of the heat transfer and thermal stresses are performed on three actively cooled leading-edge geometries (models) made of three different materials to address the issue of survivability in a hostile environment. These analyses show a mixture of results from one configuration to the next. Results for each configuration are presented and discussed. Combinations of enhanced internal film coefficients and high material thermal conductivity of copper and tungsten are predicted to maintain the maximum wall temperature for each concept within acceptable operating limits. The exception is the TD nickel material which is predicted to melt for most cases. The wide range of internal impingement film coefficients (based on correlations) for these conditions can lead to a significant uncertainty in expected leading-edge wall temperatures. The equivalent plastic strain, inherent in each configuration which results from the high thermal gradients, indicates a need for further cyclic analysis to determine component life.

Surface-Pressure and Flow-Visualization Data at Mach Number of 1.60 for Three 65 deg Delta Wings Varying in Leading-Edge Radius and Camber

NASA Technical Reports Server (NTRS)

McMillin, S. Naomi; Bryd, James E.; Parmar, Devendra S.; Bezos-OConnor, Gaudy M.; Forrest, Dana K.; Bowen, Susan

1996-01-01

An experimental investigation of the effect of leading-edge radius, camber, Reynolds number, and boundary-layer state on the incipient separation of a delta wing at supersonic speeds was conducted at the Langley Unitary Plan Wind Tunnel at Mach number of 1.60 over a free-stream Reynolds number range of 1 x 106 to 5 x 106 ft-1. The three delta wing models examined had a 65 deg swept leading edge and varied in cross-sectional shape: a sharp wedge, a 20:1 ellipse, and a 20:1 ellipse with a -9.750 circular camber imposed across the span. The wings were tested with and without transition grit applied. Surface-pressure coefficient data and flow-visualization data indicated that by rounding the wing leading edge or cambering the wing in the spanwise direction, the onset of leading-edge separation on a delta wing can be raised to a higher angle of attack than that observed on a sharp-edged delta wing. The data also showed that the onset of leading-edge separation can be raised to a higher angle of attack by forcing boundary-layer transition to occur closer to the wing leading edge by the application of grit or the increase in free-stream Reynolds number.

DiMES Tests of W Leading Edge Power Loading in DIII-D

NASA Astrophysics Data System (ADS)

Nygren, R. E.; Watkins, J. G.; Rudakov, D. L.; Lasnier, C. J.; Pitts, R. A.; Stangeby, P. C.

2015-11-01

In a transient melt experiment in JET, the power to a ~1-mm-high leading edge on a W lamella in the bulk-W outer divertor was lower than expected from the geometry by factors of 5 and 2 for L-mode and H-mode discharges, respectively. We checked this surprising result in DIII-D with 3 W blocks (10 mm square) mounted radially side-by-side in DiMES with leading edges of 0.0, 0.3, 1.0 mm, single null L-mode plasmas, OSP just outside ``0.0'' block, limited scans (NBI+ECH), B-field incident at 1.5° or 2.5°, and viewed, as in JET, from above with 0.2mm/pixel resolution IRTV. Langmuir probes measured parallel power to the target. We compared probe and IR data with a detailed thermal model of the blocks and concluded provisionally that we did not reproduce the power deficit found in JET. Blurred IR images complicated fitting of temperature distributions from the thermal model. We plan an experiment with both L- and H-mode He plasmas before the APS meeting. Supported by US DOE under DE-AC04-94AL85000, 44500007360, DE-AC52-07NA27344, and DE-FC02-04ER54698.

Hypersonic research engine/aerothermodynamic integration model, experimental results. Volume 1: Mach 6 component integration

NASA Technical Reports Server (NTRS)

Andrews, E. H., Jr.; Mackley, E. A.

1976-01-01

The NASA Hypersonic Research Engine (HRE) Project was initiated for the purpose of advancing the technology of airbreathing propulsion for hypersonic flight. A large component (inlet, combustor, and nozzle) and structures development program was encompassed by the project. The tests of a full-scale (18 in. diameter cowl and 87 in. long) HRE concept, designated the Aerothermodynamic Integration Model (AIM), at Mach numbers of 5, 6, and 7. Computer program results for Mach 6 component integration tests are presented.

User's manual for interfacing a leading edge, vortex rollup program with two linear panel methods

NASA Technical Reports Server (NTRS)

Desilva, B. M. E.; Medan, R. T.

1979-01-01

Sufficient instructions are provided for interfacing the Mangler-Smith, leading edge vortex rollup program with a vortex lattice (POTFAN) method and an advanced higher order, singularity linear analysis for computing the vortex effects for simple canard wing combinations.

Materials, Manufacturing and Test Development of a Composite Fan Blade Leading Edge Subcomponent for Improved Impact Resistance

NASA Technical Reports Server (NTRS)

Handschuh, Katherine M.; Miller, Sandi G.; Sinnott, Matthew J.; Kohlman, Lee W.; Roberts, Gary D.; Pereira, J. Michael; Ruggeri, Charles R.

2014-01-01

Application of polymer matrix composite materials for jet engine fan blades is becoming attractive as an alternative to metallic blades; particularly for large engines where significant weight savings are recognized on moving to a composite structure. However, the weight benefit of the composite of is offset by a reduction of aerodynamic efficiency resulting from a necessary increase in blade thickness; relative to the titanium blades. Blade dimensions are largely driven by resistance to damage on bird strike. Further development of the composite material is necessary to allow composite blade designs to approximate the dimensions of a metallic fan blade. The reduction in thickness over the state of the art composite blades is expected to translate into structural weight reduction, improved aerodynamic efficiency, and therefore reduced fuel consumption. This paper presents test article design, subcomponent blade leading edge fabrication, test method development, and initial results from ballistic impact of a gelatin projectile on the leading edge of composite fan blades. The simplified test article geometry was developed to realistically simulate a blade leading edge while decreasing fabrication complexity. Impact data is presented on baseline composite blades and toughened blades; where a considerable improvement to impact resistance was recorded.

Materials, Manufacturing, and Test Development of a Composite Fan Blade Leading Edge Subcomponent for Improved Impact Resistance

NASA Technical Reports Server (NTRS)

Miller, Sandi G.; Handschuh, Katherine; Sinnott, Matthew J.; Kohlman, Lee W.; Roberts, Gary D.; Martin, Richard E.; Ruggeri, Charles R.; Pereira, J. Michael

2015-01-01

Application of polymer matrix composite materials for jet engine fan blades is becoming attractive as an alternative to metallic blades; particularly for large engines where significant weight savings are recognized on moving to a composite structure. However, the weight benefit of the composite is offset by a reduction of aerodynamic efficiency resulting from a necessary increase in blade thickness; relative to the titanium blades. Blade dimensions are largely driven by resistance to damage on bird strike. Further development of the composite material is necessary to allow composite blade designs to approximate the dimensions of a metallic fan blade. The reduction in thickness over the state of the art composite blades is expected to translate into structural weight reduction, improved aerodynamic efficiency, and therefore reduced fuel consumption. This paper presents test article design, subcomponent blade leading edge fabrication, test method development, and initial results from ballistic impact of a gelatin projectile on the leading edge of composite fan blades. The simplified test article geometry was developed to realistically simulate a blade leading edge while decreasing fabrication complexity. Impact data is presented on baseline composite blades and toughened blades; where a considerable improvement to impact resistance was recorded.

Model for adhesion clutch explains biphasic relationship between actin flow and traction at the cell leading edge

PubMed Central

Craig, Erin M.; Stricker, Jonathan; Gardel, Margaret L.; Mogilner, Alex

2015-01-01

Cell motility relies on the continuous reorganization of a dynamic actin-myosin-adhesion network at the leading edge of the cell, in order to generate protrusion at the leading edge and traction between the cell and its external environment. We analyze experimentally measured spatial distributions of actin flow, traction force, myosin density, and adhesion density in control and pharmacologically perturbed epithelial cells in order to develop a mechanical model of the actin-adhesion-myosin self-organization at the leading edge. A model in which the F-actin network is treated as a viscous gel, and adhesion clutch engagement is strengthened by myosin but weakened by actin flow, can explain the measured molecular distributions and correctly predict the spatial distributions of the actin flow and traction stress. We test the model by comparing its predictions with measurements of the actin flow and traction stress in cells with fast and slow actin polymerization rates. The model predicts how the location of the lamellipodium-lamellum boundary depends on the actin viscosity and adhesion strength. The model further predicts that the location of the lamellipodium-lamellum boundary is not very sensitive to the level of myosin contraction. PMID:25969948

Wind-tunnel studies of advanced cargo aircraft concepts. [leading edge vortex flaps for drag reduction

NASA Technical Reports Server (NTRS)

Rao, D. M.; Goglia, G. L.

1981-01-01

Accomplishments in vortex flap research are summarized. A singular feature of the vortex flap is that, throughout the range of angle of attack range, the flow type remains qualitatively unchanged. Accordingly, no large or sudden change in the aerodynamic characteristics, as happens when forcibly maintained attached flow suddenly reverts to separation, will occur with the vortex flap. Typical wind tunnel test data are presented which show the drag reduction potential of the vortex flap concept applied to a supersonic cruise airplane configuration. The new technology offers a means of aerodynamically augmenting roll-control effectiveness on slender wings at higher angles of attack by manipulating the vortex flow generated from leading edge separation. The proposed manipulator takes the form of a flap hinged at or close to the leading edge, normally retracted flush with the wing upper surface to conform to the airfoil shape.

Estimating Blade Section Airloads from Blade Leading-Edge Pressure Measurements

NASA Technical Reports Server (NTRS)

vanAken, Johannes M.

2003-01-01

The Tilt-Rotor Aeroacoustic Model (TRAM) test in the Duitse-Nederlandse Wind (DNW) Tunnel acquired blade pressure data for forward flight test conditions of a tiltrotor in helicopter mode. Chordwise pressure data at seven radial locations were integrated to obtain the blade section normal force. The present investigation evaluates the use of linear regression analysis and of neural networks in estimating the blade section normal force coefficient from a limited number of blade leading-edge pressure measurements and representative operating conditions. These network models are subsequently used to estimate the airloads at intermediate radial locations where only blade pressure measurements at the 3.5% chordwise stations are available.

Leading-edge vortex research: Some nonplanar concepts and current challenges

NASA Technical Reports Server (NTRS)

Campbell, J. F.; Osborn, R. F.

1986-01-01

Some background information is provided for the Vortex Flow Aerodynamics Conference and that current slender wing airplanes do not use variable leading edge geometry to improve transonic drag polar is shown. Highlights of some of the initial studies combining wing camber, or flaps, with vortex flow are presented. Current vortex flap studies were reviewed to show that there is a large subsonic data base and that transonic and supersonic generic studies have begun. There is a need for validated flow field solvers to calculate vortex/shock interactions at transonic and supersonic speeds. Many important research opportunities exist for fundamental vortex flow investigations and for designing advanced fighter concepts.

Effect of leading edge sweep on the performance of cavitating inducer of LOX booster turbopump used in semicryogenic engine

NASA Astrophysics Data System (ADS)

Mishra, Arpit; Ghosh, Parthasarathi

2017-02-01

As a part of the developmental effort towards the realization of a staged combustion cycle based liquid rocket engine, a program on simulation of the LOX booster pump for performance characterization has been taken up. Earlier reported work shows that the pump inducer works satisfactorily under cavitating conditions for the throttling range varying from 90% to 113%. However stall occurs below 90% of the designed flow rate which is to be strongly associated with the inlet backflow vortices due to flow separation [1]. It is envisaged that leading edge sweep may help in to controls the incipience and growth of the backflow vortices at the inlet leading edge tip of axial flow inducer leading to a wider operating range. In this paper, steady state 3D CFD analysis of rotating inducer is performed to examine the effect of leading edge sweep on the performance of axial flow LOX pump inducer using ANSYS® CFX and has been compared with the performance of the inducer reported by Mishra and Ghosh [1].

Dual modes of motility at the leading edge of migrating epithelial cell sheets

PubMed Central

Klarlund, Jes K.

2012-01-01

Purse-string healing is driven by contraction of actin/myosin cables that span cells at wound edges, and it is the predominant mode of closing small round wounds in embryonic and some adult epithelia. Wounds can also heal by cell crawling, and my colleagues and I have shown previously that the presence of unconstrained, straight edges in sheets of epithelial cells is a sufficient signal to induce healing by crawling. Here, it is reported that the presence of highly concave edges, which are free or physically constrained by an inert material (agarose), is sufficient to induce formation of purse strings. It was determined that neither of the two types of healing required cell damage or other potential stimuli by using the particularly gentle procedure of introducing gaps by digesting agarose blocks imbedded in the cell sheets. Movement by crawling depends on signaling by the EGF receptor (EGFR); however, this was not required for purse-string contraction. A migrating epithelial cell sheet usually produces finger-like projections of crawling cells. The cells between fingers contain continuous actin cables, which were also determined to contain myosin IIA and exhibit additional characteristics of purse strings. When crawling was blocked by inhibition of EGFR signaling, the concave regions continued to move, suggesting that both mechanisms contribute to propel the sheets forward. Wounding epithelial cell sheets causes activation of the EGFR, which triggers movement by crawling. The EGFR was found to be activated only at straight and convex edges, which explains how both types of movement can coexist at leading epithelial edges. PMID:23019364

Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

NASA Technical Reports Server (NTRS)

Arnold, James O.; Deiwert, George S.

1997-01-01

This paper surveys the use of aerothermodynamic facilities which have been useful in the study of external flows and propulsion aspects of hypersonic, air-breathing vehicles. While the paper is not a survey of all facilities, it covers the utility of shock tunnels and conventional hypersonic blow-down facilities which have been used for hypersonic air-breather studies. The problems confronting researchers in the field of aerothermodynamics are outlined. Results from the T5 GALCIT tunnel for the shock-on lip problem are outlined. Experiments on combustors and short expansion nozzles using the semi-free jet method have been conducted in large shock tunnels. An example which employed the NASA Ames 16-Inch shock tunnel is outlined, and the philosophy of the test technique is described. Conventional blow-down hypersonic wind tunnels are quite useful in hypersonic air-breathing studies. Results from an expansion ramp experiment, simulating the nozzle on a hypersonic air-breather from the NASA Ames 3.5 Foot Hypersonic wind tunnel are summarized. Similar work on expansion nozzles conducted in the NASA Langley hypersonic wind tunnel complex is cited. Free-jet air-frame propulsion integration and configuration stability experiments conducted at Langley in the hypersonic wind tunnel complex on a small generic model are also summarized.

Team Software Development for Aerothermodynamic and Aerodynamic Analysis and Design

NASA Technical Reports Server (NTRS)

Alexandrov, N.; Atkins, H. L.; Bibb, K. L.; Biedron, R. T.; Carpenter, M. H.; Gnoffo, P. A.; Hammond, D. P.; Jones, W. T.; Kleb, W. L.; Lee-Rausch, E. M.

2003-01-01

A collaborative approach to software development is described. The approach employs the agile development techniques: project retrospectives, Scrum status meetings, and elements of Extreme Programming to efficiently develop a cohesive and extensible software suite. The software product under development is a fluid dynamics simulator for performing aerodynamic and aerothermodynamic analysis and design. The functionality of the software product is achieved both through the merging, with substantial rewrite, of separate legacy codes and the authorship of new routines. Examples of rapid implementation of new functionality demonstrate the benefits obtained with this agile software development process. The appendix contains a discussion of coding issues encountered while porting legacy Fortran 77 code to Fortran 95, software design principles, and a Fortran 95 coding standard.

Impact of clocking on the aero-thermodynamics of a second stator tested in a one and a half stage HP turbine

NASA Astrophysics Data System (ADS)

Billiard, N.; Paniagua, Guillermo; Dénos, R.

2008-06-01

This paper focuses on the experimental investigation of the time-averaged and time-accurate aero-thermodynamics of a second stator tested in a 1.5 stage high-pressure turbine. The effect of clocking on aerodynamic and heat transfer are investigated. Tests are performed under engine representative conditions in the VKI compression tube CT3. The test program includes four different clocking positions, i.e. relative pitch-wise positions between the first and the second stator. Probes located upstream and downstream of the second stator provide the thermodynamic conditions of the flow field. On the second stator airfoil, measurements are taken around the blade profile at 15, 50 and 85% span with pressure sensors and thin-film gauges. Both time-averaged and time-resolved aspects of the flow field are addressed. Regarding the time-averaged results, clocking effects are mainly observed within the leading edge region of the second stator, the largest effects being observed at 15% span. The surface static pressure distribution is changed locally, hence affecting the overall airfoil performance. For one clocking position, the thermal load of the airfoil is noticeably reduced. Pressure fluctuations are attributed to the passage of the upstream transonic rotor and its associated pressure gradients. The pattern of these fluctuations changes noticeably as a function of clocking. The time-resolved variations of heat flux and static pressure are analyzed together showing that the major effect is due to a potential interaction. The time-resolved pressure distribution integrated along the second stator surface yields the unsteady forces on the vane. The magnitude of the unsteady force is very dependent on the clocking position.

Characterization of Unsteady Flow Structures Near Leading-Edge Slat. Part 1; PIV Measurements

NASA Technical Reports Server (NTRS)

Jenkins, Luther N.; Khorrami, Mehdi R.; Choudhari, Meelan

2004-01-01

A comprehensive computational and experimental study has been performed at the NASA Langley Research Center as part of the Quiet Aircraft Technology (QAT) Program to investigate the unsteady flow near a leading-edge slat of a two-dimensional, high-lift system. This paper focuses on the experimental effort conducted in the NASA Langley Basic Aerodynamics Research Tunnel (BART) where Particle Image Velocimetry (PIV) data was acquired in the slat cove and at the slat trailing edge of a three-element, high-lift model at 4, 6, and 8 degrees angle of attack and a freestream Mach Number of 0.17. Instantaneous velocities obtained from PIV images are used to obtain mean and fluctuating components of velocity and vorticity. The data show the recirculation in the cove, reattachment of the shear layer on the slat lower surface, and discrete vortical structures within the shear layer emanating from the slat cusp and slat trailing edge. Detailed measurements are used to examine the shear layer formation at the slat cusp, vortex shedding at the slat trailing edge, and convection of vortical structures through the slat gap. Selected results are discussed and compared with unsteady, Reynolds-Averaged Navier-Stokes (URANS) computations for the same configuration in a companion paper by Khorrami, Choudhari, and Jenkins (2004). The experimental dataset provides essential flow-field information for the validation of near-field inputs to noise prediction tools.

Photographer; NACA North American F-100A NASA-200 Super Sabre airplane - wing leading edge deflected

NASA Technical Reports Server (NTRS)

1958-01-01

Photographer; NACA North American F-100A NASA-200 Super Sabre airplane - wing leading edge deflected 60 degrees for increased lift with boundary=layer control; takeoff preformance was improved 10% (mar 1960)

The Influence of Mesh Density on the Impact Response of a Shuttle Leading-Edge Panel Finite Element Simulation

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.; Lyle, Karen H.; Spellman, Regina L.

2004-01-01

A study was performed to examine the influence of varying mesh density on an LS-DYNA simulation of a rectangular-shaped foam projectile impacting the space shuttle leading edge Panel 6. The shuttle leading-edge panels are fabricated of reinforced carbon-carbon (RCC) material. During the study, nine cases were executed with all possible combinations of coarse, baseline, and fine meshes of the foam and panel. For each simulation, the same material properties and impact conditions were specified and only the mesh density was varied. In the baseline model, the shell elements representing the RCC panel are approximately 0.2-in. on edge, whereas the foam elements are about 0.5-in. on edge. The element nominal edge-length for the baseline panel was halved to create a fine panel (0.1-in. edge length) mesh and doubled to create a coarse panel (0.4-in. edge length) mesh. In addition, the element nominal edge-length of the baseline foam projectile was halved (0.25-in. edge length) to create a fine foam mesh and doubled (1.0- in. edge length) to create a coarse foam mesh. The initial impact velocity of the foam was 775 ft/s. The simulations were executed in LS-DYNA version 960 for 6 ms of simulation time. Contour plots of resultant panel displacement and effective stress in the foam were compared at five discrete time intervals. Also, time-history responses of internal and kinetic energy of the panel, kinetic and hourglass energy of the foam, and resultant contact force were plotted to determine the influence of mesh density. As a final comparison, the model with a fine panel and fine foam mesh was executed with slightly different material properties for the RCC. For this model, the average degraded properties of the RCC were replaced with the maximum degraded properties. Similar comparisons of panel and foam responses were made for the average and maximum degraded models.

Integrated Design Engineering Analysis (IDEA) Environment - Aerodynamics, Aerothermodynamics, and Thermal Protection System Integration Module

NASA Technical Reports Server (NTRS)

Kamhawi, Hilmi N.

2011-01-01

This report documents the work performed during from March 2010 October 2011. The Integrated Design and Engineering Analysis (IDEA) environment is a collaborative environment based on an object-oriented, multidisciplinary, distributed environment using the Adaptive Modeling Language (AML) as the underlying framework. This report will focus on describing the work done in the area of extending the aerodynamics, and aerothermodynamics module using S/HABP, CBAERO, PREMIN and LANMIN. It will also detail the work done integrating EXITS as the TPS sizing tool.

AN EXPERIMENTAL EVALUATION OF THREE METHODS FOR INCREASING THE LEADING EDGE THICKNESS OF SUPERCAVITATING PROPELLERS.

DTIC Science & Technology

An experimental investigation of three methods of increasing the leading-edge thickness of supercavitating propeller sections was conducted. The...addition of a parabolic thickness to both the face and back of the section. Five supercavitating propeller models incorporating these three methods of

Method and System for Weakening Shock Wave Strength at Leading Edge Surfaces of Vehicle in Supersonic Atmospheric Flight

NASA Technical Reports Server (NTRS)

Pritchett, Victor E., II (Inventor); Wang, Ten-See (Inventor); Blankson, Isaiah M. (Inventor); Daso, Endwell O. (Inventor); Farr, Rebecca Ann (Inventor); Auslender, Aaron Howard (Inventor); Plotkin, Kenneth J. (Inventor)

2015-01-01

A method and system are provided to weaken shock wave strength at leading edge surfaces of a vehicle in atmospheric flight. One or more flight-related attribute sensed along a vehicle's outer mold line are used to control the injection of a non-heated, non-plasma-producing gas into a local external flowfield of the vehicle from at least one leading-edge surface location along the vehicle's outer mold line. Pressure and/or mass flow rate of the gas so-injected is adjusted in order to cause a Rankine-Hugoniot Jump Condition along the vehicle's outer mold line to be violated.

A leading edge heating array and a flat surface heating array: Final design. [for testing the thermal protection system of the space shuttle

NASA Technical Reports Server (NTRS)

1975-01-01

A heating array is described for testing full-scale sections of the leading edge and lower fuselage surfaces of the shuttle. The heating array was designed to provide a tool for development and acceptance testing of leading edge segments and large flat sections of the main body thermal protection system. The array was designed using a variable length module concept to meet test requirements using interchangeable components from one test configuration in another configuration. Heat generating modules and heat absorbing modules were employed to achieve the thermal gradient around the leading edge. A support was developed to hold the modules to form an envelope around a variety of leading edges; to supply coolant to each module; the support structure and to hold the modules in the flat surface heater configuration. An optical pyrometer system mounted within the array was designed to monitor specimen surface temperatures without altering the test article's surface.

Equations and charts for the rapid estimation of hinge-moment and effectiveness parameters for trailing-edge controls having leading and trailing edges swept ahead of the Mach lines

NASA Technical Reports Server (NTRS)

Goin, Kennith L

1951-01-01

Existing conical-flow solutions have been used to calculate the hinge-moments and effectiveness parameters of trailing-edge controls having leading and trailing edges swept ahead of the Mach lines and having streamwise root and tip chords. Equations and detailed charts are presented for the rapid estimation of these parameters. Also included is an approximate method by which these parameters may be corrected for airfoil-section thickness.

Influence of airfoil geometry on delta wing leading-edge vortices and vortex-induced aerodynamics at supersonic speeds

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Byrd, James E.; Wesselmann, Gary F.

1992-01-01

An assessment of the influence of airfoil geometry on delta wing leading edge vortex flow and vortex induced aerodynamics at supersonic speeds is discussed. A series of delta wing wind tunnel models were tested over a Mach number range from 1.7 to 2.0. The model geometric variables included leading edge sweep and airfoil shape. Surface pressure data, vapor screen, and oil flow photograph data were taken to evaluate the complex structure of the vortices and shocks on the family of wings tested. The data show that airfoil shape has a significant impact on the wing upper surface flow structure and pressure distribution, but has a minimal impact on the integrated upper surface pressure increments.

Hypersonic research engine project. Phase 2: Aerothermodynamic integration model development, data item no. 55-4-21

NASA Technical Reports Server (NTRS)

Jilly, L. F. (Editor)

1975-01-01

The design and development of the Aerothermodynamic Integration Model (AIM) of the Hypersonic Research Engine (HRE) is described. The feasibility of integrating the various analytical and experimental data available for the design of the hypersonic ramjet engine was verified and the operational characteristic and the overall performance of the selected design was determined. The HRE-AIM was designed for operation at speeds of Mach 3 through Mach 8.

Project 2000-3 Leading Edge Enterprise: Insights into Employment and Training Practices. Working Paper.

ERIC Educational Resources Information Center

Long, Michael; Fischer, John

Leading-edge firms (LEFs)--at the forefront of their industry in terms of growth or market share--may influence skill development through diffusion of technology, products, or practices and use of market power to set standards or change customer businesses. Study of LEFs can identify the type and mix of skills needed in the industry. LEFs are…

Wind-tunnel tests of a Clark Y wing with 'Maxwell' leading-edge slots

NASA Technical Reports Server (NTRS)

Gauvain, William E

1937-01-01

Aerodynamic force tests of a Clark Y wing equipped with "Maxwell" type leading-edge slots were conducted in the N.A.C.A. 7- by 10-foot tunnel to ascertain the aerodynamic characteristics, which involved the determination of the best slot-gap opening, the effects of slat width, and the effect of a trailing-edge flap. The Maxwell wing with a wide-chord slat (0.30 c(sub w)) and with a 0.211 c(sub w) split flap deflected 60 degrees had a C(sub L sub max) of 2.53 or about twice that of the plain wing. The wing with the wide slat also had, in general, improved aerodynamic characteristics over those of the Maxwell wing with slat, and had about the same aerodynamic characteristics as a Handley Page slotted wing with approximately the same size of slat.

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils, an experimental and analytical investigation

NASA Technical Reports Server (NTRS)

Schwind, R. G.; Allen, H. J.

1973-01-01

High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 times one million to 6.2 times one million on a rectangular wing of NACA 63-009 airfoil section. Measurements were also obtained with a wide selection of leading-edge serrations added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large apatial peak in rms pressure. Frequency analysis of the pressure signals in this region show a large, high-frequency energy peak which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related to the trailing-edge and boundary-layer thicknesses.

An overview of the fundamental aerodynamics branch's research activities in wing leading-edge vortex flows at supersonic speeds

NASA Technical Reports Server (NTRS)

Miller, D. S.; Wood, R. M.; Covell, P. F.

1986-01-01

For the past 3 years, a research program pertaining to the study of wing leading edge vortices at supersonic speeds has been conducted in the Fundamental Aerodynamics Branch of the High-Speed Aerodynamics Division at the Langley Research Center. The purpose of the research is to provide an understanding of the factors governing the formation and the control of wing leading-edge vortices and to evaluate the use of these vortices for improving supersonic aerodynamic performance. The studies include both experimental and theoretical investigations and focus primarily on planform, thickness and camber effects for delta wings. An overview of this research activity is presented.

Quasi-chemostat behavior in the leading edge of B. subtilis biofilms

NASA Astrophysics Data System (ADS)

Srinivasan, Siddarth; Mahadevan, Lakshminarayanan; Rubinstein, Shmuel

2015-11-01

Bacillus subtilis is a gram positive bacterium that is a model system commonly used to study biofilm formation. By performing wide-field time-lapse microscopy on a fluorescently labeled B. subtilis strain, we observe a well defined steady boundary layer at the edge of a biofilm growing on an nutrient infused agar gel substrate, within which the outward radial expansion growth predominantly occurs. Using distinct fluorescent protein markers as proxies of gene expression, we quantitatively measure how the width, velocity and ratio of motile cell to matrix cell phenotypes within this boundary layer responds to changes in environmental conditions (such as substrate agar percentage & temperature). We further propose that the steady state at the leading edge can be interpreted as a quasi-chemostat which may enable well controlled response experiments on a colony scale. Finally, we show that for low agar concentration (0.5 wt%), the cells exhibit swarming behavior, whose dynamics and swimming velocities are characterized using differential dynamic microscopy. We show the swarming state is associated with an unstable front which gives rise to fingering and branching growth patterns, illustrating the varied morphological response of the biofilm to environmental conditions

The Leading Edge of Early Childhood Education: Linking Science to Policy for a New Generation

ERIC Educational Resources Information Center

Lesaux, Nonie K., Ed.; Jones, Stephanie M., Ed.

2016-01-01

"The Leading Edge of Early Childhood Education" aims to support the effort to simultaneously scale up and improve the quality of early childhood education by bringing together relevant insights from emerging research to provide guidance for this critical, fledgling field. It reflects the growing recognition that early childhood…

Some Effects of Leading-Edge Sweep on Boundary-Layer Transition at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Chapman, Gray T.

1961-01-01

The effects of crossflow and shock strength on transition of the laminar boundary layer behind a swept leading edge have been investigated analytically and with the aid of available experimental data. An approximate method of determining the crossflow Reynolds number on a leading edge of circular cross section at supersonic speeds is presented. The applicability of the critical crossflow criterion described by Owen and Randall for transition on swept wings in subsonic flow was examined for the case of supersonic flow over swept circular cylinders. A wide range of applicability of the subsonic critical values is indicated. The corresponding magnitude of crossflow velocity necessary to cause instability on the surface of a swept wing at supersonic speeds was also calculated and found to be small. The effects of shock strength on transition caused by Tollmien-Schlichting type of instability are discussed briefly. Changes in local Reynolds number, due to shock strength, were found analytically to have considerably more effect on transition caused by Tollmien-Schlichting instability than on transition caused by crossflow instability. Changes in the mechanism controlling transition from Tollmien-Schlichting instability to crossflow instability were found to be possible as a wing is swept back and to result in large reductions in the length of laminar flow.

Front-signal-dependent accumulation of the RHOA inhibitor FAM65B at leading edges polarizes neutrophils

PubMed Central

Gao, Kun; Tang, Wenwen; Li, Yuan; Zhang, Pingzhao; Wang, Dejie; Yu, Long; Wang, Chenji; Wu, Dianqing

2015-01-01

ABSTRACT A hallmark of neutrophil polarization is the back localization of active RHOA and phosphorylated myosin light chain (pMLC, also known as MYL2). However, the mechanism for the polarization is not entirely clear. Here, we show that FAM65B, a newly identified RHOA inhibitor, is important for the polarization. When FAM65B is phosphorylated, it binds to 14-3-3 family proteins and becomes more stable. In neutrophils, chemoattractants stimulate FAM65B phosphorylation largely depending on the signals from the front of the cells that include those mediated by phospholipase Cβ (PLCβ) and phosphoinositide 3-kinase γ (PI3Kγ), leading to FAM65B accumulation at the leading edge. Concordantly, FAM65B deficiency in neutrophils resulted in an increase in RHOA activity and localization of pMLC to the front of cells, as well as defects in chemotaxis directionality and adhesion to endothelial cells under flow. These data together elucidate a mechanism for RHOA and pMLC polarization in stimulated neutrophils through direct inhibition of RHOA by FAM65B at the leading edge. PMID:25588844

Flexible Metallic Overwrap Concept Developed for On-Orbit Repair of Space Shuttle Orbiter Leading Edges

NASA Technical Reports Server (NTRS)

Ritzert, Frank J.; Nesbitt, James A.

2005-01-01

The Columbia accident has focused attention on the critical need for on-orbit repair concepts for leading edges in the event that damage is incurred during space shuttle orbiter flight. Damage that is considered as potentially catastrophic for orbiter leading edges ranges from simple cracks to holes as large as 16 in. in diameter. NASA is particularly interested in examining potential solutions for areas of larger damage since such a problem was identified as the cause for the Columbia disaster. One possible idea for the on-orbit repair of the reinforced carbon/carbon (RCC) leading edges is an overwrap concept that would use a metallic sheet flexible enough to conform to the contours of the orbiter and robust enough to protect any problem area from catastrophic failure during reentry. The simplified view of the application of a refractory metal sheet over a mockup of shuttle orbiter panel 9, which experiences the highest temperatures on the shuttle during reentry is shown. The metallic overwrap concept is attractive because of its versatility as well as the ease with which it can be included in an onboard repair kit. Reentry of the orbiter into Earth's atmosphere imposes extreme requirements on repair materials. Temperatures can exceed 1650 C for up to 15 min in the presence of an extremely oxidizing plasma environment. Several other factors are critical, including catalysity, emissivity, and vibrational and aerodynamic loads. Materials chosen for this application will need to be evaluated with respect to high-temperature capability, resistance to oxidation, strength, coefficient of thermal expansion, and thermal conductivity. The temperature profile across panel 9 during reentry as well as a schematic of the overwrap concept itself is shown.

Preliminary Investigation in the NACA Low-Turbulence Tunnel of Low-drag Airfoil Sections Suitable for Admitting Air at the Leading Edge

NASA Technical Reports Server (NTRS)

von Doenhoff, Albert E.; Horton, Elmer A.

1942-01-01

An investigation was carried out in the NACA low-turbulence tunnel to develop low-drag airfoil sections suitable for admitting air at the leading edge. A thickness distribution having the desired type of pressure distribution was found from tests of a flexible model. Other airfoil shapes were derived from this original shape by varying the thickness, the camper, the leading-edge radius, and the size of the leading-edge opening. Data are presented giving the characteristics of the airfoil shapes in the range of lift coefficients for high-speed and cruising flight. Shapes have been developed which show no substantial increases in drag over that of the same position along the chord. Many of these shapes appear to have higher critical compressibility speeds than plain airfoils of the same thickness. Low-drag airfoil sections have been developed with openings in the leading edge as large as 41.5 percent of the maximum thickness. The range of lift coefficients for low drag in several cases is nearly as large as that of the corresponding plain airfoil sections. Preliminary measurements of maximum lift characteristics indicate that nose-opening sections of the type herein considered may not produce any marked effects on the maximum lift coefficient.

The non-receptor tyrosine kinase Lyn controls neutrophil adhesion by recruiting the CrkL–C3G complex and activating Rap1 at the leading edge

PubMed Central

He, Yuan; Kapoor, Ashish; Cook, Sara; Liu, Shubai; Xiang, Yang; Rao, Christopher V.; Kenis, Paul J. A.; Wang, Fei

2011-01-01

Establishing new adhesions at the extended leading edges of motile cells is essential for stable polarity and persistent motility. Despite recent identification of signaling pathways that mediate polarity and chemotaxis in neutrophils, little is known about molecular mechanisms governing cell–extracellular-matrix (ECM) adhesion in these highly polarized and rapidly migrating cells. Here, we describe a signaling pathway in neutrophils that is essential for localized integrin activation, leading edge attachment and persistent migration during chemotaxis. This pathway depends upon Gi-protein-mediated activation and leading edge recruitment of Lyn, a non-receptor tyrosine kinase belonging to the Src kinase family. We identified the small GTPase Rap1 as a major downstream effector of Lyn to regulate neutrophil adhesion during chemotaxis. Depletion of Lyn in neutrophil-like HL-60 cells prevented chemoattractant-induced Rap1 activation at the leading edge of the cell, whereas ectopic expression of Rap1 largely rescued the defects induced by Lyn depletion. Furthermore, Lyn controls spatial activation of Rap1 by recruiting the CrkL–C3G protein complex to the leading edge. Together, these results provide novel mechanistic insights into the poorly understood signaling network that controls leading edge adhesion during chemotaxis of neutrophils, and possibly other amoeboid cells. PMID:21628423

Mechanisms of leading edge protrusion in interstitial migration

PubMed Central

Wilson, Kerry; Lewalle, Alexandre; Fritzsche, Marco; Thorogate, Richard; Duke, Tom; Charras, Guillaume

2013-01-01

While the molecular and biophysical mechanisms underlying cell protrusion on two-dimensional substrates are well understood, our knowledge of the actin structures driving protrusion in three-dimensional environments is poor, despite relevance to inflammation, development and cancer. Here we report that, during chemotactic migration through microchannels with 5 μm × 5 μm cross-sections, HL60 neutrophil-like cells assemble an actin-rich slab filling the whole channel cross-section at their front. This leading edge comprises two distinct F-actin networks: an adherent network that polymerizes perpendicular to cell-wall interfaces and a ‘free’ network that grows from the free membrane at the cell front. Each network is polymerized by a distinct nucleator and, due to their geometrical arrangement, the networks interact mechanically. On the basis of our experimental data, we propose that, during interstitial migration, medial growth of the adherent network compresses the free network preventing its retrograde movement and enabling new polymerization to be converted into forward protrusion. PMID:24305616

The FAK–Arp2/3 interaction promotes leading edge advance and haptosensing by coupling nascent adhesions to lamellipodia actin

PubMed Central

Swaminathan, Vinay; Fischer, R. S.; Waterman, Clare M.

2016-01-01

Cell migration is initiated in response to biochemical or physical cues in the environment that promote actin-mediated lamellipodial protrusion followed by the formation of nascent integrin adhesions (NAs) within the protrusion to drive leading edge advance. Although FAK is known to be required for cell migration through effects on focal adhesions, its role in NA formation and lamellipodial dynamics is unclear. Live-cell microscopy of FAK−/− cells with expression of phosphorylation deficient or a FERM-domain mutant deficient in Arp2/3 binding revealed a requirement for FAK in promoting the dense formation, transient stabilization, and timely turnover of NA within lamellipodia to couple actin-driven protrusion to adhesion and advance of the leading edge. Phosphorylation on Y397 of FAK promotes dense NA formation but is dispensable for transient NA stabilization and leading edge advance. In contrast, transient NA stabilization and advance of the cell edge requires FAK–Arp2/3 interaction, which promotes Arp2/3 localization to NA and reduces FAK activity. Haptosensing of extracellular matrix (ECM) concentration during migration requires the interaction between FAK and Arp2/3, whereas FAK phosphorylation modulates mechanosensing of ECM stiffness during spreading. Taken together, our results show that mechanistically separable functions of FAK in NA are required for cells to distinguish distinct properties of their environment during migration. PMID:26842895

Legacy of the Space Shuttle from an Aerodynamic and Aerothermodynamic Perspective

NASA Technical Reports Server (NTRS)

Martin, Fred W.

2011-01-01

The development of the Space Shuttle Orbiter thermal protection system heating environment is described from a design stand point that began in the early 1970s. The desire for a light weight, reusable heat shield required the development of new technology, relative to previous manned spacecraft, and a systems approach to the design of the vehicle, entry guidance, and thermal protection system. Several unanticipated issues had to be resolved in both the entry and ascent phases of flight, which are discussed at a high level. During the life of the Program, significant improvements in computing power and numerical methods have been applied to Space Shuttle aerodynamic and aerothermodynamic issues, with the Shuttle Program often being the motivation, and or sponsor of the analysis development.

Initial development of an ablative leading edge for the space shuttle orbiter

NASA Technical Reports Server (NTRS)

Daforno, G.; Rose, L.; Graham, J.; Roy, P.

1974-01-01

A state-of-the-art preliminary design for typical wing areas is developed. Seven medium-density ablators (with/without honeycomb, flown on Apollo, Prime, X15A2) are evaluated. The screening tests include: (1) leading-edge models sequentially subjected to ascent heating, cold soak, entry heating, post-entry pressure fluctuations, and touchdown shock, and (2) virgin/charred models subjected to bondline strains. Two honeycomb reinforced 30 pcf elastomeric ablators were selected. Roughness/recession degradation of low speed aerodynamics appears acceptable. The design, including attachments, substructure and joints, is presented.

Multi-Component Diffusion with Application To Computational Aerothermodynamics

NASA Technical Reports Server (NTRS)

Sutton, Kenneth; Gnoffo, Peter A.

1998-01-01

The accuracy and complexity of solving multicomponent gaseous diffusion using the detailed multicomponent equations, the Stefan-Maxwell equations, and two commonly used approximate equations have been examined in a two part study. Part I examined the equations in a basic study with specified inputs in which the results are applicable for many applications. Part II addressed the application of the equations in the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) computational code for high-speed entries in Earth's atmosphere. The results showed that the presented iterative scheme for solving the Stefan-Maxwell equations is an accurate and effective method as compared with solutions of the detailed equations. In general, good accuracy with the approximate equations cannot be guaranteed for a species or all species in a multi-component mixture. 'Corrected' forms of the approximate equations that ensured the diffusion mass fluxes sum to zero, as required, were more accurate than the uncorrected forms. Good accuracy, as compared with the Stefan- Maxwell results, were obtained with the 'corrected' approximate equations in defining the heating rates for the three Earth entries considered in Part II.

PIV Study on Flow around Leading-Edge Slat of 30P30N Airfoil

NASA Astrophysics Data System (ADS)

Ando, Ryosuke; Onishi, Yusaku; Sakakibara, Jun

2017-11-01

We measured flow velocity distribution around leading-edge slat using PIV. Simultaneously, noise measurement using microphone was also performed. A leading-edge slat and main wing model having a chord length of 160 mm was placed in the tunnel with free stream velocity of about 26m/s and chord Reynolds number of 2.8 x 105. Angle of attack was changed from 4 degrees to 10 degrees at two degree intervals. In this experiment, we investigated the relationship between the unsteady flow condition and the noise. At 4 degrees in the angle of attack, vortices shedding from the slat cusp were moved to the downstream. At 6 degrees or more, flow velocity distributions show that vortices were reattached on the slat lower surface and the flow in the slat cove recirculated. In FFT analysis of noise measurement, at 6 degrees in the angle of attack, there were some peaks on low frequency area and dominant peak on high frequency area was found. At 8 degrees or more, there were also some peaks on low frequency area. But dominant peak on high frequency area disappeared.

Flight test operations using an F-106B research airplane modified with a wing leading-edge vortex flap

NASA Technical Reports Server (NTRS)

Dicarlo, Daniel J.; Brown, Philip W.; Hallissy, James B.

1992-01-01

Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.

Visualization of leading edge vortices on a series of flat plate delta wings

NASA Technical Reports Server (NTRS)

Payne, Francis M.; Ng, T. Terry; Nelson, Robert C.

1991-01-01

A summary of flow visualization data obtained as part of NASA Grant NAG2-258 is presented. During the course of this study, many still and high speed motion pictures were taken of the leading edge vortices on a series of flat plate delta wings at varying angles of attack. The purpose is to present a systematic collection of photographs showing the state of vortices as a function of the angle of attack for the four models tested.

Aerothermodynamic Analyses of Towed Ballutes

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Buck, Greg; Moss, James N.; Nielsen, Eric; Berger, Karen; Jones, William T.; Rudavsky, Rena

2006-01-01

A ballute (balloon-parachute) is an inflatable, aerodynamic drag device for application to planetary entry vehicles. Two challenging aspects of aerothermal simulation of towed ballutes are considered. The first challenge, simulation of a complete system including inflatable tethers and a trailing toroidal ballute, is addressed using the unstructured-grid, Navier-Stokes solver FUN3D. Auxiliary simulations of a semi-infinite cylinder using the rarefied flow, Direct Simulation Monte Carlo solver, DSV2, provide additional insight into limiting behavior of the aerothermal environment around tethers directly exposed to the free stream. Simulations reveal pressures higher than stagnation and corresponding large heating rates on the tether as it emerges from the spacecraft base flow and passes through the spacecraft bow shock. The footprint of the tether shock on the toroidal ballute is also subject to heating amplification. Design options to accommodate or reduce these environments are discussed. The second challenge addresses time-accurate simulation to detect the onset of unsteady flow interactions as a function of geometry and Reynolds number. Video of unsteady interactions measured in the Langley Aerothermodynamic Laboratory 20-Inch Mach 6 Air Tunnel and CFD simulations using the structured grid, Navier-Stokes solver LAURA are compared for flow over a rigid spacecraft-sting-toroid system. The experimental data provides qualitative information on the amplitude and onset of unsteady motion which is captured in the numerical simulations. The presence of severe unsteady fluid - structure interactions is undesirable and numerical simulation must be able to predict the onset of such motion.

How differential deflection of the inboard and outboard leading-edge flaps affected the handling qua

NASA Technical Reports Server (NTRS)

2002-01-01

How differential deflection of the inboard and outboard leading-edge flaps affected the handling qualities of this modified F/A-18A was evaluated during the first check flight in the Active Aeroelastic Wing program at NASA's Dryden Flight Research Center. The Active Aeroelastic Wing program at NASA's Dryden Flight Research Center seeks to determine the advantages of twisting flexible wings for primary maneuvering roll control at transonic and supersonic speeds, with traditional control surfaces such as ailerons and leading-edge flaps used to aerodynamically induce the twist. From flight test and simulation data, the program intends to develop structural modeling techniques and tools to help design lighter, more flexible high aspect-ratio wings for future high-performance aircraft, which could translate to more economical operation or greater payload capability. AAW flight tests began in November, 2002 with checkout and parameter-identification flights. Based on data obtained during the first flight series, new flight control software will be developed and a second series of research flights will then evaluate the AAW concept in a real-world environment. The program uses wings that were modified to the flexibility of the original pre-production F-18 wing. Other modifications include a new actuator to operate the outboard leading edge flap over a greater range and rate, and a research flight control system to host the aeroelastic wing control laws. The Active Aeroelastic Wing Program is jointly funded and managed by the Air Force Research Laboratory and NASA Dryden Flight Research Center, with Boeing's Phantom Works as prime contractor for wing modifications and flight control software development. The F/A-18A aircraft was provided by the Naval Aviation Systems Test Team and modified for its research role by NASA Dryden technicians.

Nonlinear Modeling and Initial Condition Estimation for Identifying the Aerothermodynamic Environment of the Space Shuttle Orbiter.

DTIC Science & Technology

1983-12-01

8217°%. .. o..’% - * 2’ . *. -o- . *o.oo o ,o ;j ’:-’ List of Figures Figure Page 1. System Identification of the Aerothermodynamic Environment of... System (STS) has of fered the engineering community a unique opportunity to flight test a reentry, hypersonic vehicle. The key 4 to the Shuttle’s...of the system (Refs. 7,8,9,10). Although the initial test flights have now been completed, data analysis and expansion of the existing data base

Aeroelastic loads prediction for an arrow wing. Task 3: Evaluation of the Boeing three-dimensional leading-edge vortex code

NASA Technical Reports Server (NTRS)

Manro, M. E.

1983-01-01

Two separated flow computer programs and a semiempirical method for incorporating the experimentally measured separated flow effects into a linear aeroelastic analysis were evaluated. The three dimensional leading edge vortex (LEV) code is evaluated. This code is an improved panel method for three dimensional inviscid flow over a wing with leading edge vortex separation. The governing equations are the linear flow differential equation with nonlinear boundary conditions. The solution is iterative; the position as well as the strength of the vortex is determined. Cases for both full and partial span vortices were executed. The predicted pressures are good and adequately reflect changes in configuration.

Active control of noise amplification in the flow over a square leading-edge flat plate utilizing DBD plasma actuator

NASA Astrophysics Data System (ADS)

Yadong, HUANG; Benmou, ZHOU

2018-05-01

Perturbation is generally considered as the flow noise, and its energy can gain transient growth in the separation bubble. The amplified perturbations may cause unstable Kelvin–Helmohltz vortices which induce the three-dimensional transition. Active control of noise amplification via dielectric barrier discharge plasma actuator in the flow over a square leading-edge flat plate is numerically studied. The actuator is installed near the plate leading-edge where the separation bubble is formed. The maximum energy amplification of perturbations is positively correlated with the separation bubble scale which decreases with the increasing control parameters. As the magnitude of noise amplification is reduced, the laminar-turbulent transition is successfully suppressed.

Front-signal-dependent accumulation of the RHOA inhibitor FAM65B at leading edges polarizes neutrophils.

PubMed

Gao, Kun; Tang, Wenwen; Li, Yuan; Zhang, Pingzhao; Wang, Dejie; Yu, Long; Wang, Chenji; Wu, Dianqing

2015-03-01

A hallmark of neutrophil polarization is the back localization of active RHOA and phosphorylated myosin light chain (pMLC, also known as MYL2). However, the mechanism for the polarization is not entirely clear. Here, we show that FAM65B, a newly identified RHOA inhibitor, is important for the polarization. When FAM65B is phosphorylated, it binds to 14-3-3 family proteins and becomes more stable. In neutrophils, chemoattractants stimulate FAM65B phosphorylation largely depending on the signals from the front of the cells that include those mediated by phospholipase Cβ (PLCβ) and phosphoinositide 3-kinase γ (PI3Kγ), leading to FAM65B accumulation at the leading edge. Concordantly, FAM65B deficiency in neutrophils resulted in an increase in RHOA activity and localization of pMLC to the front of cells, as well as defects in chemotaxis directionality and adhesion to endothelial cells under flow. These data together elucidate a mechanism for RHOA and pMLC polarization in stimulated neutrophils through direct inhibition of RHOA by FAM65B at the leading edge. © 2015. Published by The Company of Biologists Ltd.

The Influence of Clocking Angle of the Projectile on the Simulated Impact Response of a Shuttle Leading Edge Wing Panel

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.; Lyle, Karen H.; Spellman, Regina L.

2005-01-01

An analytical study was conducted to determine the influence of clocking angle of a foam projectile impacting a space shuttle leading edge wing panel. Four simulations were performed using LS-DYNA. The leading edge panels are fabricated of multiple layers of reinforced carbon-carbon (RCC) material. The RCC material was represented using Mat 58, which is a material property that can be used for laminated composite fabrics. Simulations were performed of a rectangular-shaped foam block, weighing 0.23-lb., impacting RCC Panel 9 on the top surface. The material properties of the foam were input using Mat 83. The impact velocity was 1,000 ft/s along the Orbiter X-axis. In two models, the foam impacted on a corner, in one model the foam impacted the panel initially on the 2-in.-long edge, and in the last model the foam impacted the panel on the 7-in.- long edge. The simulation results are presented as contour plots of first principal infinitesimal strain and time history plots of contact force and internal and kinetic energy of the foam and RCC panel.

OAST Space Theme Workshop. Volume 3: Working group summary. 9: Aerothermodynamics (M-3). A: Statement. B: Technology needs (form 1). C. Priority assessment (form 2). D. Additional assessments

NASA Technical Reports Server (NTRS)

1976-01-01

Twelve aerothermodynamic space technology needs were identified to reduce the design uncertainties in aerodynamic heating and forces experienced by heavy lift launch vehicles, orbit transfer vehicles, and advanced single stage to orbit vehicles for the space transportation system, and for probes, planetary surface landers, and sample return vehicles for solar system exploration vehicles. Research and technology needs identified include: (1) increasing the fluid dynamics capability by at least two orders of magnitude by developing an advanced computer processor for the solution of fluid dynamic problems with improved software; (2) predicting multi-engine base flow fields for launch vehicles; and (3) developing methods to conserve energy in aerothermodynamic ground test facilities.

Alignment of leading-edge and peak-picking time of arrival methods to obtain accurate source locations

DOE Office of Scientific and Technical Information (OSTI.GOV)

Roussel-Dupre, R.; Symbalisty, E.; Fox, C.

2009-08-01

The location of a radiating source can be determined by time-tagging the arrival of the radiated signal at a network of spatially distributed sensors. The accuracy of this approach depends strongly on the particular time-tagging algorithm employed at each of the sensors. If different techniques are used across the network, then the time tags must be referenced to a common fiducial for maximum location accuracy. In this report we derive the time corrections needed to temporally align leading-edge, time-tagging techniques with peak-picking algorithms. We focus on broadband radio frequency (RF) sources, an ionospheric propagation channel, and narrowband receivers, but themore » final results can be generalized to apply to any source, propagation environment, and sensor. Our analytic results are checked against numerical simulations for a number of representative cases and agree with the specific leading-edge algorithm studied independently by Kim and Eng (1995) and Pongratz (2005 and 2007).« less

AMPK activity regulates trafficking of mitochondria to the leading edge during cell migration and matrix invasion

PubMed Central

Cunniff, Brian; McKenzie, Andrew J.; Heintz, Nicholas H.; Howe, Alan K.

2016-01-01

Cell migration is a complex behavior involving many energy-expensive biochemical events that iteratively alter cell shape and location. Mitochondria, the principal producers of cellular ATP, are dynamic organelles that fuse, divide, and relocate to respond to cellular metabolic demands. Using ovarian cancer cells as a model, we show that mitochondria actively infiltrate leading edge lamellipodia, thereby increasing local mitochondrial mass and relative ATP concentration and supporting a localized reversal of the Warburg shift toward aerobic glycolysis. This correlates with increased pseudopodial activity of the AMP-activated protein kinase (AMPK), a critically important cellular energy sensor and metabolic regulator. Furthermore, localized pharmacological activation of AMPK increases leading edge mitochondrial flux, ATP content, and cytoskeletal dynamics, whereas optogenetic inhibition of AMPK halts mitochondrial trafficking during both migration and the invasion of three-dimensional extracellular matrix. These observations indicate that AMPK couples local energy demands to subcellular targeting of mitochondria during cell migration and invasion. PMID:27385336

Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.; Tso, Jin

1993-01-01

Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Predicted Static Aeroelastic Effects on Wings with Supersonic Leading Edges and Streamwise Tips

NASA Technical Reports Server (NTRS)

Brown, Stuart C.

1959-01-01

A method is presented for calculation of static aeroelastic effects on wings with supersonic leading edges and streamwise tips. Both chord-wise and spanwise deflections are taken into account. Aerodynamic and structural forces are introduced in influence coefficient form; the former are developed from linearized supersonic wing theory and the latter are assumed to be known from load-deflection tests or theory. The predicted effects of flexibility on lateral-control effectiveness, damping in roll, and lift-curve slope are shown for a low-aspect-ratio wing at Mach numbers of 1.25 and 2.60. The control effectiveness is shown for a trailing-edge aileron, a tip aileron, and a slot-deflector spoiler located along the 0.70 chord line. The calculations indicate that the tip aileron is particularly attractive from an aeroelastic standpoint, because the changes in effectiveness with dynamic pressure are small compared to the changes in effectiveness of the trailing-edge aileron and slot-deflector spoiler. The effects of making several simplifying assumptions in the example calculations are shown. The use of a modified strip theory to determine the aerodynamic influence coefficients gave adequate results only for the high Mach number case. Elimination of chordwise bending in the structural influence coefficients exaggerated the aeroelastic effects on rolling-moment and lift coefficients for both Mach numbers.

Simulator study of the stall departure characteristics of a light general aviation airplane with and without a wing-leading-edge modification

NASA Technical Reports Server (NTRS)

Riley, D. R.

1985-01-01

A six-degree-of-freedom nonlinear simulation was developed for a two-place, single-engine, low-wing general aviation airplane for the stall and initial departure regions of flight. Two configurations, one with and one without an outboard wing-leading-edge modification, were modeled. The math models developed are presented simulation predictions and flight-test data for validation purposes and simulation results for the two configurations for various maneuvers and power settings are compared to show the beneficial influence of adding the wing-leading-edge modification.

A Review of Hypersonics Aerodynamics, Aerothermodynamics and Plasmadynamics Activities within NASA's Fundamental Aeronautics Program

NASA Technical Reports Server (NTRS)

Salas, Manuel D.

2007-01-01

The research program of the aerodynamics, aerothermodynamics and plasmadynamics discipline of NASA's Hypersonic Project is reviewed. Details are provided for each of its three components: 1) development of physics-based models of non-equilibrium chemistry, surface catalytic effects, turbulence, transition and radiation; 2) development of advanced simulation tools to enable increased spatial and time accuracy, increased geometrical complexity, grid adaptation, increased physical-processes complexity, uncertainty quantification and error control; and 3) establishment of experimental databases from ground and flight experiments to develop better understanding of high-speed flows and to provide data to validate and guide the development of simulation tools.

Effects of Alternate Leading Edge Cutback on the Space Shuttle Main Engine Low Pressure Fuel Pump

NASA Technical Reports Server (NTRS)

Mulder, Andrew; Skelley, Stephen

2016-01-01

A higher order cavitation oscillation observed in the SSME low pressure fuel pump has been eliminated in water flow testing of a modified subscale replica of the inducer. The low pressure pump was modified by removing the outboard sections of two opposing blades of the four-bladed inducer, blending the "cutback" regions into the blades at the leading edge and tip, and removing material on the suction sides to decrease the exposed leading edge thickness. The leading edge tips of the cutback blades were moved approximately 25 degrees from their previous locations, thereby increasing one blade to blade spacing, decreasing the second, while simultaneously moving the cutback tips downstream. The test was conducted in MSFC's inducer test loop at scaled operating conditions in degassed and filtered water. In addition to eliminating HOC across the entire scaled operating regime, rotating cavitation was suppressed while the range of both alternate blade and asymmetric cavitation were increased. These latter phenomena, and more significantly, the shifts between these cavitation modes also resulted in significant changes to the head coefficient at low cavitation numbers. Reverse flow was detected at a slightly larger flow coefficient with the cutback inducer and suction capability was reduced by approximately 1 velocity head at and above approximately 90% of the reference flow coefficient. These performance changes along with more intense reverse flow are consistent with poor flow area management and increased incidence in the cutback region. Although the test demonstrated that the inducer modification was successful at eliminating the higher order cavitation across the entire scaled operating regime, different, previously unobserved, cavitation oscillations were introduced and significant performance penalties were imposed.

Uncertainty Assessment of Hypersonic Aerothermodynamics Prediction Capability

NASA Technical Reports Server (NTRS)

Bose, Deepak; Brown, James L.; Prabhu, Dinesh K.; Gnoffo, Peter; Johnston, Christopher O.; Hollis, Brian

2011-01-01

The present paper provides the background of a focused effort to assess uncertainties in predictions of heat flux and pressure in hypersonic flight (airbreathing or atmospheric entry) using state-of-the-art aerothermodynamics codes. The assessment is performed for four mission relevant problems: (1) shock turbulent boundary layer interaction on a compression corner, (2) shock turbulent boundary layer interaction due a impinging shock, (3) high-mass Mars entry and aerocapture, and (4) high speed return to Earth. A validation based uncertainty assessment approach with reliance on subject matter expertise is used. A code verification exercise with code-to-code comparisons and comparisons against well established correlations is also included in this effort. A thorough review of the literature in search of validation experiments is performed, which identified a scarcity of ground based validation experiments at hypersonic conditions. In particular, a shortage of useable experimental data at flight like enthalpies and Reynolds numbers is found. The uncertainty was quantified using metrics that measured discrepancy between model predictions and experimental data. The discrepancy data is statistically analyzed and investigated for physics based trends in order to define a meaningful quantified uncertainty. The detailed uncertainty assessment of each mission relevant problem is found in the four companion papers.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 2; Small-Radius Leading Edge

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg. delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 84 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 4: Large-radius leading edge

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Aerodynamic heating on AFE due to nonequilibrium flow with variable entropy at boundary layer edge

NASA Technical Reports Server (NTRS)

Ting, P. C.; Rochelle, W. C.; Bouslog, S. A.; Tam, L. T.; Scott, C. D.; Curry, D. M.

1991-01-01

A method of predicting the aerobrake aerothermodynamic environment on the NASA Aeroassist Flight Experiment (AFE) vehicle is described. Results of a three dimensional inviscid nonequilibrium solution are used as input to an axisymmetric nonequilibrium boundary layer program to predict AFE convective heating rates. Inviscid flow field properties are obtained from the Euler option of the Viscous Reacting Flow (VRFLO) code at the boundary layer edge. Heating rates on the AFE surface are generated with the Boundary Layer Integral Matrix Procedure (BLIMP) code for a partially catalytic surface composed of Reusable Surface Insulation (RSI) times. The 1864 kg AFE will fly an aerobraking trajectory, simulating return from geosynchronous Earth orbit, with a 75 km perigee and a 10 km/sec entry velocity. Results of this analysis will provide principal investigators and thermal analysts with aeroheating environments to perform experiment and thermal protection system design.

Fracture Mechanics Analyses of the Slip-Side Joggle Regions of Wing-Leading-Edge Panels

NASA Technical Reports Server (NTRS)

Raju, Ivatury S.; Knight, Norman F., Jr.; Song, Kyongchan; Phillips, Dawn R.

2011-01-01

The Space Shuttle wing-leading edge consists of panels that are made of reinforced carbon-carbon. Coating spallation was observed near the slip-side region of the panels that experience extreme heating. To understand this phenomenon, a root-cause investigation was conducted. As part of that investigation, fracture mechanics analyses of the slip-side joggle regions of the hot panels were conducted. This paper presents an overview of the fracture mechanics analyses.

Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and From 2.50 to 4.05

NASA Technical Reports Server (NTRS)

ONeal, Robert L.

1960-01-01

A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.

Experimental study of pressure and heating rate on a swept cylindrical leading edge resulting from swept shock wave interference. M.S. Thesis

NASA Technical Reports Server (NTRS)

Glass, Christopher E.

1989-01-01

The effects of cylindrical leading edge sweep on surface pressure and heat transfer rate for swept shock wave interference were investigated. Experimental tests were conducted in the Calspan 48-inch Hypersonic Shock Tunnel at a nominal Mach number of 8, nominal unit Reynolds number of 1.5 x 10 to the 6th power per foot, leading edge and incident shock generator sweep angles of 0, 15, and 30 deg, and incident shock generator angle-of-attack fixed at 12.5 deg. Detailed surface pressure and heat transfer rate on the cylindircal leading edge of a swept shock wave interference model were measured at the region of the maximum surface pressure and heat transfer rate. Results show that pressure and heat transfer rate on the cylindrical leading edge of the shock wave interference model were reduced as the sweep was increased over the range of tested parameters. Peak surface pressure and heat transfer rate on the cylinder were about 10 and 30 times the undisturbed flow stagnation point value, respectively, for the 0 deg sweep test. A comparison of the 15 and 30 deg swept results with the 0 deg swept results showed that peak pressure was reduced about 13 percent and 44 percent, respectively, and peak heat transfer rate was reduced about 7 percent and 27 percent, respectively.

Subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration having spanwise leading-edge vortex enhancement

NASA Technical Reports Server (NTRS)

Huffman, J. K.; Fox, C. H., Jr.

1977-01-01

A configuration which integrates a close coupled canard wing combination, spanwise blowing for enhancement of the wing leading edge vortex, an engine-over-wing concept, and a wing trailing edge coanda-effect flap is studied. The data on the configuration are presented in tabular from without discussion. The investigation was conducted in the Langley 7- by 10-foot high speed tunnel at a Mach number of 0.166 through an angle-of-attack range from -2 to 22 deg. Rectangular main engine nozzles of aspect ratio 4, 6, and 8 were tested over a momentum coefficient range from 1.0 to 1.8.

Experimental Study of Shock Wave Interference Heating on a Cylindrical Leading Edge. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Wieting, Allan R.

1987-01-01

An experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet at Mach numbers of 6.3, 6.5, and 8.0 is presented. Stream Reynolds numbers ranged from 0.5 x 106 to 4.9 x 106 per ft. and stream total temperature ranged from 2100 to 3400 R. The model consisted of a 3" dia. cylinder and a shock generation wedge articulated to angles of 10, 12.5, and 15 deg. A fundamental understanding was obtained of the fluid mechanics of shock wave interference induced flow impingement on a cylindrical leading edge and the attendant surface pressure and heat flux distributions. The first detailed heat transfer rate and pressure distributions for two dimensional shock wave interference on a cylinder was provided along with insight into the effects of specific heat variation with temperature on the phenomena. Results show that the flow around a body in hypersonic flow is altered significantly by the shock wave interference pattern that is created by an oblique shock wave from an external source intersecting the bow shock wave produced in front of the body.

Computational Aerothermodynamic Assessment of Space Shuttle Orbiter Tile Damage: Open Cavities

NASA Technical Reports Server (NTRS)

Pulsonetti, Maria; Wood, William

2005-01-01

Computational aerothermodynamic simulations of Orbiter windside tile damage in flight were performed in support of the Space Shuttle Return-to-Flight effort. The simulations were performed for both hypervelocity flight and low-enthalpy wind tunnel conditions and contributed to the Return-to-Flight program by providing information to support a variety of damage scenario analyses. Computations at flight conditions were performed at or very near the peak heating trajectory point for multiple damage scenarios involving damage windside acreage reaction cured glass (RCG) coated silica tile(s). The cavities formed by the missing tile examined in this study were relatively short leading to flow features which indicated open cavity behavior. Results of the computations indicated elevated heating bump factor levels predicted for flight over the predictions for wind tunnel conditions. The peak heating bump factors, defined as the local heating to a reference value upstream of the cavity, on the cavity floor for flight simulation were 67% larger than the peak wind tunnel simulation value. On the downstream face of the cavity the flight simulation values were 60% larger than the wind tunnel simulation values. On the outer mold line (OML) downstream of the cavity, the flight values are about 20% larger than the wind tunnel simulation values. The higher heating bump factors observed in the flight simulations were due to the larger driving potential in terms of energy entering the cavity for the flight simulations. This is evidenced by the larger rate of increase in the total enthalpy through the boundary layer prior to the cavity for the flight simulation.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 3: Medium-radius leading edge

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Turbine vane leading edge gas film cooling with spanwise angled coolant holes

NASA Technical Reports Server (NTRS)

Hanus, G. J.; Lecuyer, M. R.

1976-01-01

An experimental film cooling study was conducted on a 3x size model turbine vane. Injection at the leading edge was from a single row of holes angled in a spanwise direction for two configurations of holes at 18 or 35 deg to the surface. The reduction in the local Stanton number for injection at a coolant-to-mainstream density ratio of 2.18 was calculated from heat flux measurements downstream of injection. Results indicate that optimum cooling occurs near a coolant-to-mainstream velocity ratio of 0.5. Shallow injection angles appear to be most beneficial when injecting into a highly accelerated mainstream.

A Method for Computing the Core Flow in Three-Dimensional Leading-Edge Vortices. Ph.D. Thesis - North Carolina State Univ.

NASA Technical Reports Server (NTRS)

Luckring, J. M.

1985-01-01

A theory is presented for calculating the flow in the core of a separation-induced leading-edge vortex. The method is based on matching inner and outer representations of the vortex. The inner model of the vortex is based on the quasicylindrical Navier-Stokes equations; the flow is assumed to be steady, axially symmetric, and incompressible and in addition, gradients in the radial direction are assumed to be much larger then gradients in the axial direction. The outer model is based on the three-dimensional free-vortex-sheet theory, a higher-order panel method which solves the Prandtl-Glauert equation including nonlinear boundary conditions pertinent to the concentrated vorticity representation of the leading edge vortex. The resultant flow is evaluated a posteriori for evidence of incipient vortex breakdown and the critical helix angle concept, in conjunction with an adverse longitudinal pressure gradient, is found to correlate well with the occurrence of vortex breakdown at the trailing edge of delta, arrow, and diamond wings.

Effects of discontinuous drooped wing leading-edge modifications on the spinning characteristics of a low-wing general aviation airplane

NASA Technical Reports Server (NTRS)

Dicarlo, D. J.; Stough, H. P., III; Patton, J. M., Jr.

1980-01-01

Wind tunnel and flight tests were conducted to determine the effects of several discontinuous drooped wing leading-edge configurations on the spinning characteristics of a light, single-engine, low-wing research airplane. Particular emphasis was placed on the identification of modifications which would improve the spinning characteristics. The spanwise length of a discontinuous outboard droop was varied and several additional inboard segments were added to determine the influence of such leading-edge configurations on the spin behavior. Results of the study indicated that the use of only the discontinuous outboard droop, over a specific spanwise area, was most effective towards improving spin and spin recovery characteristics, whereas the segmented configurations having both inboard and outboard droop exhibited a tendency to enter a flat spin.

Suppression of Chemotaxis by SSeCKS via Scaffolding of Phosphoinositol Phosphates and the Recruitment of the Cdc42 GEF, Frabin, to the Leading Edge

PubMed Central

Ko, Hyun-Kyung; Guo, Li-wu; Su, Bing; Gao, Lingqiu; Gelman, Irwin H.

2014-01-01

Chemotaxis is controlled by interactions between receptors, Rho-family GTPases, phosphatidylinositol 3-kinases, and cytoskeleton remodeling proteins. We investigated how the metastasis suppressor, SSeCKS, attenuates chemotaxis. Chemotaxis activity inversely correlated with SSeCKS levels in mouse embryo fibroblasts (MEF), DU145 and MDA-MB-231 cancer cells. SSeCKS loss induced chemotactic velocity and linear directionality, correlating with replacement of leading edge lamellipodia with fascin-enriched filopodia-like extensions, the formation of thickened longitudinal F-actin stress fibers reaching to filopodial tips, relative enrichments at the leading edge of phosphatidylinositol (3,4,5)P3 (PIP3), Akt, PKC-ζ, Cdc42-GTP and active Src (SrcpoY416), and a loss of Rac1. Leading edge lamellipodia and chemotaxis inhibition in SSeCKS-null MEF could be restored by full-length SSeCKS or SSeCKS deleted of its Src-binding domain (ΔSrc), but not by SSeCKS deleted of its three MARCKS (myristylated alanine-rich C kinase substrate) polybasic domains (ΔPBD), which bind PIP2 and PIP3. The enrichment of activated Cdc42 in SSeCKS-null leading edge filopodia correlated with recruitment of the Cdc42-specific guanine nucleotide exchange factor, Frabin, likely recruited via multiple PIP2/3-binding domains. Frabin knockdown in SSeCKS-null MEF restores leading edge lamellipodia and chemotaxis inhibition. However, SSeCKS failed to co-immunoprecipitate with Rac1, Cdc42 or Frabin. Consistent with the notion that chemotaxis is controlled by SSeCKS-PIP (vs. -Src) scaffolding activity, constitutively-active phosphatidylinositol 3-kinase could override the ability of the Src inhibitor, SKI-606, to suppress chemotaxis and filopodial enrichment of Frabin in SSeCKS-null MEF. Our data suggest a role for SSeCKS in controlling Rac1 vs. Cdc42-induced cellular dynamics at the leading chemotactic edge through the scaffolding of phospholipids and signal mediators, and through the reorganization of the

Clathrin-independent carriers form a high capacity endocytic sorting system at the leading edge of migrating cells

PubMed Central

Howes, Mark T.; Kirkham, Matthew; Riches, James; Cortese, Katia; Walser, Piers J.; Simpson, Fiona; Hill, Michelle M.; Jones, Alun; Lundmark, Richard; Lindsay, Margaret R.; Hernandez-Deviez, Delia J.; Hadzic, Gordana; McCluskey, Adam; Bashir, Rumasia; Liu, Libin; Pilch, Paul; McMahon, Harvey; Robinson, Phillip J.; Hancock, John F.; Mayor, Satyajit

2010-01-01

Although the importance of clathrin- and caveolin-independent endocytic pathways has recently emerged, key aspects of these routes remain unknown. Using quantitative ultrastructural approaches, we show that clathrin-independent carriers (CLICs) account for approximately three times the volume internalized by the clathrin-mediated endocytic pathway, forming the major pathway involved in uptake of fluid and bulk membrane in fibroblasts. Electron tomographic analysis of the 3D morphology of the earliest carriers shows that they are multidomain organelles that form a complex sorting station as they mature. Proteomic analysis provides direct links between CLICs, cellular adhesion turnover, and migration. Consistent with this, CLIC-mediated endocytosis of key cargo proteins, CD44 and Thy-1, is polarized at the leading edge of migrating fibroblasts, while transient ablation of CLICs impairs their ability to migrate. These studies provide the first quantitative ultrastructural analysis and molecular characterization of the major endocytic pathway in fibroblasts, a pathway that provides rapid membrane turnover at the leading edge of migrating cells. PMID:20713605

The FAK-Arp2/3 interaction promotes leading edge advance and haptosensing by coupling nascent adhesions to lamellipodia actin.

PubMed

Swaminathan, Vinay; Fischer, R S; Waterman, Clare M

2016-04-01

Cell migration is initiated in response to biochemical or physical cues in the environment that promote actin-mediated lamellipodial protrusion followed by the formation of nascent integrin adhesions (NAs) within the protrusion to drive leading edge advance. Although FAK is known to be required for cell migration through effects on focal adhesions, its role in NA formation and lamellipodial dynamics is unclear. Live-cell microscopy of FAK(-/-)cells with expression of phosphorylation deficient or a FERM-domain mutant deficient in Arp2/3 binding revealed a requirement for FAK in promoting the dense formation, transient stabilization, and timely turnover of NA within lamellipodia to couple actin-driven protrusion to adhesion and advance of the leading edge. Phosphorylation on Y397 of FAK promotes dense NA formation but is dispensable for transient NA stabilization and leading edge advance. In contrast, transient NA stabilization and advance of the cell edge requires FAK-Arp2/3 interaction, which promotes Arp2/3 localization to NA and reduces FAK activity. Haptosensing of extracellular matrix (ECM) concentration during migration requires the interaction between FAK and Arp2/3, whereas FAK phosphorylation modulates mechanosensing of ECM stiffness during spreading. Taken together, our results show that mechanistically separable functions of FAK in NA are required for cells to distinguish distinct properties of their environment during migration. © 2016 Swaminathan et al. This article is distributed by The American Society for Cell Biology under license from the author(s). Two months after publication it is available to the public under an Attribution–Noncommercial–Share Alike 3.0 Unported Creative Commons License (http://creativecommons.org/licenses/by-nc-sa/3.0).

Aerothermodynamic Design, Review on Ground Testing and CFD (Conception aerothermodynamique, revue sur les essais au sol et dynamique des fluides informatisee)

DTIC Science & Technology

2010-04-01

Aerothermodynamic Design, Review on Ground Testing and CFD (RTO-EN-AVT-186) Executive Summary The Lecture Series focus on the presentation of...impulsions ITAM et les tubes à choc DLR HEG. Les sondes à réponse rapide et les techniques de mesures instables ont été présentées ainsi que les outils de

The effects of leading edge and downstream film cooling on turbine vane heat transfer

NASA Astrophysics Data System (ADS)

Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.

1988-11-01

The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

Fracture Mechanics Analyses of the Slip-Side Joggle Regions of Wing-Leading Edge Panels

NASA Technical Reports Server (NTRS)

Raju, Ivatury S.; Knight, Norman F., Jr.; Song, Kyongchan; Phillips, Dawn R.

2010-01-01

The Space Shuttle Orbiter wing comprises of 22 leading edge panels on each side of the wing. These panels are part of the thermal protection system that protects the Orbiter wings from extreme heating that take place on the reentry in to the earth atmosphere. On some panels that experience extreme heating, liberation of silicon carbon (SiC) coating was observed on the slip side regions of the panels. Global structural and local fracture mechanics analyses were performed on these panels as a part of the root cause investigation of this coating liberation anomaly. The wing-leading-edge reinforced carbon-carbon (RCC) panels, Panel 9, T-seal 10, and Panel 10, are shown in Figure 1 and the progression of the stress analysis models is presented in Figure 2. The global structural analyses showed minimal interaction between adjacent panels and the T-seal that bridges the gap between the panels. A bounding uniform temperature is applied to a representative panel and the resulting stress distribution is examined. For this loading condition, the interlaminar normal stresses showed negligible variation in the chord direction and increased values in the vicinity of the slip-side joggle shoulder. As such, a representative span wise slice on the panel can be taken and the cross section can be analyzed using plane strain analysis.

The effects of leading edge and downstream film cooling on turbine vane heat transfer

NASA Technical Reports Server (NTRS)

Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.

1988-01-01

The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

Investigation of airfoil leading edge separation control with nanosecond plasma actuator

NASA Astrophysics Data System (ADS)

Zheng, J. G.; Cui, Y. D.; Zhao, Z. J.; Li, J.; Khoo, B. C.

2016-11-01

A combined numerical and experimental investigation of airfoil leading edge flow separation control with a nanosecond dielectric barrier discharge (DBD) plasma actuator is presented. Our study concentrates on describing dynamics of detailed flow actuation process and elucidating the nanosecond DBD actuation mechanism. A loose coupling methodology is employed to perform simulation, which consists of a self-similar plasma model for the description of pulsed discharge and two-dimensional Reynolds averaged Navier-Stokes (RANS) equations for the calculation of external airflow. A series of simulations of poststall flows around a NACA0015 airfoil is conducted with a Reynolds number range covering both low and high Re at Re=(0.05 ,0.15 ,1.2 ) ×106 . Meanwhile, wind-tunnel experiment is performed for two low Re flows to measure aerodynamic force on airfoil model and transient flow field with time-resolved particle image velocimetry (PIV). The PIV measurement provides possibly the clearest view of flow reattachment process under the actuation of a nanosecond plasma actuator ever observed in experiments, which is highly comparable to that predicted by simulation. It is found from the detailed simulation that the discharge-induced residual heat rather than shock wave plays a dominant role in flow control. For any leading edge separations, the preliminary flow reattachment is realized by residual heat-induced spanwise vortices. After that, the nanosecond actuator functions by continuing exciting flow instability at poststall attack angles or acting as an active trip near stall angle. As a result, the controlled flow is characterized by a train of repetitive, downstream moving vortices over suction surface or an attached turbulent boundary layer, which depends on both angle of attack and Reynolds number. The advection of residual temperature with external flow offers a nanosecond plasma actuator a lot of flexibility to extend its influence region. Animations are provided for

A Unit-Problem Investigation of Blunt Leading-Edge Separation Motivated by AVT-161 SACCON Research

NASA Technical Reports Server (NTRS)

Luckring, James M.; Boelens, Okko J.

2011-01-01

A research effort has been initiated to examine in more detail some of the challenging flow fields discovered from analysis of the SACCON configuration aerodynamics. This particular effort is oriented toward a diamond wing investigation specifically designed to isolate blunt leading-edge separation phenomena relevant to the SACCON investigations of the present workshop. The approach taken to design this new effort is reviewed along with the current status of the program.

Rarefaction and Non-equilibrium Effects in Hypersonic Flows about Leading Edges of Small Bluntness

NASA Astrophysics Data System (ADS)

Ivanov, Mikhail; Khotyanovsky, Dmitry; Kudryavtsev, Alexey; Shershnev, Anton; Bondar, Yevgeniy; Yonemura, Shigeru

2011-05-01

A hypersonic flow about a cylindrically blunted thick plate at a zero angle of attack is numerically studied with the kinetic (DSMC) and continuum (Navier-Stokes equations) approaches. The Navier-Stokes equations with velocity slip and temperature jump boundary conditions correctly predict the flow fields and surface parameters for values of the Knudsen number (based on the radius of leading edge curvature) smaller than 0.1. The results of computations demonstrate significant effects of the entropy layer on the boundary layer characteristics.

Aerodynamic Investigation of C-141 Leading Edge Modifications for Cruise Drag Reduction--Test 2.

DTIC Science & Technology

1981-04-01

82171.6 1 ((1 I III 11 1 - MI( RO( )FY R~ I IN II I (I 041PI AD A099 6 6 2 _ AFWAL-TR-81-3032 AERODYNAMIC INVESTIGATION OF C-141 LEADING EDGE MODIFICATIONS...FOR CRUISE DRAG REDUCTION -- TEST 2 Robert A. Large, Captain, USAF Aerodynamics & Airframe Branch Aeromechanics Division Flight Dynamics Laboratory...KEMloUA Ch, Aerodynamics & Airframe Br Aeromechanics Division FOR THE COMMANDER hie" Aeromechanics Division "If your address has changed, if you wish to

Stagnation region gas film cooling for turbine blade leading edge applications

NASA Technical Reports Server (NTRS)

Luckey, D. W.; Winstanley, D. K.; Hanus, G. J.; Lecuyer, M. R.

1976-01-01

An experimental investigation was conducted to model the film-cooling performance for a turbine-vane leading edge using the stagnation region of a cylinder in cross flow. Experiments were conducted with a single row of spanwise-angled coolant holes for a range of the coolant blowing ratio with a freestream-to-wall temperature ratio of about 2.1 and a Reynolds number of 170,000, characteristic of the gas-turbine environment. Data from local heat-flux measurements are presented for coolant-hole injection angles of 25, 35, and 45 deg with the row of holes located at three positions relative to the stagnation line on the cylinder. Results show the spanwise (hole-to-hole) variation of heat-flux reduction due to film cooling and indicate conditions for the optimum film-cooling performance.

Insect Residue Contamination on Wing Leading Edge Surfaces: A Materials Investigation for Mitigation

NASA Technical Reports Server (NTRS)

Lorenzi, Tyler M.; Wohl, Christopher J.; Penner, Ronald K.; Smith, Joseph G.; Siochi, Emilie J.

2011-01-01

Flight tests have shown that residue from insect strikes on aircraft wing leading edge surfaces may induce localized transition of laminar to turbulent flow. The highest density of insect populations have been observed between ground level and 153 m during light winds (2.6 -- 5.1 m/s), high humidity, and temperatures from 21 -- 29 C. At a critical residue height, dependent on the airfoil and Reynolds number, boundary layer transition from laminar to turbulent results in increased drag and fuel consumption. Although this represents a minimal increase in fuel burn for conventional transport aircraft, future aircraft designs will rely on maintaining laminar flow across a larger portion of wing surfaces to reduce fuel burn during cruise. Thus, insect residue adhesion mitigation is most critical during takeoff and initial climb to maintain laminar flow in fuel-efficient aircraft configurations. Several exterior treatments investigated to mitigate insect residue buildup (e.g., paper, scrapers, surfactants, flexible surfaces) have shown potential; however, implementation has proven to be impractical. Current research is focused on evaluation of wing leading edge surface coatings that may reduce insect residue adhesion. Initial work under NASA's Environmentally Responsible Aviation Program focused on evaluation of several commercially available products (commercial off-the-shelf, COTS), polymers, and substituted alkoxy silanes that were applied to aluminum (Al) substrates. Surface energies of these coatings were determined from contact angle data and were correlated to residual insect excrescence on coated aluminum substrates using a custom-built "bug gun." Quantification of insect excrescence surface coverage was evaluated by a series of digital photographic image processing techniques.

AERO2S - SUBSONIC AERODYNAMIC ANALYSIS OF WINGS WITH LEADING- AND TRAILING-EDGE FLAPS IN COMBINATION WITH CANARD OR HORIZONTAL TAIL SURFACES (CDC VERSION)

NASA Technical Reports Server (NTRS)

Darden, C. M.

1994-01-01

This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is

Experimental study of flow separation control on a low- Re airfoil using leading-edge protuberance method

NASA Astrophysics Data System (ADS)

Zhang, M. M.; Wang, G. F.; Xu, J. Z.

2014-04-01

An experimental study of flow separation control on a low- Re c airfoil was presently investigated using a newly developed leading-edge protuberance method, motivated by the improvement in the hydrodynamics of the giant humpback whale through its pectoral flippers. Deploying this method, the control effectiveness of the airfoil aerodynamics was fully evaluated using a three-component force balance, leading to an effectively impaired stall phenomenon and great improvement in the performances within the wide post-stall angle range (22°-80°). To understand the flow physics behind, the vorticity field, velocity field and boundary layer flow field over the airfoil suction side were examined using a particle image velocimetry and an oil-flow surface visualization system. It was found that the leading-edge protuberance method, more like low-profile vortex generator, effectively modified the flow pattern of the airfoil boundary layer through the chordwise and spanwise evolutions of the interacting streamwise vortices generated by protuberances, where the separation of the turbulent boundary layer dominated within the stall region and the rather strong attachment of the laminar boundary layer still existed within the post-stall region. The characteristics to manipulate the flow separation mode of the original airfoil indicated the possibility to further optimize the control performance by reasonably designing the layout of the protuberances.

Influence of a heated leading edge on boundary layer growth, stability, and transition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Landrum, D.B.; Macha, J.M.

1987-01-01

This paper presents the results of a combined theoretical and experimental study of the influence of a heated leading edge on the growth, stability, and transition of a two-dimensional boundary layer. The findings are directly applicable to aircraft wings and nacelles that use surface heating for anti-icing protection. The potential effects of the non-adiabatic condition are particularly important for laminar-flow sections where even small perturbations can result in significantly degraded aerodynamic performance. The results of the study give new insight to the fundamental coupling between streamwise pressure gradient and surface heat flux in laminar and transitional boundary layers.

Heat transfer characteristics of hypersonic waveriders with an emphasis on leading edge effects

NASA Technical Reports Server (NTRS)

Vanmol, Denis O.; Anderson, John D., Jr.

1992-01-01

The present analysis of the heat-transfer characteristics of a family of viscous-optimized, 60 m-long waverider hypersonic vehicles gives attention to the transition from laminar to turbulent flow, and to how the transition affects aerodynamic heating distributions over the waverider surface. Two different constant-dynamic-pressure flight trajectories are considered, at 0.2 and 1.0 freestream atmospheres. For Mach numbers below 10, it is found that passive radiative cooling of the surface is sufficient. The degree of leading-edge bluntness required by aerodynamic heating constraints does not significantly degrade the aerodynamic performance of these waveriders.

A leading edge heating array and a flat surface heating array - operation, maintenance and repair manual

NASA Technical Reports Server (NTRS)

1975-01-01

A general description of the leading edge/flat surface heating array is presented along with its components, assembly instructions, installation instructions, operation procedures, maintenance instructions, repair procedures, schematics, spare parts lists, engineering drawings of the array, and functional acceptance test log sheets. The proper replacement of components, correct torque values, step-by-step maintenance instructions, and pretest checkouts are described.

Roughness Effects on the Formation of a Leading Edge Vortex

NASA Astrophysics Data System (ADS)

Elliott, Cassidy; Lang, Amy; Wahidi, Redha; Wilroy, Jacob

2016-11-01

Microscopic scales cover the wings of Monarch butterflies, creating a patterned surface. This patterning is an important natural flow control mechanism that is thought to delay the growth of the leading edge vortex (LEV) produced by the flapping motion of a wing. The increased skin friction caused by the scales leads to a weaker LEV being shed into the butterfly's wake, lessening drag and increasing flight efficiency. To test this theory, a plate of random roughness was designed in SolidWorks and printed on the Objet 30 Pro 3D printer. A 2x3x5 cubic foot tow tank was used to test the rough plate at Reynold's numbers of 1500, 3000, and 6000 (velocities of 8, 16, and 32 mm/s) at an angle of attack of 45 degrees. Particle Image Velocimetry (PIV) captured images of the LEV generated by the plate when towed upwards through the particle-seeded flow. Codes written in MatLab were used to automatically track and determine the strength of the LEV. Circulation values for the randomly-rough plate were then compared to the same values generated in a previous experiment that used a smooth plate and a grooved plate to determine the effect of the patterning on vortex development. Funding provided by NSF REU site Grant EEC 1358991 and CBET 1628600.

Material Modeling of Space Shuttle Leading Edge and External Tank Materials For Use in the Columbia Accident Investigation

NASA Technical Reports Server (NTRS)

Carney, Kelly; Melis, Matthew; Fasanella, Edwin L.; Lyle, Karen H.; Gabrys, Jonathan

2004-01-01

Upon the commencement of the analytical effort to characterize the impact dynamics and damage of the Space Shuttle Columbia leading edge due to External Tank insulating foam, the necessity of creating analytical descriptions of these materials became evident. To that end, material models were developed of the leading edge thermal protection system, Reinforced Carbon Carbon (RCC), and a low density polyurethane foam, BX-250. Challenges in modeling the RCC include its extreme brittleness, the differing behavior in compression and tension, and the anisotropic fabric layup. These effects were successfully included in LS-DYNA Material Model 58, *MAT_LAMINATED_ COMPOSITE_ FABRIC. The differing compression and tension behavior was modeled using the available damage parameters. Each fabric layer was given an integration point in the shell element, and was allowed to fail independently. Comparisons were made to static test data and coupon ballistic impact tests before being utilized in the full scale analysis. The foam's properties were typical of elastic automotive foams; and LS-DYNA Material Model 83, *MAT_FU_CHANG_FOAM, was successfully used to model its behavior. Material parameters defined included strain rate dependent stress-strain curves for both loading and un-loading, and for both compression and tension. This model was formulated with static test data and strain rate dependent test data, and was compared to ballistic impact tests on load-cell instrumented aluminum plates. These models were subsequently utilized in analysis of the Shuttle leading edge full scale ballistic impact tests, and are currently being used in the Return to Flight Space Shuttle re-certification effort.

Twin lead ballistic conductor based on nanoribbon edge transport

NASA Astrophysics Data System (ADS)

Konôpka, Martin; Dieška, Peter

2018-03-01

If a device like a graphene nanoribbon (GNR) has all its four corners attached to electric current leads, the device becomes a quantum junction through which two electrical circuits can interact. We study such system theoretically for stationary currents. The 4-point energy-dependent conductance matrix of the nanostructure and the classical resistors in the circuits are parameters of the model. The two bias voltages in the circuits are the control variables of the studied system while the electrochemical potentials at the device's terminals are non-trivially dependent on the voltages. For the special case of the linear-response regime analytical formulae for the operation of the coupled quantum-classical device are derived and applied. For higher bias voltages numerical solutions are obtained. The effects of non-equilibrium Fermi levels are captured using a recursive algorithm in which self-consistency between the electrochemical potentials and the currents is reached within few iterations. The developed approach allows to study scenarios ranging from independent circuits to strongly coupled ones. For the chosen model of the GNR with highly conductive zigzag edges we determine the regime in which the single device carries two almost independent currents.

Origin of long lifetime of band-edge charge carriers in organic-inorganic lead iodide perovskites.

PubMed

Chen, Tianran; Chen, Wei-Liang; Foley, Benjamin J; Lee, Jooseop; Ruff, Jacob P C; Ko, J Y Peter; Brown, Craig M; Harriger, Leland W; Zhang, Depei; Park, Changwon; Yoon, Mina; Chang, Yu-Ming; Choi, Joshua J; Lee, Seung-Hun

2017-07-18

Long carrier lifetime is what makes hybrid organic-inorganic perovskites high-performance photovoltaic materials. Several microscopic mechanisms behind the unusually long carrier lifetime have been proposed, such as formation of large polarons, Rashba effect, ferroelectric domains, and photon recycling. Here, we show that the screening of band-edge charge carriers by rotation of organic cation molecules can be a major contribution to the prolonged carrier lifetime. Our results reveal that the band-edge carrier lifetime increases when the system enters from a phase with lower rotational entropy to another phase with higher entropy. These results imply that the recombination of the photoexcited electrons and holes is suppressed by the screening, leading to the formation of polarons and thereby extending the lifetime. Thus, searching for organic-inorganic perovskites with high rotational entropy over a wide range of temperature may be a key to achieve superior solar cell performance.

Origin of long lifetime of band-edge charge carriers in organic–inorganic lead iodide perovskites

PubMed Central

Chen, Tianran; Chen, Wei-Liang; Foley, Benjamin J.; Lee, Jooseop; Ruff, Jacob P. C.; Ko, J. Y. Peter; Brown, Craig M.; Harriger, Leland W.; Zhang, Depei; Park, Changwon; Yoon, Mina; Chang, Yu-Ming; Choi, Joshua J.; Lee, Seung-Hun

2017-01-01

Long carrier lifetime is what makes hybrid organic–inorganic perovskites high-performance photovoltaic materials. Several microscopic mechanisms behind the unusually long carrier lifetime have been proposed, such as formation of large polarons, Rashba effect, ferroelectric domains, and photon recycling. Here, we show that the screening of band-edge charge carriers by rotation of organic cation molecules can be a major contribution to the prolonged carrier lifetime. Our results reveal that the band-edge carrier lifetime increases when the system enters from a phase with lower rotational entropy to another phase with higher entropy. These results imply that the recombination of the photoexcited electrons and holes is suppressed by the screening, leading to the formation of polarons and thereby extending the lifetime. Thus, searching for organic–inorganic perovskites with high rotational entropy over a wide range of temperature may be a key to achieve superior solar cell performance. PMID:28673975

DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev, , Dr.

2017-04-01

The flow over a 2D leading-edge flat plate is studied at Mach number Ma =(Uinf / \\setmn √{kBTinf / m}) in the range leading-edge flat plate boundary layer at high Mach number. Here, LT is the characteristic dimension, Uinfand Tinfare the free stream velocity and temperature, rhoinf is the free stream density, m is the molecular mass, muinf is the molecular viscosity based on the free stream temperature Tinf , and kB is the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.

DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev, , Dr.

2016-11-01

The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf /√{kBTinf / m }) in the range leading-edge flat plate boundary layer at high Mach number. Here, LT is the characteristic dimension, Uinf and Tinf are the free stream velocity and temperature, ρinf is the free stream density, mis the molecular mass, μinf is the molecular viscosity based on the free stream temperature Tinf, and kB is the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.

DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev, , Dr.

2017-01-01

The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf /√{kBTinf / m }) in the range leading-edge flat plate boundary layer at high Mach number. Here, LT is the characteristic dimension, Uinf and Tinf are the free stream velocity and temperature, rhoinf is the free stream density, m is the molecular mass, muinf is the molecular viscosity based on the free stream temperature Tinf , and kB is the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.

DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev

2016-10-01

The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf / {kBTinf /m}) in the range leading-edge flat plate boundary layer at high Mach number. Here, LT is the characteristic dimension, Uinf and Tinf are the free stream velocity and temperature, rhoinf is the free stream density, m is the molecular mass, muinf is the molecular viscosity based on the free stream temperature Tinf , and kB is the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.

DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev, , Dr.

The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf / ∖ sqrt{kBTinf / m})in the range leading-edge flat plate boundary layer at high Mach number. Here, LTis the characteristic dimension, Uinfand Tinfare the free stream velocity and temperature, rhoinf is the free stream density, mis the molecular mass, muinf is the molecular viscosity based on the free stream temperature Tinf , and kBis the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain. Indian Institute of Science Bangalore-560 012, India.

Influence of optimized leading-edge deflection and geometric anhedral on the low-speed aerodynamic characteristics of a low-aspect-ratio highly swept arrow-wing configuration. [langley 7 by 10 foot tunnel

NASA Technical Reports Server (NTRS)

Coe, P. L., Jr.; Huffman, J. K.

1979-01-01

An investigation conducted in the Langley 7 by 10 foot tunnel to determine the influence of an optimized leading-edge deflection on the low speed aerodynamic performance of a configuration with a low aspect ratio, highly swept wing. The sensitivity of the lateral stability derivative to geometric anhedral was also studied. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to spanwise variation of unwash, the resulting optimized leading edge was a smooth, continuously warped surface for which the deflection varied from 16 deg at the side of body to 50 deg at the wing tip. For the particular configuration studied, levels of leading-edge suction on the order of 90 percent were achieved. The results of tests conducted to determine the sensitivity of the lateral stability derivative to geometric anhedral indicate values which are in reasonable agreement with estimates provided by simple vortex-lattice theories.

First experimental assessment of RCS plume-flow field interaction on Hermes leading edge thruster configuration

NASA Astrophysics Data System (ADS)

Poertner, T.

1993-11-01

Glow discharge flow visualization experiments are demonstrated which have been performed to enable a first assessment of the HERMES 1.0 leading edge thruster configuration concerning interference between the thruster plumes of the reaction control system (RCS) and the surrounding flow field. The results of the flow visualization tests are presented in exemplary selected photographs. Additional Pitot pressure measurements support assumptions concerning interference induced pressure changes that may result from the observed significant flow field disturbances.

Effect of RANS-Type Turbulence Models on Adiabatic Film Cooling Effectiveness over a Scaled Up Gas Turbine Blade Leading Edge Surface

NASA Astrophysics Data System (ADS)

Yepuri, Giridhara Babu; Talanki Puttarangasetty, Ashok Babu; Kolke, Deepak Kumar; Jesuraj, Felix

2016-06-01

Increasing the gas turbine inlet temperature is one of the key technologies in raising gas turbine engine power output. Film cooling is one of the efficient cooling techniques to cool the hot section components of a gas turbine engines in turn the turbine inlet temperature can be increased. This study aims at investigating the effect of RANS-type turbulence models on adiabatic film cooling effectiveness over a scaled up gas turbine blade leading edge surfaces. For the evaluation, five different two equation RANS-type turbulent models have been taken in consideration, which are available in the ANSYS-Fluent. For this analysis, the gas turbine blade leading edge configuration is generated using Solid Works. The meshing is done using ANSYS-Workbench Mesh and ANSYS-Fluent is used as a solver to solve the flow field. The considered gas turbine blade leading edge model is having five rows of film cooling circular holes, one at stagnation line and the two each on either side of stagnation line at 30° and 60° respectively. Each row has the five holes with the hole diameter of 4 mm, pitch of 21 mm arranged in staggered manner and has the hole injection angle of 30° in span wise direction. The experiments are carried in a subsonic cascade tunnel facility at heat transfer lab of CSIR-National Aerospace Laboratory with a Reynolds number of 1,00,000 based on leading edge diameter. From the Computational Fluid Dynamics (CFD) evaluation it is found that K-ɛ Realizable model gives more acceptable results with the experimental values, compared to the other considered turbulence models for this type of geometries. Further the CFD evaluated results, using K-ɛ Realizable model at different blowing ratios are compared with the experimental results.

Aerothermodynamic Environment Definition for the Genesis Sample Return Capsule

NASA Technical Reports Server (NTRS)

Cheatwood, F. McNeil; Merski, N. Ronald, Jr.; Riley, Christopher J.; Mitcheltree, Robert A.

2001-01-01

NASA's Genesis sample return mission will be the first to return material from beyond the Earth-Moon system. NASA Langley Research Center supported this mission with aerothermodynamic analyses of the sample return capsule. This paper provides an overview of that effort. The capsule is attached through its forebody to the spacecraft bus. When the attachment is severed prior to Earth entry, forebody cavities remain. The presence of these cavities could dramatically increase the heating environment in their vicinity and downstream. A combination of computational fluid dynamics calculations and wind tunnel phosphor thermography tests were employed to address this issue. These results quantify the heating environment in and around the cavities, and were a factor in the decision to switch forebody heat shield materials. A transition map is developed which predicts that the flow aft of the penetrations will still be laminar at the peak heating point of the trajectory. As the vehicle continues along the trajectory to the peak dynamic pressure point, fully turbulent flow aft of the penetrations could occur. The integrated heat load calculations show that a heat shield sized to the stagnation point levels will be adequate for the predicted environment aft of the penetrations.

Carbon dioxide gas purification and analytical measurement for leading edge 193nm lithography

NASA Astrophysics Data System (ADS)

Riddle Vogt, Sarah; Landoni, Cristian; Applegarth, Chuck; Browning, Matt; Succi, Marco; Pirola, Simona; Macchi, Giorgio

2015-03-01

The use of purified carbon dioxide (CO2) has become a reality for leading edge 193 nm immersion lithography scanners. Traditionally, both dry and immersion 193 nm lithographic processes have constantly purged the optics stack with ultrahigh purity compressed dry air (UHPCDA). CO2 has been utilized for a similar purpose as UHPCDA. Airborne molecular contamniation (AMC) purification technologies and analytical measurement methods have been extensively developed to support the Lithography Tool Manufacturers purity requirements. This paper covers the analytical tests and characterizations carried out to assess impurity removal from 3.0 N CO2 (beverage grade) for its final utilization in 193 nm and EUV scanners.

Influence of a heated leading edge on boundary layer growth, stability, and transition

DOE Office of Scientific and Technical Information (OSTI.GOV)

Landrum, D.B.; Macha, J.M.

1987-06-01

This paper presents the results of a combined theoretical and experimental study of the influence of a heated leading edge on the growth, stability, and transition of a two-dimensional boundary layer. The findings are directly applicable to aircraft wings and nacelles that use surface heating for anti-icing protection. The potential effects of the non-adiabatic condition are particularly important for laminar-flow sections where even small perturbations can result in significantly degraded aerodynamic performance. The results of the study give new insight to the fundamental coupling between streamwise pressure gradient and surface heat flux in laminar and transitional boundary layers. 13 references.

Calibration of sonic valves for the laminar flow control, leading-edge flight test

NASA Technical Reports Server (NTRS)

Petley, D. H.; Alexander, W., Jr.; Wright, A. S., Jr.; Vallas, M.

1985-01-01

Sonic needle valves were calibrated to measure and control airflow in the suction system for the leading-edge flight test. The procedure and results for the calibration flow test of 4:41 flight valves are given. Mass-flow rates, which ranged from 0.001 to 0.012 lbm/sec, and maximum back pressure were measured for total temperatures from -30 F to 75 F and total pressures from 120 to 540 psf. Correlating equations are obtained for mass-flow rate as a function of total pressure, total temperature, and valve opening length. The most important aspect of flow measurement and control is found to be the measurement of valve opening length.

AERO2S - SUBSONIC AERODYNAMIC ANALYSIS OF WINGS WITH LEADING- AND TRAILING-EDGE FLAPS IN COMBINATION WITH CANARD OR HORIZONTAL TAIL SURFACES (IBM PC VERSION)

NASA Technical Reports Server (NTRS)

Carlson, H. W.

1994-01-01

This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is

Full-scale wind-tunnel investigation of the effects of wing leading-edge modifications on the high angle-of-attack aerodynamic characteristics of a low-wing general aviation airplane

NASA Technical Reports Server (NTRS)

Johnson, J. L., Jr.; Newsom, W. A.; Satran, D. R.

1980-01-01

The paper presents the results of a recent investigation to determine the effects of wing leading-edge modifications on the high angle-of-attack aerodynamic characteristics of a low-wing general aviation airplane in the Langley Full-Scale Wind Tunnel. The investigation was conducted to provide aerodynamic information for correlation and analysis of flight-test results obtained for the configuration. The wind-tunnel investigation consisted of force and moment measurements, wing pressure measurements, flow surveys, and flow visualization studies utilizing a tuft grid, smoke and nonintrusive mini-tufts which were illuminated by ultra-violet light. In addition to the tunnel scale system which measured overall forces and moments, the model was equipped with an auxiliary strain-gage balance within the left wing panel to measure lift and drag forces on the outer wing panel independent of the tunnel scale system. The leading-edge modifications studied included partial- and full-span leading-edge droop arrangements as well as leading-edge slats.

Preliminary results of the calculated and experimental studies of the basic aerothermodynamic parameters of the ExoMars landing module

NASA Astrophysics Data System (ADS)

Finchenko, V. S.; Ivankov, A. A.; Shmatov, S. I.; Mordvinkin, A. S.

2015-12-01

The article presents the initial data for the ExoMars landing module aerothermodynamic calculations, used calculation methods, the calculation results of aerodynamic characteristics of the landing module shape and structural parameters of thermal protection selected during the conceptual design phase. Also, the test results of the destruction of the thermal protection material and comparison of the basic characteristics of the landing module with a front shield in the form of a cone and a spherical segment are presented.

Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils

NASA Technical Reports Server (NTRS)

Lindsey, Walter F; Landrum, Emma Jean

1958-01-01

Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.

Multiple leading edge vortices of unexpected strength in freely flying hawkmoth

PubMed Central

Johansson, L. Christoffer; Engel, Sophia; Kelber, Almut; Heerenbrink, Marco Klein; Hedenström, Anders

2013-01-01

The Leading Edge Vortex (LEV) is a universal mechanism enhancing lift in flying organisms. LEVs, generally illustrated as a single vortex attached to the wing throughout the downstroke, have not been studied quantitatively in freely flying insects. Previous findings are either qualitative or from flappers and tethered insects. We measure the flow above the wing of freely flying hawkmoths and find multiple simultaneous LEVs of varying strength and structure along the wingspan. At the inner wing there is a single, attached LEV, while at mid wing there are multiple LEVs, and towards the wingtip flow separates. At mid wing the LEV circulation is ~40% higher than in the wake, implying that the circulation unrelated to the LEV may reduce lift. The strong and complex LEV suggests relatively high flight power in hawmoths. The variable LEV structure may result in variable force production, influencing flight control in the animals. PMID:24253180

"Partners in Science": A Model Cooperative Program Introducing High School Teachers and Students to Leading-Edge Pharmaceutical Science

ERIC Educational Resources Information Center

Woska, Joseph R., Jr.; Collins, Danielle M.; Canney, Brian J.; Arcario, Erin L.; Reilly, Patricia L.

2005-01-01

"Partners in Science" is a cooperative program between Boehringer Ingelheim Pharmaceuticals, Inc. and area high schools in the community surrounding our Connecticut campus. It is a two-phase program that introduces high school students and teachers to the world of drug discovery and leading-edge pharmaceutical research. Phase 1 involves…

Turbine Vane External Heat Transfer. Volume 1: Analytical and Experimental Evaluation of Surface Heat Transfer Distributions with Leading Edge Showerhead Film Cooling

NASA Technical Reports Server (NTRS)

Turner, E. R.; Wilson, M. D.; Hylton, L. D.; Kaufman, R. M.

1985-01-01

Progress in predictive design capabilities for external heat transfer to turbine vanes was summarized. A two dimensional linear cascade (previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils) was used to examine the effect of leading edge shower head film cooling on downstream heat transfer. The data were used to develop and evaluate analytical models. Modifications to the two dimensional boundary layer model are described. The results were used to formulate and test an effective viscosity model capable of predicting heat transfer phenomena downstream of the leading edge film cooling array on both the suction and pressure surfaces, with and without mass injection.

Characteristics of surface roughness associated with leading edge ice accretion

NASA Technical Reports Server (NTRS)

Shin, Jaiwon

1994-01-01

Detailed size measurements of surface roughness associated with leading edge ice accretions are presented to provide information on characteristics of roughness and trends of roughness development with various icing parameters. Data was obtained from icing tests conducted in the Icing Research Tunnel (IRT) at NASA Lewis Research Center (LeRC) using a NACA 0012 airfoil. Measurements include diameters, heights, and spacing of roughness elements along with chordwise icing limits. Results confirm the existence of smooth and rough ice zones and that the boundary between the two zones (surface roughness transition region) moves upstream towards stagnation region with time. The height of roughness grows as the air temperature and the liquid water content increase, however, the airspeed has little effect on the roughness height. Results also show that the roughness in the surface roughness transition region grows during a very early stage of accretion but reaches a critical height and then remains fairly constant. Results also indicate that a uniformly distributed roughness model is only valid at a very initial stage of the ice accretion process.

DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

NASA Astrophysics Data System (ADS)

Pradhan, Sahadev

2016-09-01

The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf /√{kBTinf / m }) in the range leading-edge flat plate boundary layer at high Mach number. Here, LT is the characteristic dimension, Uinf and Tinf are the free stream velocity and temperature, rhoinf is the free stream density, m is the molecular mass, muinfis the molecular viscosity based on the free stream temperature Tinf , and kB is the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip ((Pradhan and Kumaran, J. Fluid Mech-2011); (Kumaran and Pradhan, J. Fluid Mech-2014)) at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.

Transport Mechanisms Governing initial Leading-Edge Vortex Development on a Pitching Wing

NASA Astrophysics Data System (ADS)

Wabick, Kevin; Berdon, Randall; Buchholz, James; Johnson, Kyle; Thurow, Brian

2017-11-01

The formation and evolution of Leading Edge Vortices (LEVs) are ubiquitous in natural fliers and maneuvering wings, and have a profound impact on aerodynamic loads. The formation of an LEV is experimentally investigated on a pitching flat-plate wing of aspect-ratio 2, and dimensionless pitch rates of k = Ωc / 2 U of 0.1, 0.2, and 0.5, at a Reynolds number of 104. The sources and sinks of vorticity that contribute to the growth and evolution of the LEV are investigated at spanwise regions of interest, and their relative balance is compared to other wing kinematics, and the case of a two-dimensional pitching wing. This work is supported by the Air Force Office of Scientific Research (Grant Number FA9550-16-1-0107, Dr. Douglas Smith, program manager).

AMELIA CESTOL Test: Acoustic Characteristics of Circulation Control Wing with Leading- and Trailing-Edge Slot Blowing

NASA Technical Reports Server (NTRS)

Horne, William C.; Burnside, Nathan J.

2013-01-01

The AMELIA Cruise-Efficient Short Take-off and Landing (CESTOL) configuration concept was developed to meet future requirements of reduced field length, noise, and fuel burn by researchers at Cal Poly, San Luis Obispo and Georgia Tech Research Institute under sponsorship by the NASA Fundamental Aeronautics Program (FAP), Subsonic Fixed Wing Project. The novel configuration includes leading- and trailing-edge circulation control wing (CCW), over-wing podded turbine propulsion simulation (TPS). Extensive aerodynamic measurements of forces, surfaces pressures, and wing surface skin friction measurements were recently measured over a wide range of test conditions in the Arnold Engineering Development Center(AEDC) National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Ft Wind Tunnel. Acoustic measurements of the model were also acquired for each configuration with 7 fixed microphones on a line under the left wing, and with a 48-element, 40-inch diameter phased microphone array under the right wing. This presentation will discuss acoustic characteristics of the CCW system for a variety of tunnel speeds (0 to 120 kts), model configurations (leading edge(LE) and/or trailing-edge(TE) slot blowing, and orientations (incidence and yaw) based on acoustic measurements acquired concurrently with the aerodynamic measurements. The flow coefficient, Cmu= mVSLOT/qSW varied from 0 to 0.88 at 40 kts, and from 0 to 0.15 at 120 kts. Here m is the slot mass flow rate, VSLOT is the slot exit velocity, q is dynamic pressure, and SW is wing surface area. Directivities at selected 1/3 octave bands will be compared with comparable measurements of a 2-D wing at GTRI, as will as microphone array near-field measurements of the right wing at maximum flow rate. The presentation will include discussion of acoustic sensor calibrations as well as characterization of the wind tunnel background noise environment.

Rotational accelerations stabilize leading edge vortices on revolving fly wings.

PubMed

Lentink, David; Dickinson, Michael H

2009-08-01

The aerodynamic performance of hovering insects is largely explained by the presence of a stably attached leading edge vortex (LEV) on top of their wings. Although LEVs have been visualized on real, physically modeled, and simulated insects, the physical mechanisms responsible for their stability are poorly understood. To gain fundamental insight into LEV stability on flapping fly wings we expressed the Navier-Stokes equations in a rotating frame of reference attached to the wing's surface. Using these equations we show that LEV dynamics on flapping wings are governed by three terms: angular, centripetal and Coriolis acceleration. Our analysis for hovering conditions shows that angular acceleration is proportional to the inverse of dimensionless stroke amplitude, whereas Coriolis and centripetal acceleration are proportional to the inverse of the Rossby number. Using a dynamically scaled robot model of a flapping fruit fly wing to systematically vary these dimensionless numbers, we determined which of the three accelerations mediate LEV stability. Our force measurements and flow visualizations indicate that the LEV is stabilized by the ;quasi-steady' centripetal and Coriolis accelerations that are present at low Rossby number and result from the propeller-like sweep of the wing. In contrast, the unsteady angular acceleration that results from the back and forth motion of a flapping wing does not appear to play a role in the stable attachment of the LEV. Angular acceleration is, however, critical for LEV integrity as we found it can mediate LEV spiral bursting, a high Reynolds number effect. Our analysis and experiments further suggest that the mechanism responsible for LEV stability is not dependent on Reynolds number, at least over the range most relevant for insect flight (100

The Influence of Projectile Trajectory Angle on the Simulated Impact Response of a Shuttle Leading Edge Wing Panel

NASA Technical Reports Server (NTRS)

Spellman, Regina L.; Jones, Lisa E.; Lyle, Karen H.; Jackson, Karen E.; Fasanella, Edwin L.

2005-01-01

In support of recommendations by the Columbia Accident Investigation Board, a team has been studying the effect of debris impacting the reinforced carbon-carbon panels of the shuttle leading edge. The objective of this study was to examine the effect of varying parameters of the debris trajectory on the damage tolerance. Impacts at the upper and lower surface and the apex of the leading edge were examined. For each location, trajectory variances included both the alpha and beta directions. The results of the analysis indicated in all cases the beta sweep decreased the amount of damage to the panel. The increases in alpha resulted in a significant increase in damage to the RCC panel. In particular, for the lower surface, where the alpha can increase by 10 degrees, there was a nearly 40% increase in the impulse. As a result, it is recommended that for future analyses, a 10 degree offset in alpha from the nominal trajectory is included for impacts on the lower surface. It is also recommended to assume a straight aft, or zero beta, trajectory for a more conservative analysis.

Lift augmentation on a delta wing via leading edge fences and the Gurney flap. M.S. Thesis

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.

1992-01-01

Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Performance of laminar-flow leading-edge test articles in cloud encounters

NASA Technical Reports Server (NTRS)

Davis, Richard E.; Maddalon, Dal V.; Wagner, Richard D.

1987-01-01

An extensive data bank of concurrent measurements of laminar flow (LF), particle concentration, and aircraft charging state was gathered for the first time. From this data bank, 13 flights in the simulated airline service (SAS) portion were analyzed to date. A total of 6.86 hours of data at one-second resolution were analyzed. An extensive statistical analysis, for both leading-edge test articles, shows that there is a significant effect of cloud and haze particles on the extent of laminar flow obtained. Approximately 93 percent of data points simulating LFC flight were obtained in clear air conditions; approximately 7 percent were obtained in cloud and haze. These percentages are consistent with earlier USAF and NASA estimates and results. The Hall laminar flow loss criteria was verified qualitatively. Larger particles and higher particle concentrations have a more marked effect on LF than do small particles. A particle spectrometer of a charging patch are both acceptable as diagnostic indicators of the presence of particles detrimental to laminar flow.

Roughness Effects on the Formation of a Leading Edge Vortex

NASA Astrophysics Data System (ADS)

Elliott, Cassidy; Lang, Amy; Wahidi, Redha; Wilroy, Jacob

2017-11-01

Microscopic scales cover the wings of Monarch butterflies, creating a patterned surface that acts as a natural energy capture mechanism. This patterning is thought to delay the growth of the leading edge vortex (LEV) produced by the flapping motion of a wing. Increased skin friction caused by the scales leads to a weaker LEV being shed into the butterfly's wake, lessening drag and increasing flight efficiency. To test how this roughness effects LEV formation, a plate of random roughness was designed in SolidWorks and printed on the Objet 30 Pro 3D printer. A 2x3x5 cubic foot tow tank was used to test the rough plate at Reynold's numbers of 1500, 3000, and 6000 (velocities of 8, 16, and 32 mm/s) at an angle of attack of 45 degrees. Images were captured of the LEV generated when the plate was towed upwards through the particle-seeded flow. These images were used to determine the XY velocity of the particles using a technique called Digital Particle Image Velocimetry (DPIV). Codes written in MATLAB were used to track and measure the strength of the LEV. Circulation values for the randomly-rough plate were then compared to the same values generated in a previous experiment that used a smooth plate and a grooved plate to determine the effect of the patterning on vortex development. Funding for this research project was provided by the National Science Foundation under the Research Experience for Undergraduates (REU) program (REU Supplement CBET 1628600 under CBET 1335848).

Wind-tunnel tests on model wing with Fowler flap and specially developed leading-edge slot

NASA Technical Reports Server (NTRS)

Weick, Fred E; Platt, Robert C

1933-01-01

An investigation was made in the NACA 7 by 10 foot wind tunnel to find the increase in maximum lift coefficient which could be obtained by providing a model wing with both a Fowler trailing-edge extension flap and a Handley Page type leading-edge slot. A conventional Handley page slot proportioned to operate on the plain wing without a flap gave but a slight increase with the flap; so a special form of slot was developed to work more effectively with the flap. With the best combined arrangement the maximum lift coefficient based on the original area was increased from 3.17, for the Fowler wing, to 3.62. The minimum drag coefficient with both devices retracted was increased in approximately the same proportion. Tests were also made with the special-type slot on the plain wing without the flap. The special slot, used either with or without the Fowler flap, gave definitely higher values of the maximum lift coefficient than the slots of conventional form, with an increase of the same order in the minimum drag coefficient.

Modeling Shock Train Leading Edge Detection in Dual-Mode Scramjets

NASA Astrophysics Data System (ADS)

Ladeinde, Foluso; Lou, Zhipeng; Li, Wenhai

2016-11-01

The objective of this study is to accurately model the detection of shock train leading edge (STLE) in dual-mode scramjet (DMSJ) engines intended for hypersonic flight in air-breathing propulsion systems. The associated vehicles have applications in military warfare and intelligence, and there is commercial interest as well. Shock trains are of interest because they play a significant role in the inability of a DMSJ engine to develop the required propulsive force. The experimental approach to STLE detection has received some attention; as have numerical calculations. However, virtually all of the numerical work focus on mechanically- (i.e., pressure-) generated shock trains, which are much easier to model relative to the phenomenon in the real system where the shock trains are generated by combustion. A focus on combustion, as in the present studies, enables the investigation of the effects of equivalence ratio, which, together with the Mach number, constitutes an important parameter determining mode transition. The various numerical approaches implemented in our work will be reported, with result comparisons to experimental data. The development of an STLE detection procedure in an a priori manner will also be discussed.

Large-Scale Wind-Tunnel Tests and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat

NASA Technical Reports Server (NTRS)

Hickey, David H.; Aoyagi, Kiyoshi

1960-01-01

A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.

Effects of continuum breakdown on hypersonic aerothermodynamics for reacting flow

NASA Astrophysics Data System (ADS)

Holman, Timothy D.; Boyd, Iain D.

2011-02-01

This study investigates the effects of continuum breakdown on the surface aerothermodynamic properties (pressure, stress, and heat transfer rate) of a sphere in a Mach 25 flow of reacting air in regimes varying from continuum to a rarefied gas. Results are generated using both continuum [computational fluid dynamics (CFD)] and particle [direct simulation Monte Carlo (DSMC)] approaches. The DSMC method utilizes a chemistry model that calculates the backward rates from an equilibrium constant. A preferential dissociation model is modified in the CFD method to better compare with the vibrationally favored dissociation model that is utilized in the DSMC method. Tests of these models are performed to confirm their validity and to compare the chemistry models in both numerical methods. This study examines the effect of reacting air flow on continuum breakdown and the surface properties of the sphere. As the global Knudsen number increases, the amount of continuum breakdown in the flow and on the surface increases. This increase in continuum breakdown significantly affects the surface properties, causing an increase in the differences between CFD and DSMC. Explanations are provided for the trends observed.

Nonequilibrium effects on the aerothermodynamics of transatmospheric and aerobraking vehicles

NASA Technical Reports Server (NTRS)

Hassan, Basil; Candler, Graham V.

1993-01-01

A 3D CFD algorithm is used to study the effect of thermal and chemical nonequilibrium on slender and blunt body aerothermodynamics. Both perfect gas and reacting gas air models are used to compute the flow over a generic transatmospheric vehicle and a proposed lunar transfer vehicle. The reacting air is characterized by a translational-rotational temperature and a vibrational-electron-electronic temperature and includes eight chemical species. The effects of chemical reaction, vibrational excitation, and ionization on lift-to-drag ratio and trim angle are investigated. Results for the NASA Ames All-body Configuration show a significant difference in center of gravity location for a reacting gas flight case when compared to a perfect gas wind tunnel case at the same Mach number, Reynolds number, and angle of attack. For the same center of gravity location, the wind tunnel model trims at lower angle of attack than the full-scale flight case. Nonionized and ionized results for a proposed lunar transfer vehicle compare well to computational results obtained from a previously validated reacting gas algorithm. Under the conditions investigated, effects of weak ionization on the heat transfer and aerodynamic coefficients were minimal.

Myosin II activity is required for functional leading-edge cells and closure of epidermal sheets in fish skin ex vivo.

PubMed

Morita, Toshiyuki; Tsuchiya, Akiko; Sugimoto, Masazumi

2011-09-01

Re-epithelialization in skin wound healing is a process in which epidermal sheets grow and close the wound. Although the actin-myosin system is thought to have a pivotal role in re-epithelialization, its role is not clear. In fish skin, re-epithelialization occurs around 500 μm/h and is 50 times faster than in mammalian skin. We had previously reported that leading-edge cells of the epidermal outgrowth have both polarized large lamellipodia and "purse string"-like actin filament cables in the scale-skin culture system of medaka fish, Oryzias latipes (Cell Tissue Res, 2007). The actin purse-string (APS) is a supracellular contractile machinery in which adherens junctions (AJs) link intracellular myosin II-including actin cables between neighboring cells. In this study, we developed a modified "face-to-face" scale-skin culture system as an ex vivo model to study epidermal wound healing, and examined the role of the actin-myosin system in the rapid re-epithelialization using a myosin II ATPase inhibitor, blebbistatin. A low level of blebbistatin suppressed the formation of APS and induced the dissociation of keratocytes from the leading edge without attenuating the growth of the epidermal sheet or the migration rate of solitary keratocytes. AJs in the superficial layer showed no obvious changes elicited by blebbistatin. However, two epidermal sheets without APSs did not make a closure with each other, which was confirmed by inhibiting the connecting AJs between the superficial layers. These results suggest that myosin II activity is required for functional leading-edge cells and for epidermal closure.

Hypersonic Airbreathing Propulsion: An Aerodynamics, Aerothermodynamics, and Acoustics Competency White Paper

NASA Technical Reports Server (NTRS)

Drummond, J. Philip; Cockrell, Charles E., Jr.; Pellett, Gerald L.; Diskin, Glenn S.; Auslender, Aaron H.; Exton, Reginald J.; Guy, R. Wayne; Hoppe, John C.; Puster, Richard L.; Rogers, R. Clayton

2002-01-01

This White Paper examines the current state of Hypersonic Airbreathing Propulsion at the NASA Langley Research Center and the factors influencing this area of work and its personnel. Using this knowledge, the paper explores beyond the present day and suggests future directions and strategies for the field. Broad views are first taken regarding potential missions and applications of hypersonic propulsion. Then, candidate propulsion systems that may be applicable to these missions are suggested and discussed. Design tools and experimental techniques for developing these propulsion systems are then described, and approaches for applying them in the design process are considered. In each case, current strategies are reviewed and future approaches that may improve the techniques are considered. Finally, the paper concentrates on the needs to be addressed in each of these areas to take advantage of the opportunities that lay ahead for both the NASA Langley Research Center and the Aerodynamic Aerothermodynamic, and Aeroacoustics Competency. Recommendations are then provided so that the goals set forth in the paper may be achieved.

Aerothermodynamic Design of the Mars Science Laboratory Backshell and Parachute Cone

NASA Technical Reports Server (NTRS)

Edquist, Karl T.; Dyakonov, Artem A.; Wright, Michael J.; Tang, Chun Y.

2009-01-01

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule backshell and parachute cone. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux) design entry trajectories from a 2009 launch. Transient interference effects from reaction control system thruster plumes were included in the design environments when necessary. The limiting backshell design heating conditions of 6.3 W/sq cm for heat flux and 377 J/sq cm for total heat load are not influenced by thruster firings. Similarly, the thrusters do not affect the parachute cover lid design environments (13 W/sq cm and 499 J/sq cm). If thruster jet firings occur near peak dynamic pressure, they will augment the design environments at the interface between the backshell and parachute cone (7 W/sq cm and 174 J/sq cm). Localized heat fluxes are higher near the thruster fairing during jet firings, but these areas did not require additional thermal protection material. Finally, heating bump factors were developed for antenna radomes on the parachute cone

Timing discriminator using leading-edge extrapolation

DOEpatents

Gottschalk, Bernard

1983-01-01

A discriminator circuit to recover timing information from slow-rising pulses by means of an output trailing edge, a fixed time after the starting corner of the input pulse, which is nearly independent of risetime and threshold setting. This apparatus comprises means for comparing pulses with a threshold voltage; a capacitor to be charged at a certain rate when the input signal is one-third threshold voltage, and at a lower rate when the input signal is two-thirds threshold voltage; current-generating means for charging the capacitor; means for comparing voltage capacitor with a bias voltage; a flip-flop to be set when the input pulse reaches threshold voltage and reset when capacitor voltage reaches the bias voltage; and a clamping means for discharging the capacitor when the input signal returns below one-third threshold voltage.

Timing discriminator using leading-edge extrapolation

DOEpatents

Gottschalk, B.

1981-07-30

A discriminator circuit to recover timing information from slow-rising pulses by means of an output trailing edge, a fixed time after the starting corner of the input pulse, which is nearly independent of risetime and threshold setting is described. This apparatus comprises means for comparing pulses with a threshold voltage; a capacitor to be charged at a certain rate when the input signal is one-third threshold voltage, and at a lower rate when the input signal is two-thirds threshold voltage; current-generating means for charging the capacitor; means for comparing voltage capacitor with a bias voltage; a flip-flop to be set when the input pulse reaches threshold voltage and reset when capacitor voltage reaches the bias voltage; and a clamping means for discharging the capacitor when the input signal returns below one-third threshold voltage.

Theoretical effect of modifications to the upper surface of two NACA airfoils using smooth polynomial additional thickness distributions which emphasize leading edge profile and which vary quadratically at the trailing edge. [using flow equations and a CDC 7600 computer

NASA Technical Reports Server (NTRS)

Merz, A. W.; Hague, D. S.

1975-01-01

An investigation was conducted on a CDC 7600 digital computer to determine the effects of additional thickness distributions to the upper surface of the NACA 64-206 and 64 sub 1 - 212 airfoils. The additional thickness distribution had the form of a continuous mathematical function which disappears at both the leading edge and the trailing edge. The function behaves as a polynomial of order epsilon sub 1 at the leading edge, and a polynomial of order epsilon sub 2 at the trailing edge. Epsilon sub 2 is a constant and epsilon sub 1 is varied over a range of practical interest. The magnitude of the additional thickness, y, is a second input parameter, and the effect of varying epsilon sub 1 and y on the aerodynamic performance of the airfoil was investigated. Results were obtained at a Mach number of 0.2 with an angle-of-attack of 6 degrees on the basic airfoils, and all calculations employ the full potential flow equations for two dimensional flow. The relaxation method of Jameson was employed for solution of the potential flow equations.

Effects of wing leading-edge flap deflections on subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing

NASA Technical Reports Server (NTRS)

Henderson, W. P.

1978-01-01

An investigation was conducted to determine the effects of wing leading-edge flap deflections on the subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing. The tests were conducted at Mach numbers from 0.40 to 0.85, corresponding to Reynolds numbers (based on wing mean geometric chord) of 2.37 x 1,000,000 to 4.59 x 1,000,000 and at angles of attack from -3 deg to 22 deg. The configurations under study included a wing-fuselage configuration and a wing-fuselage-strake configuration. Each configuration had multisegmented, constant-chord leading-edge flaps which could be deflected independently or in various combinations.

The Columbia River--on the Leading Edge

NASA Astrophysics Data System (ADS)

O'Connor, J. E.

2005-05-01

On the leading edge of the North American plate, the Columbia River is the largest of the world's 40 or so rivers with drainage areas greater than 500,000 square kilometers to drain toward a convergent plate boundary. This unique setting results in a unique continental river basin; marked by episodic and cataclysmic geologic disturbance, but also famously fecund with perhaps 10 to 16 million salmon historically spawning in its waters each year. Now transformed by dams, transportation infrastructure, dikes and diversions, the Columbia River presents an expensive conundrum for management of its many values. Inclusion of river ecology and geomorphology in discussions of river management is generally limited to observations of the last 200 years-a time period of little natural disturbance and low sediment transport. However, consideration of longer timescales provides additional perspective of historical ecologic and geomorphic conditions. Only 230 km from its mouth, the Columbia River bisects the volcanic arc of the Cascade Range, forming the Columbia River Gorge. Cenozoic lava flows have blocked the river, forcing diversions and new canyon cutting. Holocene eruptions of Mount Mazama (Crater Lake), Mount Hood, Mount St. Helens, and Mount Rainier have shed immense quantities of sediment into the lower Columbia River, forming a large percentage of the Holocene sediment transported through the lower river. Quaternary landslides, perhaps triggered by great earthquakes, have descended from the 1000-m-high gorge walls, also blocking and diverting the river, one as recently as 550 years ago. These geologic disturbances, mostly outside the realm of historical observation and operating at timescales of 100s to 1000s of years in the gorge and elsewhere, have clearly affected basin geomorphology, riverine ecology, and past and present cultural utilization of river resources. The historic productivity of the river, however, hints at extraordinary resilience (and perhaps

Effects of Wing Leading Edge Penetration with Venting and Exhaust Flow from Wheel Well at Mach 24 in Flight

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.

2003-01-01

A baseline solution for CFD Point 1 (Mach 24) in the STS-107 accident investigation was modified to include effects of: (1) holes through the leading edge into a vented cavity; and (2) a scarfed, conical nozzle directed toward the centerline of the vehicle from the forward, inboard corner of the landing gear door. The simulations were generated relatively quickly and early in the investigation because simplifications were made to the leading edge cavity geometry and an existing utility to merge scarfed nozzle grid domains with structured baseline external domains was implemented. These simplifications in the breach simulations enabled: (1) a very quick grid generation procedure; and (2) high fidelity corroboration of jet physics with internal surface impingements ensuing from a breach through the leading edge, fully coupled to the external shock layer flow at flight conditions. These simulations provided early evidence that the flow through a two-inch diameter (or larger) breach enters the cavity with significant retention of external flow directionality. A normal jet directed into the cavity was not an appropriate model for these conditions at CFD Point 1 (Mach 24). The breach diameters were of the same order or larger than the local, external boundary-layer thickness. High impingement heating and pressures on the downstream lip of the breach were computed. It is likely that hole shape would evolve as a slot cut in the direction of the external streamlines. In the case of the six-inch diameter breach the boundary layer is fully ingested. The intent of externally directed jet simulations in the second scenario was to approximately model aerodynamic effects of a relatively large internal wing pressure, fueled by combusting aluminum, which deforms the corner of the landing gear door and directs a jet across the windside surface. These jet interactions, in and of themselves, were not sufficiently large to explain observed aerodynamic behavior.

Fracture Mechanics Analyses of Reinforced Carbon-Carbon Wing-Leading-Edge Panels

NASA Technical Reports Server (NTRS)

Raju, Ivatury S.; Phillips, Dawn R.; Knight, Norman F., Jr.; Song, Kyongchan

2010-01-01

Fracture mechanics analyses of subsurface defects within the joggle regions of the Space Shuttle wing-leading-edge RCC panels are performed. A 2D plane strain idealized joggle finite element model is developed to study the fracture behavior of the panels for three distinct loading conditions - lift-off and ascent, on-orbit, and entry. For lift-off and ascent, an estimated bounding aerodynamic pressure load is used for the analyses, while for on-orbit and entry, thermo-mechanical analyses are performed using the extreme cold and hot temperatures experienced by the panels. In addition, a best estimate for the material stress-free temperature is used in the thermo-mechanical analyses. In the finite element models, the substrate and coating are modeled separately as two distinct materials. Subsurface defects are introduced at the coating-substrate interface and within the substrate. The objective of the fracture mechanics analyses is to evaluate the defect driving forces, which are characterized by the strain energy release rates, and determine if defects can become unstable for each of the loading conditions.

Computation of leading edge film cooling from a CONSOLE geometry (CONverging Slot hOLE)

NASA Astrophysics Data System (ADS)

Guelailia, A.; Khorsi, A.; Hamidou, M. K.

2016-01-01

The aim of this study is to investigate the effect of mass flow rate on film cooling effectiveness and heat transfer over a gas turbine rotor blade with three staggered rows of shower-head holes which are inclined at 30° to the spanwise direction, and are normal to the streamwise direction on the blade. To improve film cooling effectiveness, the standard cylindrical holes, located on the leading edge region, are replaced with the converging slot holes (console). The ANSYS CFX has been used for this computational simulation. The turbulence is approximated by a k-ɛ model. Detailed film effectiveness distributions are presented for different mass flow rate. The numerical results are compared with experimental data.

Smoothed Two-Dimensional Edges for Laminar Flow

NASA Technical Reports Server (NTRS)

Holmes, B. J.; Liu, C. H.; Martin, G. L.; Domack, C. S.; Obara, C. J.; Hassan, A.; Gunzburger, M. D.; Nicolaides, R. A.

1986-01-01

New concept allows passive method for installing flaps, slats, iceprotection equipment, and other leading-edge devices on natural-laminar-flow (NLF) wings without causing loss of laminar flow. Two-dimensional roughness elements in laminar boundary layers strategically shaped to increase critical (allowable) height of roughness. Facilitates installation of leading-edge devices by practical manufacturing methods.

Mechanism of Pressure-Induced Phase Transitions, Amorphization, and Absorption-Edge Shift in Photovoltaic Methylammonium Lead Iodide.

PubMed

Szafrański, Marek; Katrusiak, Andrzej

2016-09-01

Our single-crystal X-ray diffraction study of methylammonium lead triiodide, MAPbI3, provides the first comprehensive structural information on the tetragonal phase II in the pressure range to 0.35 GPa, on the cubic phase IV stable between 0.35 and 2.5 GPa, and on the isostructural cubic phase V observed above 2.5 GPa, which undergoes a gradual amorphization. The optical absorption study confirms that up to 0.35 GPa, the absorption edge of MAPbI3 is red-shifted, allowing an extension of spectral absorption. The transitions to phases IV and V are associated with the abrupt blue shifts of the absorption edge. The strong increase of the energy gap in phase V result in a spectacular color change of the crystal from black to red around 3.5 GPa. The optical changes have been correlated with the pressure-induced strain of the MAPbI3 inorganic framework and its frustration, triggered by methylammonium cations trapped at random orientations in the squeezed voids.

Influence of leading edge bluntness on hypersonic flow in a generic internal-compression inlet

NASA Astrophysics Data System (ADS)

Borovoy, V.; Egorov, I.; Mosharov, V.; Radchenko, V.; Skuratov, A.; Struminskaya, I.

2015-06-01

Flow and heat transfer inside a generic inlet are investigated experimentally. The cross section of the inlet is rectangular. The inlet is installed on a flat plat at a significant distance from the leading edge. The experiments are performed in TsAGI wind tunnel UT-1M working in the Ludwieg tube mode at Mach number M∞ = 5 and Reynolds numbers (based on the plate length L = 320 mm) Re∞L = 23 · 106 and 13 · 106. Steady flow duration is 40 ms. Optical panoramic methods are used for investigation of flow outside and inside the inlet as well. For this purpose, the cowl and one of two compressing wedges are made of a transparent material. Heat flux distribution is measured by thin luminescent Temperature Sensitive Paint (TSP). Surface flow and shear stress visualization is performed by viscous oil containing luminophor particles. The investigation shows that at high contraction ratio of the inlet, an increase of plate or cowl bluntness to some critical value leads to sudden change of the flow structure.

Power spectrum weighted edge analysis for straight edge detection in images

NASA Astrophysics Data System (ADS)

Karvir, Hrishikesh V.; Skipper, Julie A.

2007-04-01

Most man-made objects provide characteristic straight line edges and, therefore, edge extraction is a commonly used target detection tool. However, noisy images often yield broken edges that lead to missed detections, and extraneous edges that may contribute to false target detections. We present a sliding-block approach for target detection using weighted power spectral analysis. In general, straight line edges appearing at a given frequency are represented as a peak in the Fourier domain at a radius corresponding to that frequency, and a direction corresponding to the orientation of the edges in the spatial domain. Knowing the edge width and spacing between the edges, a band-pass filter is designed to extract the Fourier peaks corresponding to the target edges and suppress image noise. These peaks are then detected by amplitude thresholding. The frequency band width and the subsequent spatial filter mask size are variable parameters to facilitate detection of target objects of different sizes under known imaging geometries. Many military objects, such as trucks, tanks and missile launchers, produce definite signatures with parallel lines and the algorithm proves to be ideal for detecting such objects. Moreover, shadow-casting objects generally provide sharp edges and are readily detected. The block operation procedure offers advantages of significant reduction in noise influence, improved edge detection, faster processing speed and versatility to detect diverse objects of different sizes in the image. With Scud missile launcher replicas as target objects, the method has been successfully tested on terrain board test images under different backgrounds, illumination and imaging geometries with cameras of differing spatial resolution and bit-depth.

A Survey of Factors Affecting Blunt Leading-Edge Separation for Swept and Semi-Slender Wings

NASA Technical Reports Server (NTRS)

Luckring, James M.

2010-01-01

A survey is presented of factors affecting blunt leading-edge separation for swept and semi-slender wings. This class of separation often results in the onset and progression of separation-induced vortical flow over a slender or semi-slender wing. The term semi-slender is used to distinguish wings with moderate sweeps and aspect ratios from the more traditional highly-swept, low-aspect-ratio slender wing. Emphasis is divided between a selection of results obtained through literature survey a section of results from some recent research projects primarily being coordinated through NATO s Research and Technology Organization (RTO). An aircraft context to these studies is included.

Reduction of the RCS of the leading edge of a conducting wing-shaped structure by means of lossless dielectric material

NASA Astrophysics Data System (ADS)

Booysen, A. J.; Pistorius, C. W. I.; Malherbe, J. A. G.

1991-06-01

The radar cross section of the leading edge of a conducting wing-shaped structure is reduced by replacing part of the structure with a lossless dielectric material. The structure retains its original external shape, thereby ensuring that the aerodynamic properties are not altered by the structural changes needed to reduce the radar cross section.

Leading edge embedded fan airfoil concept -- A new powered high lift technology

NASA Astrophysics Data System (ADS)

Phan, Nhan Huu

A new powered-lift airfoil concept called Leading Edge Embedded Fan (LEEF) is proposed for Extremely Short Take-Off and Landing (ESTOL) and Vertical Take-Off and Landing (VTOL) applications. The LEEF airfoil concept is a powered-lift airfoil concept capable of generating thrust and very high lift-coefficient at extreme angles-of attack (AoA). It is designed to activate only at the take-off and landing phases, similar to conventional flaps or slats, allowing the aircraft to operate efficiently at cruise in its conventional configuration. The LEEF concept consists of placing a crossflow fan (CFF) along the leading-edge (LE) of the wing, and the housing is designed to alter the airfoil shape between take-off/landing and cruise configurations with ease. The unique rectangular cross section of the crossflow fan allows for its ease of integration into a conventional subsonic wing. This technology is developed for ESTOL aircraft applications and is most effectively applied to General Aviation (GA) aircraft. Another potential area of application for LEEF is tiltrotor aircraft. Unlike existing powered high-lift systems, the LEEF airfoil uses a local high-pressure air source from cross-flow fans, does not require ducting, and is able to be deployed using distributed electric power systems throughout the wing. In addition to distributed lift augmentation, the LEEF system can provide additional thrust during takeoff and landing operation to supplement the primary cruise propulsion system. Two-dimensional (2D) and three-dimensional (3D) Computational Fluid Dynamics (CFD) simulations of a conventional airfoil/wing using the NACA 63-3-418 section, commonly used in GA, and a LEEF airfoil/wing embedded into the same airfoil section were carried out to evaluate the advantages of and the costs associated with implementing the LEEF concept. Computational results show that significant lift and augmented thrust are available during LEEF operation while requiring only moderate fan power

Development of a design model for airfoil leading edge film cooling

NASA Astrophysics Data System (ADS)

Wadia, A. R.; Nealy, D. A.

1985-03-01

A series of experiments on scaled cylinder models having injection through holes inclined at 20, 30, 45, and 90 degrees are presented. The experiments were conducted in a wind tunnel on several stainless steel test specimens in which flow and heat transfer parameters were measured over simulated airfoil leading edge surfaces. On the basis of the experimental results, an engineering design model is proposed that treats the gas-to-surface heat transfer coefficient with film cooling in a manner suggested by Luckey and L'Ecuyer (1981). It is shown that the main factor influencing the averaged film cooling effectiveness in the showerhead region is the inclination of the injection holes. The effectiveness parameter was not affected by variations in the coolant-to-gas stream pressure ratio, the freestream Mach number, the gas to coolant temperature ratio, or the gas stream Reynolds number. Experience in the wind tunnel tests is reflected in the design of the model in which the coolant side heat transfer coefficient is offset by a simultaneous increase in the gas side film coefficient. The design applications of the analytical model are discussed, with emphasis given to high temperature first stage turbine vanes and rotor blades.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 1; Sharp Leading Edge; [conducted in the Langley National Transonic Facility (NTF)

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

A method of predicting flow rates required to achieve anti-icing performance with a porous leading edge ice protection system

NASA Technical Reports Server (NTRS)

Kohlman, D. L.; Albright, A. E.

1983-01-01

An analytical method was developed for predicting minimum flow rates required to provide anti-ice protection with a porous leading edge fluid ice protection system. The predicted flow rates compare with an average error of less than 10 percent to six experimentally determined flow rates from tests in the NASA Icing Research Tunnel on a general aviation wing section.

Intermediate experimental vehicle, ESA program aerodynamics-aerothermodynamics key technologies for spacecraft design and successful flight

NASA Astrophysics Data System (ADS)

Dutheil, Sylvain; Pibarot, Julien; Tran, Dac; Vallee, Jean-Jacques; Tribot, Jean-Pierre

2016-07-01

With the aim of placing Europe among the world's space players in the strategic area of atmospheric re-entry, several studies on experimental vehicle concepts and improvements of critical re-entry technologies have paved the way for the flight of an experimental space craft. The successful flight of the Intermediate eXperimental Vehicle (IXV), under ESA's Future Launchers Preparatory Programme (FLPP), is definitively a significant step forward from the Atmospheric Reentry Demonstrator flight (1998), establishing Europe as a key player in this field. The IXV project objectives were the design, development, manufacture and ground and flight verification of an autonomous European lifting and aerodynamically controlled reentry system, which is highly flexible and maneuverable. The paper presents, the role of aerodynamics aerothermodynamics as part of the key technologies for designing an atmospheric re-entry spacecraft and securing a successful flight.

New Hypersonic Shock Tunnel at the Laboratory of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu

DOE Office of Scientific and Technical Information (OSTI.GOV)

Toro, P. G. P.; Minucci, M. A. S.; Chanes, J. B. Jr

The new 0.60-m. nozzle exit diameter hypersonic shock tunnel was designed to study advanced air-breathing propulsion system such as supersonic combustion and/or laser technologies. In addition, it may be used for hypersonic flow studies and investigations of the electromagnetic (laser) energy addition for flow control. This new hypersonic shock tunnel was designed and installed at the Laboratory for of Aerothermodynamics and Hypersonics Prof. Henry T. Nagamatsu, IEAv-CTA, Brazil. The design of the tunnel enables relatively long test times, 2-10 milliseconds, suitable for the experiments performed at the laboratory. Free stream Mach numbers ranging from 6 to 25 can be producedmore » and stagnation pressures and temperatures up to 360 atm. and up to 9,000 K, respectively, can be generated. Shadowgraph and schlieren optical techniques will be used for flow visualization.« less

Edge of polar cap patches

NASA Astrophysics Data System (ADS)

Hosokawa, K.; Taguchi, S.; Ogawa, Y.

2016-04-01

On the night of 4 December 2013, a sequence of polar cap patches was captured by an all-sky airglow imager (ASI) in Longyearbyen, Norway (78.1°N, 15.5°E). The 630.0 nm airglow images from the ASI of 4 second exposure time, oversampled the emission of natural lifetime (with quenching) of at least ˜30 sec, introduce no observational blurring effects. By using such high-quality ASI images, we succeeded in visualizing an asymmetry in the gradients between the leading/trailing edges of the patches in a 2-D fashion. The gradient in the leading edge was found to be 2-3 times steeper than that in the trailing edge. We also identified fingerlike structures, appearing only along the trailing edge of the patches, whose horizontal scale size ranged from 55 to 210 km. These fingers are considered to be manifestations of plasma structuring through the gradient-drift instability (GDI), which is known to occur only along the trailing edge of patches. That is, the current 2-D observations visualized, for the first time, how GDI stirs the patch plasma and such a mixing process makes the trailing edge more gradual. This result strongly implies a close connection between the GDI-driven plasma stirring and the asymmetry in the large-scale shape of patches and then suggests that the fingerlike structures can be used as markers to estimate the fine-scale structure in the plasma flow within patches.

Pressure investigation of NASA leading edge vortex flaps on a 60 deg Delta wing

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Donatelli, D. A.; Terry, J. E.

1983-01-01

Pressure distributions on a 60 deg Delta Wing with NASA designed leading edge vortex flaps (LEVF) were found in order to provide more pressure data for LEVF and to help verify NASA computer codes used in designing these flaps. These flaps were intended to be optimized designs based on these computer codes. However, the pressure distributions show that the flaps wre not optimum for the size and deflection specified. A second drag-producing vortex forming over the wing indicated that the flap was too large for the specified deflection. Also, it became apparent that flap thickness has a possible effect on the reattachment location of the vortex. Research is continuing to determine proper flap size and deflection relationships that provide well-behaved flowfields and acceptable hinge-moment characteristics.

Experimental study of wing leading-edge devices for improved maneuver performance of a supercritical maneuvering fighter configuration. [Langley 7- by 10-ft high speed tunnel tests

NASA Technical Reports Server (NTRS)

Mann, M. J.; Huffman, J. K.; Fox, C. H., Jr.; Campbell, R. L.

1983-01-01

Wind tunnel tests were conducted to examine the use of wing leading-edge devices for improved subsonic and transonic maneuver performance. These devices were tested on a fighter configuration which utilized supercritical-wing technology. The configuration had a leading-edge sweep of 45 deg and an aspect ratio of 3.28. The tests were conducted at Mach numbers of 0.60 and 0.85 with angles of attack from -0.5 deg to 22 deg. At both Mach numbers, sharp leading-edge flaps produced vortices which greatly altered the flow pattern on the wing and resulted in substantial reductions in drag at high lift. Underwing or pylon-type vortex generators also reduced drag at high lift. The vortex generators worked better at a Mach number of 0.60. The vortex generators gave the best overall results with zero toe-in angle and when mounted on either the outboard part of the wing or at both an outboard location and halfway out the semispan. Both the flaps and the vortex generators had a minor effect on the pitching moment. Fluorescent minitufts were found to be useful for flow visualization at transonic maneuver conditions.

The effect of butterfly-scale inspired patterning on leading-edge vortex growth

NASA Astrophysics Data System (ADS)

Wilroy, Jacob; Lang, Amy; Wahidi, Redha

2014-11-01

Leading edge vortices (LEVs) are important for generating thrust and lift in flapping flight, and the surface patterning (scales) on butterfly wings is hypothesized to play a role in the vortex formation of the LEV. To simplify this complex flow problem, we designed an experiment to focus on the alteration of 2-D vortex development with a variation in surface patterning. Specifically we are interested in the secondary vorticity generated by the LEV interacting at the patterned surface and how this can affect the growth rate of the circulation in the LEV. For this experiment we used rapid-prototyped longitudinal and transverse square grooves attached to a flat plate and compared the vortex formation as the plate moved vertically. The plate is impulsively started in quiescent water and flow fields at Re = 1500, 3000, and 6000 are examined using Digital Particle Image Velocimetry (DPIV). The vortex formation time is 0.6 and is based on the flat plate travel length and chord length. Support for this research came from NSF REU Grant 1358991 and CBET 1335848.

The effect of butterfly-scale inspired patterning on leading-edge vortex growth

NASA Astrophysics Data System (ADS)

Wilroy, Jacob; Lang, Amy

2015-11-01

Leading edge vortices (LEVs) are important for generating thrust and lift in flapping flight, and the surface patterning (scales) on butterfly wings is hypothesized to play a role in the vortex formation of the LEV. To simplify this complex flow problem, an experiment was designed to focus on the alteration of 2-D vortex development with a variation in surface patterning. Specifically, the secondary vorticity generated by the LEV interacting at the patterned surface was studied and the subsequent affect on the growth rate of the circulation in the LEV. For this experiment we used butterfly inspired grooves attached to a flat plate and compared the vortex formation to a smooth plate case as the plate moved vertically. The plate is impulsively started in quiescent water and flow fields at Re = 1500, 3000, and 6000 are examined using Digital Particle Image Velocimetry (DPIV). The vortex formation time is 3.0 and is based on the flat plate travel length and chord length. We would like to thank the National Science Foundation REU Site Award 1358991 for funding this research.

Quantifying edge significance on maintaining global connectivity

PubMed Central

Qian, Yuhua; Li, Yebin; Zhang, Min; Ma, Guoshuai; Lu, Furong

2017-01-01

Global connectivity is a quite important issue for networks. The failures of some key edges may lead to breakdown of the whole system. How to find them will provide a better understanding on system robustness. Based on topological information, we propose an approach named LE (link entropy) to quantify the edge significance on maintaining global connectivity. Then we compare the LE with the other six acknowledged indices on the edge significance: the edge betweenness centrality, degree product, bridgeness, diffusion importance, topological overlap and k-path edge centrality. Experimental results show that the LE approach outperforms in quantifying edge significance on maintaining global connectivity. PMID:28349923

Moving Uphill: Microbial Facilitation at the Leading Edge of Plant Species Distributional Shifts

NASA Astrophysics Data System (ADS)

Suding, K.; Farrer, E.; Spasojevic, M.; Porazinska, D.; Bueno de Mesquita, C.; Schmidt, S. K.

2016-12-01

Climate change is expected to influence species distributions and reshuffle patterns of biodiversity. A key challenge to our understanding of these effects is that biotic interactions - new species to compete with, new stressors that increase dependence on facilitation, new prey or predators - will likely affect the ability of species to track climate at the leading edges of their distributional range. While it is well established that soil biota strongly influence plant abundance and diversity, it has been difficult to quantify the key belowground dynamics. This presentation will investigate the influence of one key biotic interaction, between plants and soil microbiota, on the ability of plant species to track climate change and expand their range uphill in a high montane system in the Front Range of Colorado. High-resolution photography from 1972 and 2008 indicate colonization of tundra vegetation in formerly unvegetated areas. Observational work on the distributions patterns of both plants and soil microbiota (bacteria, fungi and nematodes) in a spatially-explicit grid at the upper edge of plant distributions indicate strong, mostly positive, associations between plant species and soil taxa. Abiotic factors, while important, consistently underpredicted the occurrence of plant species and, in nine of the 12 most common tundra plants, co-occurring microbial taxa were important predictors of plant occurrence. Comparison of plant and microbial distributional patterns in 2007 and 2015 indicate the influence of microbial community composition on assembly and beta-diversity of the plant community over time. Plant colonization patterns in this region previously devoid of vegetation will likely influence carbon, nitrogen and phosphorus dynamics, with downstream consequences on nutrient limitation and phytoplankton composition in alpine lakes.

Mars Science Laboratory Entry Capsule Aerothermodynamics and Thermal Protection System

NASA Technical Reports Server (NTRS)

Edquist, Karl T.; Hollis, Brian R.; Dyakonov, Artem A.; Laub, Bernard; Wright, Michael J.; Rivellini, Tomasso P.; Slimko, Eric M.; Willcockson, William H.

2007-01-01

The Mars Science Laboratory (MSL) spacecraft is being designed to carry a large rover (greater than 800 kg) to the surface of Mars using a blunt-body entry capsule as the primary decelerator. The spacecraft is being designed for launch in 2009 and arrival at Mars in 2010. The combination of large mass and diameter with non-zero angle-of-attack for MSL will result in unprecedented convective heating environments caused by turbulence prior to peak heating. Navier-Stokes computations predict a large turbulent heating augmentation for which there are no supporting flight data1 and little ground data for validation. Consequently, an extensive experimental program has been established specifically for MSL to understand the level of turbulent augmentation expected in flight. The experimental data support the prediction of turbulent transition and have also uncovered phenomena that cannot be replicated with available computational methods. The result is that the flight aeroheating environments predictions must include larger uncertainties than are typically used for a Mars entry capsule. Finally, the thermal protection system (TPS) being used for MSL has not been flown at the heat flux, pressure, and shear stress combinations expected in flight, so a test program has been established to obtain conditions relevant to flight. This paper summarizes the aerothermodynamic definition analysis and TPS development, focusing on the challenges that are unique to MSL.

Comparison of Theoretical and Experimental Unsteady Aerodynamics of Linear Oscillating Cascade With Supersonic Leading-Edge Locus

NASA Technical Reports Server (NTRS)

Ramsey, John K.; Erwin, Dan

2004-01-01

An experimental influence coefficient technique was used to obtain unsteady aerodynamic influence coefficients and, consequently, unsteady pressures for a cascade of symmetric airfoils oscillating in pitch about mid-chord. Stagger angles of 0 deg and 10 deg were investigated for a cascade with a gap-to-chord ratio of 0.417 operating at an axial Mach number of 1.9, resulting in a supersonic leading-edge locus. Reduced frequencies ranged from 0.056 to 0.2. The influence coefficients obtained determine the unsteady pressures for any interblade phase angle. The unsteady pressures were compared with those predicted by several algorithms for interblade phase angles of 0 deg and 180 deg.

CFD study on NACA 4415 airfoil implementing spherical and sinusoidal Tubercle Leading Edge

PubMed Central

2017-01-01

The Humpback whale tubercles have been studied for more than a decade. Tubercle Leading Edge (TLE) effectively reduces the separation bubble size and helps in delaying stall. They are very effective in case of low Reynolds number flows. The current Computational Fluid Dynamics (CFD) study is on NACA 4415 airfoil, at a Reynolds number 120,000. Two TLE shapes are tested on NACA 4415 airfoil. The tubercle designs implemented on the airfoil are sinusoidal and spherical. A parametric study is also carried out considering three amplitudes (0.025c, 0.05c and 0.075c), the wavelength (0.25c) is fixed. Structured mesh is utilized to generate grid and Transition SST turbulence model is used to capture the flow physics. Results clearly show spherical tubercles outperform sinusoidal tubercles. Furthermore experimental study considering spherical TLE is carried out at Reynolds number 200,000. The experimental results show that spherical TLE improve performance compared to clean airfoil. PMID:28850622

CFD study on NACA 4415 airfoil implementing spherical and sinusoidal Tubercle Leading Edge.

PubMed

Aftab, S M A; Ahmad, K A

2017-01-01

The Humpback whale tubercles have been studied for more than a decade. Tubercle Leading Edge (TLE) effectively reduces the separation bubble size and helps in delaying stall. They are very effective in case of low Reynolds number flows. The current Computational Fluid Dynamics (CFD) study is on NACA 4415 airfoil, at a Reynolds number 120,000. Two TLE shapes are tested on NACA 4415 airfoil. The tubercle designs implemented on the airfoil are sinusoidal and spherical. A parametric study is also carried out considering three amplitudes (0.025c, 0.05c and 0.075c), the wavelength (0.25c) is fixed. Structured mesh is utilized to generate grid and Transition SST turbulence model is used to capture the flow physics. Results clearly show spherical tubercles outperform sinusoidal tubercles. Furthermore experimental study considering spherical TLE is carried out at Reynolds number 200,000. The experimental results show that spherical TLE improve performance compared to clean airfoil.

Experimental investigation of edge hardening and edge cracking sensitivity of burr-free parts

NASA Astrophysics Data System (ADS)

Senn, Sergei; Liewald, Mathias

2018-05-01

This experimental study is focused on characterisation of edge hardening of sheet metal and remaining formability of differently prepared cutted edges. Edge cracking sensitivity of counter cutted, shear cutted, recutted and water-jet cutted components are compared and evaluated. Subsequently, edge hardening and hole expansion ratio were correlated for material HC420 LA with sheet thickness of t = 2 mm. As other studies show, the cutting edge surface quality influences the hole expansion ratio: a high clear cut surface increases formability of cutting edges, whereas micro cracks and rough surfaces result into a large fracture surface, which impact remaining formability noticeably. Thus, cutting edges with lower edge hardening behaviour in conjunction with a higher clear cut surface exhibit higher hole expansion ratios. Counter cutting and the recutting do show a similar effect on edge hardening. Using the hole expansion test, it was possible to prove that counter cutted components show a significantly lower edge cracking sensitivity in comparison to conventionally shear cutted components. The hole expansion ratio of counter cutted specimens looks balanced and is comparable to the hole expansion ratio measured from specimens with recutted or water jet cutted edges. The significant difference of the investigated cutting processes is characterized by size of clear cutting area. This area of recutted edges emerges larger than the area of counter cutted specimens, which evidently leads to an increased hole expansion ratio of recutted specimens compared to conventionally shear cutted ones. However, it is important to note that the hole expansion ratio of counter cutted and recutted specimens appear fairly balanced, but counter cutted samples indeed can be produced burr-free. Using counter cutting technology, it is possible to produce burr free surfaces with high edge formability.

Wind-Tunnel Tests on Various Types of Dive Brakes Mounted in Proximity of the Leading Edge of the Wing

NASA Technical Reports Server (NTRS)

Lattanzi, Bernardino; Bellante, Erno

1949-01-01

The present report is concerned with a series of tests on a model airplane fitted with four types of dive flaps of various shapes, positions, and incidence located near the leading edge of the wing (from 5 to 20 percent of the wing chord). Tests were also made on a stub airfoil fitted with a ventral dive (located at 8 percent of the wing chord). The hinge moments of the dive flaps were measured.

Dynamic Stall Measurements and Computations for a VR-12 Airfoil with a Variable Droop Leading Edge

NASA Technical Reports Server (NTRS)

Martin, P. B.; McAlister, K. W.; Chandrasekhara, M. S.; Geissler, W.

2003-01-01

High density-altitude operations of helicopters with advanced performance and maneuver capabilities have lead to fundamental research on active high-lift system concepts for rotor blades. The requirement for this type of system was to improve the sectional lift-to-drag ratio by alleviating dynamic stall on the retreating blade while simultaneously reducing the transonic drag rise of the advancing blade. Both measured and computational results showed that a Variable Droop Leading Edge (VDLE) airfoil is a viable concept for application to a rotor high-lift system. Results are presented for a series of 2D compressible dynamic stall wind tunnel tests with supporting CFD results for selected test cases. These measurements and computations show a dramatic decrease in the drag and pitching moment associated with severe dynamic stall when the VDLE concept is applied to the Boeing VR-12 airfoil. Test results also show an elimination of the negative pitch damping observed in the baseline moment hysteresis curves.

Losing your edge: climate change and the conservation value of range-edge populations.

PubMed

Rehm, Evan M; Olivas, Paulo; Stroud, James; Feeley, Kenneth J

2015-10-01

Populations occurring at species' range edges can be locally adapted to unique environmental conditions. From a species' perspective, range-edge environments generally have higher severity and frequency of extreme climatic events relative to the range core. Under future climates, extreme climatic events are predicted to become increasingly important in defining species' distributions. Therefore, range-edge genotypes that are better adapted to extreme climates relative to core populations may be essential to species' persistence during periods of rapid climate change. We use relatively simple conceptual models to highlight the importance of locally adapted range-edge populations (leading and trailing edges) for determining the ability of species to persist under future climates. Using trees as an example, we show how locally adapted populations at species' range edges may expand under future climate change and become more common relative to range-core populations. We also highlight how large-scale habitat destruction occurring in some geographic areas where many species range edge converge, such as biome boundaries and ecotones (e.g., the arc of deforestation along the rainforest-cerrado ecotone in the southern Amazonia), can have major implications for global biodiversity. As climate changes, range-edge populations will play key roles in helping species to maintain or expand their geographic distributions. The loss of these locally adapted range-edge populations through anthropogenic disturbance is therefore hypothesized to reduce the ability of species to persist in the face of rapid future climate change.

Enhancing the hydrodynamic performance of a tapered swept-back wing through leading-edge tubercles

NASA Astrophysics Data System (ADS)

Wei, Zhaoyu; Lian, Lian; Zhong, Yisen

2018-06-01

The hydrodynamic benefit of implementing leading-edge (LE) tubercles on wings at very low Reynolds numbers ( Res) has not been thoroughly elucidated to date, though their benefits at relatively higher Res are well-studied. Through wind tunnel testing at Re = 5.5 × 104, we found that the LE tubercles increase the lift at all pitch angles tested and slightly reduce the drag at a pitch angle of 4° < α < 10°, which finally results in a significant hydrodynamic performance enhancement at lower pitch angles. Flow visualization reveals that the hydrodynamic performance enhancement is due to the favourable attached flows downstream of the tubercle peaks. The attached flows are believed to be closely related to the downwash and momentum exchange within the boundary layers, which originate from surface and streamwise-aligned counter-rotating vortex pairs (CVPs).

An experimental study of pressures on 60 deg Delta wings with leading edge vortex flaps

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Terry, J. E.; Donatelli, D. A.

1983-01-01

An experimental study was conducted in the Virginia Tech Stability Wind Tunnel to determine surface pressures over a 60 deg sweep delta wing with three vortex flap designs. Extensive pressure data was collected to provide a base data set for comparison with computational design codes and to allow a better understanding of the flow over vortex flaps. The results indicated that vortex flaps can be designed which will contain the leading edge vortex with no spillage onto the wing upper surface. However, the tests also showed that flaps designed without accounting for flap thickness will not be optimum and the result can be oversized flaps, early flap vortex reattachment and a second separation and vortex at the wing/flap hinge line.

Comparison of two leading uniform theories of edge diffraction with the exact uniform asymptotic solution

NASA Technical Reports Server (NTRS)

Boersma, J.; Rahmat-Samii, Y.

1980-01-01

The diffraction of an arbitrary cylindrical wave by a half-plane has been treated by Rahmat-Samii and Mittra who used a spectral domain approach. In this paper, their exact solution for the total field is expressed in terms of a new integral representation. For large wave number k, two rigorous procedures are described for the exact uniform asymptotic expansion of the total field solution. The uniform expansions obtained are valid in the entire space, including transition regions around the shadow boundaries. The final results are compared with the formulations of two leading uniform theories of edge diffraction, namely, the uniform asymptotic theory and the uniform theory of diffraction. Some unique observations and conclusions are made in relating the two theories.

Sensitivity of F-106B Leading-Edge-Vortex Images to Flight and Vapor-Screen Parameters

NASA Technical Reports Server (NTRS)

Lamar, John E.; Johnson, Thomas D., Jr.

1988-01-01

A flight test was undertaken at NASA Langley Research Center with vapor-screen and image-enhancement techniques to obtain qualitative and quantitative information about near-field vortex flows above the wings of fighter aircraft. In particular, the effects of Reynolds and Mach numbers on the vortex system over an angle-of-attack range were sought. The relevance of these flows stems from their present and future use at many points in the flight envelope, especially during transonic maneuvers. The aircraft used in this flight program was the F-106B because it was available and had sufficient wing sweep (60 deg) to generate a significant leading-edge vortex system. The sensitivity of the visual results to vapor screen hardware and to onset flow changes is discussed.

Adding-point strategy for reduced-order hypersonic aerothermodynamics modeling based on fuzzy clustering

NASA Astrophysics Data System (ADS)

Chen, Xin; Liu, Li; Zhou, Sida; Yue, Zhenjiang

2016-09-01

Reduced order models(ROMs) based on the snapshots on the CFD high-fidelity simulations have been paid great attention recently due to their capability of capturing the features of the complex geometries and flow configurations. To improve the efficiency and precision of the ROMs, it is indispensable to add extra sampling points to the initial snapshots, since the number of sampling points to achieve an adequately accurate ROM is generally unknown in prior, but a large number of initial sampling points reduces the parsimony of the ROMs. A fuzzy-clustering-based adding-point strategy is proposed and the fuzzy clustering acts an indicator of the region in which the precision of ROMs is relatively low. The proposed method is applied to construct the ROMs for the benchmark mathematical examples and a numerical example of hypersonic aerothermodynamics prediction for a typical control surface. The proposed method can achieve a 34.5% improvement on the efficiency than the estimated mean squared error prediction algorithm and shows same-level prediction accuracy.

Rate and topography of peptidoglycan synthesis during cell division in Escherichia coli: Concept of a leading edge

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wientjes, F.B.; Nanninga, N.

1989-06-01

The rate at which the peptidoglycan of Escherichia coli is synthesized during the division cycle was studied with two methods. One method involved synchronization of E. coli MC4100 lysA cultures by centrifugal elutriation and subsequent pulse-labeling of the synchronously growing cultures with (meso-{sup 3}H)diaminopimelic acid (({sup 3}H)Dap). The second method was autoradiography of cells pulse-labeled with ({sup 3}H)Dap. It was found that the peptidoglycan is synthesized at a more or less exponentially increasing rate during the division cycle with a slight acceleration in this rate as the cells start to constrict. Apparently, polar cap formation requires synthesis of extra surfacemore » components, presumably to accommodate for a change in the surface-to-volume ratio. Furthermore, it was found that the pool size of Dap was constant during the division cycle. Close analysis of the topography of ({sup 3}H)Dap incorporation at the constriction site revealed that constriction proceeded by synthesis of peptidoglycan at the leading edge of the invaginating cell envelope. During constriction, no reallocation of incorporation occurred, i.e., the incorporation at the leading edge remained high throughout the process of constriction. Impairment of penicillin-binding protein 3 by mutation or by the specific {beta}-lactam antibiotic furazlocillin did not affect ({sup 3}H)Dap incorporation during initiation of constriction. However, the incorporation at the constriction site was inhibited in later stages of the constriction process. It is concluded that during division at least two peptidoglycan-synthesizing systems are operating sequentially.« less

Morphological Variations of Leading-Edge Serrations in Owls (Strigiformes).

PubMed

Weger, Matthias; Wagner, Hermann

2016-01-01

Owls have developed serrations, comb-like structures, along the leading edge of their wings. Serrations were investigated from a morphological and a mechanical point of view, but were not yet quantitatively compared for different species. Such a comparative investigation of serrations from species of different sizes and activity patterns may provide new information about the function of the serrations. Serrations on complete wings and on tenth primary remiges of seven owl species were investigated. Small, middle-sized, and large owl species were investigated as well as species being more active during the day and owls being more active during the night. Serrations occurred at the outer parts of the wings, predominantly at tenth primary remiges, but also on further wing feathers in most species. Serration tips were oriented away from the feather rachis so that they faced into the air stream during flight. The serrations of nocturnal owl species were higher developed as demonstrated by a larger inclination angle (the angle between the base of the barb and the rachis), a larger tip displacement angle (the angle between the tip of the serration and the base of the serration) and a longer length. Putting the measured data into a clustering algorithm yielded dendrograms that suggested a strong influence of activity pattern, but only a weak influence of size on the development of the serrations. Serrations are supposed to be involved in noise reduction during flight and also depend on the aerodynamic properties that in turn depend on body size. Since especially nocturnal owls have to rely on hearing during prey capture, the more pronounced serrations of nocturnal species lend further support to the notion that serrations have an important function in noise reduction. The differences in shape of the serrations investigated indicate that a silent flight requires well-developed serrations.

Morphological Variations of Leading-Edge Serrations in Owls (Strigiformes)

PubMed Central

Weger, Matthias; Wagner, Hermann

2016-01-01

Background Owls have developed serrations, comb-like structures, along the leading edge of their wings. Serrations were investigated from a morphological and a mechanical point of view, but were not yet quantitatively compared for different species. Such a comparative investigation of serrations from species of different sizes and activity patterns may provide new information about the function of the serrations. Results Serrations on complete wings and on tenth primary remiges of seven owl species were investigated. Small, middle-sized, and large owl species were investigated as well as species being more active during the day and owls being more active during the night. Serrations occurred at the outer parts of the wings, predominantly at tenth primary remiges, but also on further wing feathers in most species. Serration tips were oriented away from the feather rachis so that they faced into the air stream during flight. The serrations of nocturnal owl species were higher developed as demonstrated by a larger inclination angle (the angle between the base of the barb and the rachis), a larger tip displacement angle (the angle between the tip of the serration and the base of the serration) and a longer length. Putting the measured data into a clustering algorithm yielded dendrograms that suggested a strong influence of activity pattern, but only a weak influence of size on the development of the serrations. Conclusions Serrations are supposed to be involved in noise reduction during flight and also depend on the aerodynamic properties that in turn depend on body size. Since especially nocturnal owls have to rely on hearing during prey capture, the more pronounced serrations of nocturnal species lend further support to the notion that serrations have an important function in noise reduction. The differences in shape of the serrations investigated indicate that a silent flight requires well-developed serrations. PMID:26934104

An improved panel method for the solution of three-dimensional leading edge vortex flows Volume 2: User's guide and programmer's document

NASA Technical Reports Server (NTRS)

Tinoco, E. N.; Lu, P.; Johnson, F. T.

1980-01-01

A computer program developed for solving the subsonic, three dimensional flow over wing-body configurations with leading edge vortex separation is presented. Instructions are given for the proper set up and input of a problem into the computer code. Program input formats and output are described, as well as the overlay structure of the program. The program is written in FORTRAN.

Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

NASA Technical Reports Server (NTRS)

Micol, J. R.

1998-01-01

Description, capabilities, recent upgrades, and utilization of the NASA Langley Research Center (LaRC) Aerothermodynamic Facilities Complex (AFC) are presented. The AFC consists of five hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 20, unit Reynolds number from 0.04 to 22 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of three different test gases (air, helium, and tetrafluoromethane), thereby making several of the facilities unique. The Complex represents nearly three-fourths of the conventional (as opposed to impulse)-type hypersonic wind tunnels operational in this country. AFC facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data fr generating the flight aerodynamic databook and final design of the thermal protection system (TPS) (e.g., establishment of flight limitations not to exceed TPS design limits). Modifications and enhancements of AFC hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. Examples illustrating facility utilization in recent years to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs are presented. These programs include Phase I of the Reusable Launch Vehicle (RLV) Advanced Technology Demonstrator, X-33 program, PHase II of the X-33 program, X-34 program, the Hyper-X program ( a Mach 5,7, and 10 airbreathing propulsion flight experiment), and the X-38 program (Experimental Crew Return Vehicle, X-CRV). Current upgrades/enchancements and future plans for the AFC are discussed.

Organization of actin in the leading edge of cultured cells: influence of osmium tetroxide and dehydration on the ultrastructure of actin meshworks

PubMed Central

1981-01-01

The ordered structure of the leading edge (lamellipodium) of cultured fibroblasts is readily revealed in cells extracted briefly in Triton X- 100-glutaraldehyde mixtures, fixed further in glutaraldehyde, and then negatively stained for electron microscopy. By this procedure, the leading edge regions show a highly organised, three-dimensional network of actin filaments together with variable numbers of radiating actin filament bundles or microspikes. The use of Phalloidin after glutaraldehyde fixation resulted in a marginal improvement in filament order. Processing of the cytoskeletons though the additional steps generally employed for conventional electron microscopy resulted in a marked deterioration or complete disruption of the order of the actin filament networks. In contrast, the actin filaments of the stress fiber bundles were essentially unaffected. Thus, postfixation in osmium tetroxide (1% for 7 min at room temperature) transformed the networks to a reticulum of kinked fibers, resembling those produced by the exposure of muscle F-actin to OsO4 in vitro (P. Maupin-Szamier and T. D. Pollard. 1978. J. Cell Biol. 77:837--852). While limited exposure to OsO4 (0.2+ for 20 min at 0 degrees C) obviated this destruction, dehydration in acetone or ethanol, with or without post-osmication, caused a further and unavoidable disordering and aggregation of the meshwork filaments. The meshwork regions of the leading edge then showed a striking resemblance to the networks hitherto described in critical point-dried preparations of cultured cells. I conclude that much of the "microtrabecular lattice" described by Wolosewick and Porter (1979. J. Cell Biol. 82:114--139) in the latter preparations constitutes actin meshworks and actin filament arrays, with their associated components, that have been distorted and aggregated by the preparative procedures employed. PMID:6799521

An improved panel method for the solution of three-dimensional leading-edge vortex flows. Volume 1: Theory document

NASA Technical Reports Server (NTRS)

Johnson, F. T.; Lu, P.; Tinoco, E. N.

1980-01-01

An improved panel method for the solution of three dimensional flow and wing and wing-body combinations with leading edge vortex separation is presented. The method employs a three dimensional inviscid flow model in which the configuration, the rolled-up vortex sheets, and the wake are represented by quadratic doublet distributions. The strength of the singularity distribution as well as shape and position of the vortex spirals are computed in an iterative fashion starting with an assumed initial sheet geometry. The method calculates forces and moments as well as detail surface pressure distributions. Improvements include the implementation of improved panel numerics for the purpose of elimination the highly nonlinear effects of ring vortices around double panel edges, and the development of a least squares procedure for damping vortex sheet geometry update instabilities. A complete description of the method is included. A variety of cases generated by the computer program implementing the method are presented which verify the mathematical assumptions of the method and which compare computed results with experimental data to verify the underlying physical assumptions made by the method.

A method to design blended rolled edges for compact range reflectors

NASA Technical Reports Server (NTRS)

Gupta, Inder J.; Burnside, Walter D.

1989-01-01

A method to design blended rolled edges for arbitrary rim shape compact range reflectors is presented. The reflectors may be center-fed or offset-fed. The method leads to rolled edges with minimal surface discontinuities. It is shown that the reflectors designed using the prescribed method can be defined analytically using simple expressions. A procedure to obtain optimum rolled edges parameter is also presented. The procedure leads to blended rolled edges that minimize the diffracted fields emanating from the junction between the paraboloid and the rolled edge surface while satisfying certain constraints regarding the reflector size and the minimum operating frequency of the system.

A method to design blended rolled edges for compact range reflectors

NASA Technical Reports Server (NTRS)

Gupta, Inder J.; Ericksen, Kurt P.; Burnside, Walter D.

1990-01-01

A method to design blended rolled edges for arbitrary rim shape compact range reflectors is presented. The reflectors may be center-fed or offset-fed. The method leads to rolled edges with minimal surface discontinuities. It is shown that the reflectors designed using the prescribed method can be defined analytically using simple expressions. A procedure to obtain optimum rolled edges parameters is also presented. The procedure leads to blended rolled edges that minimize the diffracted fields emanating from the junction between the paraboloid and the rolled edge surface while satisfying certain constraints regarding the reflector size and the minimum operating frequency of the system.

CFD Analysis of the Aerodynamics of a Business-Jet Airfoil with Leading-Edge Ice Accretion

NASA Technical Reports Server (NTRS)

Chi, X.; Zhu, B.; Shih, T. I.-P.; Addy, H. E.; Choo, Y. K.

2004-01-01

For rime ice - where the ice buildup has only rough and jagged surfaces but no protruding horns - this study shows two dimensional CFD analysis based on the one-equation Spalart-Almaras (S-A) turbulence model to predict accurately the lift, drag, and pressure coefficients up to near the stall angle. For glaze ice - where the ice buildup has two or more protruding horns near the airfoil's leading edge - CFD predictions were much less satisfactory because of the large separated region produced by the horns even at zero angle of attack. This CFD study, based on the WIND and the Fluent codes, assesses the following turbulence models by comparing predictions with available experimental data: S-A, standard k-epsilon, shear-stress transport, v(exp 2)-f, and differential Reynolds stress.

Computational Modeling of Three-Dimensional Compressible Leading Edge Vortices