Science.gov

Sample records for leading edge aerothermodynamics

  1. Aerothermal/FEM Analysis of Hypersonic Sharp Leading Edges

    NASA Technical Reports Server (NTRS)

    Kolodziej, Paul; Bull, Jeffrey D.; Kowalski, Thomas R.; Rasky, Daniel J. (Technical Monitor)

    1995-01-01

    Advanced hypersonic vehicles, like wave riders, will have sharp leading edges to minimize drag. These designs require accurate finite element modeling (FEM) of the thermal-structural behavior of a diboride ceramic matrix composite sharp leading edge. By coupling the FEM solver to an engineering model of the aerothermodynamic heating environment the impact of non catalytic surfaces, rarefied flow effects, and multidimensional conduction on the performance envelopes of sharp leading edges can be examined.

  2. Supersonic Leading Edge Receptivity

    NASA Technical Reports Server (NTRS)

    Maslov, Anatoly A.

    1998-01-01

    This paper describes experimental studies of leading edge boundary layer receptivity for imposed stream disturbances. Studies were conducted in the supersonic T-325 facility at ITAM and include data for both sharp and blunt leading edges. The data are in agreement with existing theory and should provide guidance for the development of more complete theories and numerical computations of this phenomena.

  3. Computational aerothermodynamics

    NASA Technical Reports Server (NTRS)

    Deiwert, George S.; Green, Michael J.

    1987-01-01

    Computational aerothermodynamics (CAT) has in the past contributed to the understanding of real-gas flows encountered by hypervelocity reentry vehicles. With advances in computational fluid dynamics, in the modeling of high temperature phenomena, and in computer capability, CAT is an enabling technology for the design of many future space vehicles. An overview of the current capabilities of CAT is provided by describing available methods and their applications. Technical challenges that need to be met are discussed.

  4. Moveable Leading Edge Device for a Wing

    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 section on a wing of an aircraft is extended into a deployed position. A deformable section connects the leading edge section to a trailing section. The deformable section changes from a deformed shape to an original shape when the leading edge section is moved into the deployed position. The leading edge section on the wing is moved from the deployed position to an undeployed position. The deformable section changes to the deformed shape inside of the wing.

  5. Aerothermal Performance Constraints for Small Radius Leading Edges Operating at Hypervelocity

    NASA Technical Reports Server (NTRS)

    Kolodziej, Paul; Bull, Jeffrey D.; Milos, Frank S.; Squire, Thomas H.

    1997-01-01

    Small radius leading edges and nosetips were used to minimize wave drag in early hypervelocity vehicle concepts until further analysis demonstrated that extreme aerothermodynamic heating blunted the available thermal protection system materials. Recent studies indicate that ultra-high temperature composite (UHTC) materials are shape stable at temperatures approaching 3033 K and will be available for use as sharp leading edge components in the near future. Steady-state aerothermal performance constraints for UHTC components are presented in this paper to identify their non-ablating operational capability at altitudes from sea level to 90 km. An integrated design tool was developed to estimate these constraints. The tool couples aerothermodynamic heating with material response using commercial finite element analysis software and is capable of both steady-state and transient analysis. Performance during entry is analyzed by transient thermal analysis along the trajectory. The thermal load condition from the transient thermal analysis is used to estimate thermal stress. Applying the tool to UHTC materials shows that steady-state, non-ablating operation of a HfB2/SiC(A-7) (A-7) component is possible at velocities approaching Earth's circular orbital velocity of 7.9 km/s at altitudes approaching 70 km.

  6. Aerothermal Performance Constraints for Hypervelocity Small Radius Unswept Leading Edges and Nosetips

    NASA Technical Reports Server (NTRS)

    Kolodziej, Paul

    1997-01-01

    Small radius leading edges and nosetips were utilized to minimize wave drag in early hypervelocity vehicle concepts until further analysis demonstrated that extreme aerothermodynamic heating would cause severe ablation or blunting of the available thermal protection system materials. Recent studies indicate that ultrahigh temperature ceramic (UHTC) materials are shape stable at temperatures approaching 3033 K and will be available for use as sharp UHTC leading edge components in the near future. Aerothermal performance constraints for sharp components made from these materials are presented in this work to demonstrate the effects of convective blocking, surface catalycity, surface emissivity, and rarefied flow effects on steady state operation at altitudes from sea level to 90 km. These components are capable of steady state operation at velocities up to 7.9 km/s at attitudes near 90 km.

  7. Cavitation on hydrofoils with sinusoidal leading edge

    NASA Astrophysics Data System (ADS)

    Johari, H.

    2015-12-01

    Cavitation characteristics of hydrofoils with sinusoidal leading edge were examined experimentally at a Reynolds number of 7.2 × 105. The hydrofoils had an underlying NACA 634-021 profile and an aspect ratio of 4.3. The sinusoidal leading edge geometries included three amplitudes of 2.5%, 5%, and 12% and two wavelengths of 25% and 50% of the mean chord length. Results revealed that cavitation on the leading edge-modified hydrofoils existed in pockets behind the troughs whereas the baseline hydrofoil produced cavitation along its entire span. Moreover, cavitation on the modified hydrofoils appeared at consistently lower angles of attack than on the baseline hydrofoil.

  8. Leading edge protection for composite blades

    NASA Technical Reports Server (NTRS)

    Brantley, J. W.; Irwin, T. P. (Inventor)

    1977-01-01

    A laminated filament composite structure, such as an airfoil for use in an environment in which it is subjected to both foreign object impact and bending is provided with improved leading edge protection. At least one fine wire mesh layer is partially bonded within the composite structure along its neutral bending axis. A portion of the wire mesh layer extends beyond the neutral bending axis and partially around the leading edge where it is bonded to the outer periphery of the primary composite structure. The wire mesh is clad with a metal such as nickel to provide an improved leading edge protective device which is firmly anchored within the composite structure. Also described is a novel method of constructing a composite airfoil so as to further minimize the possibility of losing the leading edge protective device due to delamination caused by impact and bending.

  9. Computation of leading-edge vortex flows

    NASA Technical Reports Server (NTRS)

    Newsome, R. W.; Thomas, J. L.

    1986-01-01

    The simulation of the leading edge vortex flow about a series of conical delta wings through solution of the Navier-Stokes and Euler equations is studied. The occurrence, the validity, and the usefulness of separated flow solutions to the Euler equations of particular interest. Central and upwind difference solutions to the governing equations are compared for a series of cross sectional shapes, including both rounded and sharp tip geometries. For the rounded leading edge and the flight condition considered, viscous solutions obtained with either central or upwind difference methods predict the classic structure of vortical flow over a highly swept delta wing. Predicted features include the primary vortex due to leading edge separation and the secondary vortex due to crossflow separation. Central difference solutions to the Euler equations show a marked sensitivity to grid refinement. On a coarse grid, the flow separates due to numerical error and a primary vortex which resembles that of the viscous solution is predicted. In contrast, the upwind difference solutions to the Euler equations predict attached flow even for first-order solutions on coarse grids. On a sufficiently fine grid, both methods agree closely and correctly predict a shock-curvature-induced inviscid separation near the leeward plane of symmetry. Upwind difference solutions to the Navier-Stokes and Euler equations are presented for two sharp leading edge geometries. The viscous solutions are quite similar to the rounded leading edge results with vortices of similar shape and size. The upwind Euler solutions predict attached flow with no separation for both geometries. However, with sufficient grid refinement near the tip or through the use of more accurate spatial differencing, leading edge separation results. Once the leading edge separation is established, the upwind solution agrees with recently published central difference solutions to the Euler equations.

  10. 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.

  11. 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.

  12. 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.

  13. Wing Leading Edge Joint Laminar Flow Tests

    NASA Technical Reports Server (NTRS)

    Drake, Aaron; Westphal, Russell V.; Zuniga, Fanny A.; Kennelly, Robert A., Jr.; Koga, Dennis J.

    1996-01-01

    An F-104G aircraft at NASA's Dryden Flight Research Center has been equipped with a specially designed and instrumented test fixture to simulate surface imperfections of the type likely to be present near the leading edge on the wings of some laminar flow aircraft. The simulated imperfections consisted of five combinations of spanwise steps and gaps of various sizes. The unswept fixture yielded a pressure distribution similar to that of some laminar flow airfoils. The experiment was conducted at cruise conditions typical for business-jets and light transports: Mach numbers were in the range 0.5-0.8, and unit Reynolds numbers were 1.5-2.5 million per foot. Skin friction measurements indicated that laminar flow was often maintained for some distance downstream of the surface imperfections. Further work is needed to more precisely define transition location and to extend the experiments to swept-wing conditions and a broader range of imperfection geometries.

  14. 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.

  15. 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.

  16. 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.

  17. Studies of leading-edge thrust phenomena

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

    A study of practical limitations on achievement of theoretical leading-edge thrust has been made and an empirical method for estimation of attainable thrust has been developed. The method is based on a theoretical analysis of a set of two-dimensional airfoils to define thrust dependence on airfoil geometric characteristics and arbitrarily defined limiting pressures, an examination of two-dimensional airfoil experimental data to provide an estimate of limiting pressure dependence on local Mach number and Reynolds number, and employment of simple sweep theory to adapt the method to three-dimensional wings. Because the method takes into account the spanwise variation of airfoil section characteristics, an opportunity is afforded for design by iteration to maximize the attainable thrust and the attendant performance benefits. The applicability of the method was demonstrated by comparisons of theoretical and experimental aerodynamic characteristics for a series of wing-body configurations. Generally, good predictions of the attainable thrust and its influence on lift and drag characteristics were obtained over a range of Mach numbers from 0.24 to 2.0.

  18. Aerothermodynamic Data Base

    NASA Technical Reports Server (NTRS)

    1984-01-01

    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. A list of documentation of DMS processed data arranged sequentially and by space shuttle configuration is presented. The listing provides an up to date record of all applicable aerothermodynamic data collected, processed, or summarized during the space shuttle program. Tables are designed to provide survey information to the various space shuttle managerial and technical levels.

  19. 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.

  20. Separation bubble around a leading edge of compressor blade

    NASA Astrophysics Data System (ADS)

    Liu, Huo-xing; Liu, Bao-jie; Li, Ling; Jiang, Hao-Kang

    2003-04-01

    This paper presents an experimental study of the influence of 2D leading-edge geometry on transition and performance. The measurements were conducted on a special large-scale experimental facility, the pressure distribution and flow field were measured. The test model used in this study consists of circular leading edge and elliptic leading edge. Results are presented for a range of incidence. The measurement result indicated that the leading edge shape has a large influence on flow details, separation and transition as well as the boundary layer properties after reattached point.

  1. 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.

  2. Turbulent Wing-Leading-Edge Correlation Assessment for the Shuttle Orbiter

    NASA Technical Reports Server (NTRS)

    King, Rudolph A.; Vaughan, Matthew P.

    2009-01-01

    This study was conducted in support of the Orbiter damage assessment activity that takes place for each Shuttle mission since STS-107 (STS - Space Transportation System). As part of the damage assessment activity, the state of boundary layer (laminar or turbulent) during reentry needs to be estimated in order to define the aerothermal environment on the Orbiter. Premature turbulence on the wing leading edge (WLE) is possible if a surface irregularity promotes early transition and the resulting turbulent wedge flow contaminates the WLE flow. The objective of this analysis is to develop a criterion to determine if and when the flow along the WLE experiences turbulent heating given an incoming turbulent boundary layer that contaminates the attachment line. The data to be analyzed were all obtained as part of the MH-13 Space Shuttle Orbiter Aerothermodynamic Test conducted on a 1.8%-scale Orbiter model at Calspan/University of Buffalo Research Center in the Large Energy National Shock Tunnels facility. A rational framework was used to develop a means to assess the state of the WLE flow on the Orbiter during reentry given a contaminated attachment-line flow. Evidence of turbulent flow on the WLE has been recently documented for a few STS missions during the Orbiter s flight history, albeit late in the reentry trajectory. The criterion developed herein will be compared to these flight results.

  3. 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.

  4. Leading-Edge "Pop-Up" Spoiler For Airfoil

    NASA Technical Reports Server (NTRS)

    Wilson, John C.; Lance, Michael B.

    1991-01-01

    New concept places spoiler in leading edge of airfoil, hinged along its trailing edge, so airflow helps to deploy it and force it against mechanical stop. Deployed "pop-up" spoiler quickly eliminates almost all aerodynamic lift of stabilator. Designed to be added to leading edge of existing stabilator, without major rework. Though initial application to be on helicopter stabilators, equally applicable to wings or winglike components.

  5. Numerical study of delta wing leading edge blowing

    NASA Technical Reports Server (NTRS)

    Yeh, David; Tavella, Domingo; Roberts, Leonard

    1988-01-01

    Spanwise and tangential leading edge blowing as a means of controlling the position and strength of the leading edge vortices are studied by numerical solution of the three-dimensional Navier-Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of spanwise leading edge blowing at moderate angle of attack magnifies the size and strength of the leading edge vortices, and moves the vortex cores outboard and upward. The increase in lift primarily comes from the greater nonlinear vortex lift. However, spanwise blowing causes earlier vortex breakdown, thus decreasing the stall angle. The effects of tangential blowing at low to moderate angles of attack tend to reduce the pressure peaks associated with leading edge vortices and to increase the suction peak around the leading edge, so that the integrated value of the surface pressure remains about the same. Tangential leading edge blowing in post-stall conditions is shown to re-establish vortical flow and delay vortex bursting, thus increasing C sub L sub max and stall angle.

  6. 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.

  7. 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.

  8. 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.

  9. Vortex leading edge flap assembly for supersonic airplanes

    NASA Technical Reports Server (NTRS)

    Rudolph, Peter K. C. (Inventor)

    1997-01-01

    A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

  10. 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.

  11. 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.

  12. 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.

  13. Effect of Leading Edge Tubercles on Swept Humpback Whale Flipper

    NASA Astrophysics Data System (ADS)

    Murray, Mark; Miklosovic, David; Fish, Frank; Howle, Laurens

    2004-11-01

    The effect of leading edge tubercles on the performance of idealized humpback whale flipper models at sweep angles of 15 and 30 degrees is analyzed. We present the experimental results based on precision wind tunnel testing, comparing the data obtained on idealized model sets with and without leading edge tubercles. We have found a significant difference in the lift and drag coefficients over a large range of angle of attack.

  14. Stall Delay by Leading Edge Tubercles on Humpback Whale Flipper

    NASA Astrophysics Data System (ADS)

    Murray, Mark; Miklosovic, David; Fish, Frank; Howle, Laurens

    2003-11-01

    The effect of leading edge tubercles on the performance of idealized humpback whale flipper models is analyzed. We present the experimental results based on precision wind tunnel testing, comparing the data obtained on idealized model sets with and without leading edge tubercles. We have found a significant increase in the angle of attack required for stall on the flipper with tubercles and a smaller drag coefficient at these higher angles of attack.

  15. 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.

  16. Leading and trailing edge noise of an airfoil

    NASA Astrophysics Data System (ADS)

    Amiet, R. K.

    Theoretical and experimental predictions of the noise produced when a rigid surface, e.g., an airfoil, with a sharp edge is introduced into a turbulent flow are compared. For an airfoil in rectilinear motion agreement is good. It is better for leading edge than for trailing edge noise because of lack of knowledge of boundary layer surface pressure. For a rotating airfoil, leading edge noise produces spectral peaking around harmonics of blade passage frequency because of multiple eddy chopping. Trailing edge noise produces a broad spectrum. For skewed inflow to a rotor, e.g., a helicopter in forward flight, narrow band tones rapidly degenerate because of the turbulent eddies in the rotor plane. Theory and measurement agree well for helicopters, but not as closely as for airfoils.

  17. 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. PMID:24116405

  18. 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.

  19. Task 4 supporting technology. Part 1: Detailed test plan for leading edge tile development. Leading edge material development and testing

    NASA Technical Reports Server (NTRS)

    Hogenson, P. A.; Staszak, Paul; Hinkle, Karrie

    1995-01-01

    This task develops two alternative candidate tile materials for leading edge applications: coated alumina enhanced thermal barrier (AETB) tile and silicone impregnated reusable ceramic ablator (SIRCA) tile. Upon reentry of the X-33/RLV space vehicle, the leading edges experience the highest heating rates and temperatures. The wing leading edge and nose cap experience peak temperatures in the range 2000 to 2700 F. Replacing reinforced carbon-carbon (RCC) with tile-based thermal protection system (TPS) materials is the primary objective. Weight, complexity, coating impact damage, and repairability are among the problems that this tile technology development addresses. The following subtasks will be performed in this development effort: tile coating development; SIRCA tile development; robustness testing of tiles; tile repair development; tile operations/processing; tile leading edge configuration; and life cycle testing.

  20. Task 4 supporting technology. Part 1: Detailed test plan for leading edge tile development. Leading edge material development and testing

    NASA Astrophysics Data System (ADS)

    Hogenson, P. A.; Staszak, Paul; Hinkle, Karrie

    1995-05-01

    This task develops two alternative candidate tile materials for leading edge applications: coated alumina enhanced thermal barrier (AETB) tile and silicone impregnated reusable ceramic ablator (SIRCA) tile. Upon reentry of the X-33/RLV space vehicle, the leading edges experience the highest heating rates and temperatures. The wing leading edge and nose cap experience peak temperatures in the range 2000 to 2700 F. Replacing reinforced carbon-carbon (RCC) with tile-based thermal protection system (TPS) materials is the primary objective. Weight, complexity, coating impact damage, and repairability are among the problems that this tile technology development addresses. The following subtasks will be performed in this development effort: tile coating development; SIRCA tile development; robustness testing of tiles; tile repair development; tile operations/processing; tile leading edge configuration; and life cycle testing.

  1. 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.

  2. Design and Analysis of UHTC Leading Edge Attachment

    NASA Technical Reports Server (NTRS)

    Thomas, David J.; Nemeth, Noel N. (Technical Monitor)

    2002-01-01

    NASA Glenn Research Center was contacted to provide technical support to NASA Ames Research Center in the design and analysis of an ultra high temperature ceramic (UHTC) leading edge. UHTC materials are being considered for reusable launch vehicles because their high temperature capability may allow for un-cooled sharp leading edge designs. While ceramic materials have the design benefit of allowing subcomponents to run hot, they also provide a design challenge in that they invariably must be in contact with cooler subcomponents elsewhere in the structure. NASA Glenn Research Center proposed a modification to an existing attachment design. Thermal and structural analyses of the leading edge assembly were carried out using ABAQUS finite element software. Final results showed that the proposed modifications aided in thermally isolating hot and cold subcomponents and reducing bearing stresses at the attachment location.

  3. Fluid-thermal-structural study of aerodynamically heated leading edges

    NASA Technical Reports Server (NTRS)

    Deuchamphai, Pramote; Thornton, Earl A.; Wieting, Allan R.

    1988-01-01

    A finite element approach for integrated fluid-thermal-structural analysis of aerodynamically heated leading edges is presented. The Navier-Stokes equations for high speed compressible flow, the energy equation, and the quasi-static equilibrium equations for the leading edge are solved using a single finite element approach in one integrated, vectorized computer program called LIFTS. The fluid-thermal-structural coupling is studied for Mach 6.47 flow over a 3-in diam cylinder for which the flow behavior and the aerothermal loads are calibrated by experimental data. Issues of the thermal-structural response are studied for hydrogen-cooled, super thermal conducting leading edges subjected to intense aerodynamic heating.

  4. 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.

  5. 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

  6. 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.

  7. 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.

  8. Digitally Literate Teachers in Leading Edge Schools in Norway

    ERIC Educational Resources Information Center

    Almas, Aslaug Grov; Krumsvik, Rune

    2007-01-01

    This paper highlights digitally literate, in-service teachers in leading edge schools in Norway and focuses on how they, in their professional development, adapt digital literacy. Today we find a consensus among policy-makers, researchers, teacher-educators and teachers that digital literacy must be given high priority and needs to be explored…

  9. 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.

  10. 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

  11. Experimental Aerothermodynamics In Support Of The Columbia Accident Investigation

    NASA Technical Reports Server (NTRS)

    Horvath, Thomas J.

    2004-01-01

    The technical foundation for the most probable damage scenario reported in the Columbia Accident Investigation Board's final report was largely derived from synergistic aerodynamic/aerothermodynamic wind tunnel measurements and inviscid predictions made at NASA Langley Research Center and later corroborated with engineering analysis, high fidelity numerical viscous simulations, and foam impact testing near the close of the investigation. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at NASA Langley and illustrates how the ground-based heating measurements provided early insight that guided the direction and utilization of agency resources in support of the investigation. Global surface heat transfer mappings, surface streamline patterns, and shock shapes were measured on 0.0075 scale models of the Orbiter configuration with and without postulated damage to the thermal protection system. Test parametrics include angle of attack from 38 to 42 degs, sideslip angles of 38 to 42 degs, sideslip angles of plus or minus 1 deg, Reynolds numbers based upon model length from 0.05 x 10(exp 6) to 6.5 x 10(exp 6), and normal shock density ratios of 5 (Mach 6 Air) and 12 (Mach 6 CF4). The primary objective of the testing was to provide surface heating characteristics on scaled Orbiter models with outer mold line perturbations to simulate various forms of localized surface damage to the thermal protection system. Initial experimental testing conducted within two weeks of the accident simulated a broad spectrum of thermal protection system damage to the Orbiter windward surface and was used to refute several hypothesized forms of thermal protection system damage, which included gouges in the windward thermal protection system tiles, breaches through the wing new the main landing gear door, and protuberances along the wing leading edge that produced asymmetric boundary layer transition. As the forensic phase of the investigation

  12. Leading-edge slat optimization for maximum airfoil lift

    NASA Technical Reports Server (NTRS)

    Olson, L. E.; Mcgowan, P. R.; Guest, C. J.

    1979-01-01

    A numerical procedure for determining the position (horizontal location, vertical location, and deflection) of a leading edge slat that maximizes the lift of multielement airfoils is presented. The structure of the flow field is calculated by iteratively coupling potential flow and boundary layer analysis. This aerodynamic calculation is combined with a constrained function minimization analysis to determine the position of a leading edge slat so that the suction peak on the nose of the main airfoil is minized. The slat position is constrained by the numerical procedure to ensure an attached boundary layer on the upper surface of the slat and to ensure negligible interaction between the slat wake and the boundary layer on the upper surface of the main airfoil. The highest angle attack at which this optimized slat position can maintain attached flow on the main airfoil defines the optimum slat position for maximum lift. The design method is demonstrated for an airfoil equipped with a leading-edge slat and a trailing edge, single-slotted flap. The theoretical results are compared with experimental data, obtained in the Ames 40 by 80 Foot Wind Tunnel, to verify experimentally the predicted slat position for maximum lift. The experimentally optimized slat position is in good agreement with the theoretical prediction, indicating that the theoretical procedure is a feasible design method.

  13. 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

  14. Aerothermodynamic heating and performance analysis of a high-lift aeromaneuvering AOTV concept

    NASA Technical Reports Server (NTRS)

    Menees, G. P.; Brown, K. G.; Wilson, J. F.; Davies, C. B.

    1985-01-01

    The thermal-control requirements for design-optimized aeromaneuvering performance are determined for space-based applications and low-earth orbit sorties involving large, multiple plane-inclination changes. The leading-edge heating analysis is the most advanced developed for hypersonic-rarefied flow over lifting surfaces at incidence. The effects of leading-edge bluntness, low-density viscous phenomena, and finite-rate flow-field chemistry and surface catalysis are accounted for. The predicted aerothermodynamic heating characteristics are correlated with thermal-control and flight-performance capabilities. The mission payload capability for delivery, retrieval, and combined operations is determined for round-trip sorties extending to polar orbits. Recommendations are given for future design refinements. The results help to identify technology issues required to develop prototype operational systems.

  15. Space Shuttle hypersonic aerodynamic and aerothermodynamic flight research and the comparison to ground test results

    NASA Technical Reports Server (NTRS)

    Iliff, Kenneth W.; Shafer, Mary F.

    1993-01-01

    Aerodynamic and aerothermodynamic comparisons between flight and ground test for the Space Shuttle at hypersonic speeds are discussed. All of the comparisons are taken from papers published by researchers active in the Space Shuttle program. The aerodynamic comparisons include stability and control derivatives, center-of-pressure location, and reaction control jet interaction. Comparisons are also discussed for various forms of heating, including catalytic, boundary layer, top centerline, side fuselage, OMS pod, wing leading edge, and shock interaction. The jet interaction and center-of-pressure location flight values exceeded not only the predictions but also the uncertainties of the predictions. Predictions were significantly exceeded for the heating caused by the vortex impingement on the OMS pods and for heating caused by the wing leading-edge shock interaction.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. At the leading edge of three-dimensional cell migration

    PubMed Central

    Petrie, Ryan J.; Yamada, Kenneth M.

    2012-01-01

    Summary Cells migrating on flat two-dimensional (2D) surfaces use actin polymerization to extend the leading edge of the plasma membrane during lamellipodia-based migration. This mode of migration is not universal; it represents only one of several mechanisms of cell motility in three-dimensional (3D) environments. The distinct modes of 3D migration are strongly dependent on the physical properties of the extracellular matrix, and they can be distinguished by the structure of the leading edge and the degree of matrix adhesion. How are these distinct modes of cell motility in 3D environments related to each other and regulated? Recent studies show that the same type of cell migrating in 3D extracellular matrix can switch between different leading edge structures. This mode-switching behavior, or plasticity, by a single cell suggests that the apparent diversity of motility mechanisms is integrated by a common intracellular signaling pathway that governs the mode of cell migration. In this Commentary, we propose that the mode of 3D cell migration is governed by a signaling axis involving cell–matrix adhesions, RhoA signaling and actomyosin contractility, and that this might represent a universal mechanism that controls 3D cell migration. PMID:23378019

  1. Leading-edge vortex solutions with large total pressure losses

    NASA Technical Reports Server (NTRS)

    Murman, Earll M.; Powell, Kenneth G.; Goodsell, Aga M.; Landahl, Marten T.

    1987-01-01

    Computations are presented for a Lambda = 75 deg delta wing in a supersonic freestream under two conditions which lead to leading-edge vortices. For one condition, analysis of the computed vortical flow reveals a closed streamline in the core. From varying computational parameters, it appears that this is due to truncation error of the convective derivatives. For the other condition, comparisons are made with wind-tunnel data, and good agreement is noted for pitot pressure distributions, flow angles on the symmetry plane, and the position of an embedded shock. Many of the aerodynamic parameters are shown to be insensitive to grid spacing.

  2. Optimization of the leading edge segment of a corrugated wing

    NASA Astrophysics Data System (ADS)

    Khurana, Manas; Chahl, Javaan

    2014-03-01

    Insect wings consist of flat plates of membranes stiffened by spars. The effect of this structure is that the wings appear as corrugated surfaces when considered on chordwise sections. We know that aerodynamically efficient insects such as a dragonfly engage in fixed wing flight modes for extended periods. The analysis in the literature has shown that the aerodynamic efficiency (cl/cd) of a corrugated aerofoil is sensitive to Reynolds number (Re) and angle-of-attack (AoA), yet the conclusions established are on the basis of flow analysis on a single baseline shape only. The sample size of the aerofoils must be extended further so that the influence and merits of corrugated shape features can be established. In this work, a design-of-experiments (DoE) approach is applied to induce systematic shape perturbations on a select, off-the-shelf baseline shape one feature at a time over a set number of increments. At each shape increment, the aerodynamic forces are established using a high fidelity CFD solver. The design space is modeled at a Re of 20,000 and 34,000 and at flow angle of 4.0° to represent a Micro Air Vehicle (MAV) in glide. The results confirmed the importance of the leading and trailing edge deflections on cl/cd. At Re = 20, 000, cl/cd of a corrugated aerofoil with deflection at the leading edge region only is 16% higher than the baseline shape, and 39% higher than the flat plate. At Re = 34, 000, cl/cd performance is sensitive to the trailing edge deflection. At the optimum deflection setting, cl/cd is 18% higher than the baseline shape and 23% higher than the flat plate. The results confirm that the leading and trailing edge deflections are critical to cl/cd for a MAV in glide.

  3. Computational aeroacoustic simulations of leading-edge slat flow

    NASA Astrophysics Data System (ADS)

    Takeda, K.; Zhang, X.; Nelson, P. A.

    2004-02-01

    High-lift devices on modern airliners are a major contributor to overall airframe noise. In this paper the aeroacoustics of leading-edge slat devices in a high-lift configuration are investigated computationally. A hierarchical methodology is used to enable the rapid evaluation of different slat configurations. The overall goal is to gain a deeper understanding of the noise generation and amplification mechanisms in and around the slat, and the effects of slat system geometry. In order to perform parametric studies of the aeroacoustics, a simplified 2-D model of the slat is used. The flow and aeroacoustics are computed using a compressible, unsteady, Reynolds-Averaged Navier-Stokes code. A robust buffer zone boundary condition is used to prevent the reflection of outgoing acoustic waves from contaminating the long-time solution. A Ffowcs Williams-Hawkings solver is used to compute the far field acoustic field from the unsteady flow solution and determine the directivity. The spanwise correlation length used is derived from experimental data of this high-lift configuration. The effect of spanwise correlation length on the acoustic far field is examined. The aeroacoustics of the slat system are largely governed by the geometry, especially in terms of slat overlap. We perform a study of the effects of trailing edge thickness, horizontal and vertical overlap settings for the slat on near field wave propagation and far field directivity. The implications for low-noise leading edge slat design are discussed.

  4. Leading edge vortex dynamics on a pitching delta wing

    NASA Technical Reports Server (NTRS)

    Lemay, S. P.; Batill, S. M.; Nelson, R. C.

    1988-01-01

    A study of the dynamic behavior of the leading edge vortices on a delta wing undergoing oscillatory pitching motion is presented. A sharp edge, flat plate, delta wing having a sweep angle of 70 deg was used in this investigation. The wing was sinusoidally pitched about its 1/2 chord position at reduced frequencies ranging from k = 2(pi)fc/u = 0.05 to 0.30 at chord Reynolds numbers between 90,000 and 350,000, for angle of attack ranges of 29 to 39 deg and 0 to 45 deg. During these dynamic motions, visualization of the leading edge vortices was obtained by marking the vortices with TiCl4 introduced through ports located near the model apex. The location of vortex breakdown was recorded using high speed motion picture photography. The motion picture records were analyzed to determine vortex trajectory and breakdown position as a function of angle of attack. When the wing was sinusoidally pitched, a hysteresis was observed in the location of breakdown position. This hysteresis increased with reduced frequency. The velocity of breakdown propagation along the wing, and the phase lag between model motion and breakdown location were also determined. Detailed information was also obtained on the oscillation of breakdown position in both static and dynamic cases.

  5. 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.

  6. Generation of instability waves at a leading edge

    NASA Technical Reports Server (NTRS)

    Goldstein, M. E.

    1982-01-01

    This paper describes the generation of instability waves downstream of a leading edge by an imposed upstream disturbance. Two cases are considered. The first is concerned with mean flows of the Blasius type wherein the instabilities are represented by Tollmien-Schlichting waves. It is shown that the latter are generated fairly far downstream of the edge and are the result of a wave length reduction process that tunes the free stream disturbances to the Tollmien-Schlichting wave length. The other case is concerned with inflectional, uni-directional, transversely sheared mean flows. Such idealized flows provide a fairly good local representation to the nearly parallel flows in jets. They can support inviscid instabilities of the Kelvin-Helmholtz type. The various mathematically permissible mechanisms that can couple these instabilities to the upstream disturbances are discussed. The results are compared to some acoustic measurements and conclusions are drawn about the generation of the instabilities in these flows.

  7. 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; Lester, Dean

    2012-01-01

    In response to the Columbia Accident Investigation Board report, a plug repair kit has been developed to enable astronauts to repair the space shuttle's wing leading edge (WLE) during orbit. The plug repair kit consists of several 17.78- cm-diameter carbon/silicon carbide (C/SiC) cover plates of various curvatures that can be attached to the refractory carbon-carbon WLE panels using a TZM refractory metal attach mechanism. The attach mechanism is inserted through the damage in the WLE panel and, as it is tightened, the cover plate flexes to conform to the curvature of the WLE panel within 0.050 mm. An astronaut installs the repair during an extravehicular activity (EVA). After installing the plug repair, edge gaps are checked and the perimeter of the repair is sealed using a proprietary material, developed to fill cracks and small holes in the WLE.

  8. Pulsed film cooling on a turbine blade leading edge

    NASA Astrophysics Data System (ADS)

    Rutledge, James L.

    2009-12-01

    Unsteadiness in gas turbine film cooling jets may arise due to inherent unsteadiness of the flow through an engine or may be induced as a means of flow control. The traditional technique used to evaluate the performance of a steady film cooling scheme is demonstrated to be insufficient for use with unsteady film cooling and is modified to account for the cross coupling of the time dependent adiabatic effectiveness and heat transfer coefficient. The addition of a single term to the traditional steady form of the net heat flux reduction equation with time averaged quantities accounts for the unsteady effects. An experimental technique to account for the influence of the new term was devised and used to measure the influence of a pulsating jet on the net heat flux in the leading edge region of a turbine blade. High spatial resolution data was acquired in the near-hole region using infrared thermography coupled with experimental techniques that allowed application of the appropriate thermal boundary conditions immediately adjacent to the film cooling hole. The turbine blade leading edge was simulated by a half cylinder in cross flow with a blunt afterbody. The film cooling geometry consisted of a coolant hole located 21.5° from the leading edge, angled 20° to the surface and 90° from the streamwise direction. Investigated parameters include pulsation frequency, duty cycle, and waveform shape. Separate experiments were conducted in a water channel to provide visualization of the unsteady coolant propagation behavior. Further insight into the flow physics was obtained through computational simulations of the experimental apparatus. The computational results afforded time resolved flow field and net heat flux reduction data unobtainable with the experimental techniques. A technique to predict the performance of an unsteady film cooling scheme through knowledge of only the steady film cooling behavior was developed and demonstrated to be effective.

  9. Method Improvements in Thermal Analysis of Mach 10 Leading Edges

    NASA Technical Reports Server (NTRS)

    Amundsen, Ruth M.

    2001-01-01

    Several improvements have recently been made in the thermal analysis methods for leading edges of a hypersonic vehicle. The leading edges of this vehicle undergo exceptionally high heat loads that incorporate extreme spatial gradients as well as severe transients. Due to the varying flight conditions, complex geometry, and need for thermal loads at many points along the trajectory, full computational fluid dynamics (CFD) analysis of the aeroheating; loads is not feasible. Thus, engineering methods must be used to determine the aeroheating on the vehicle surfaces, and that must be utilized in the thermal analysis. Over the last year, the thermal analysis of a hypersonic vehicle has been enhanced in several ways. Two different engineering codes are used to predict aeroheating loads: one over the curve near the stagnation point, and the other on flat surfaces downstream of the leading edge. These two are matched together at the intersection point using a method that allows closer approximation of CFD results. User-developed FORTRAN, which is part of the thermal solver PATRAN Thermal, is used to accomplish this. The customizable FORTRAN code also allows use of many different time- and space-dependent factors, interpolation of the heat load in time and space, and inclusion of both highly swept and unswept grid structures. This FORTRAN is available to other PATRAN users who may want to accomplish a similar objective in analysis. Flux, rather than convective coefficient, is used to define heat loads, which allows more accurate analysis as well as better application of margins. Improvements have also been made in more efficient utilization of imported CAD geometry, by creating faces on solids to facilitate load application.

  10. 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

    The feasibility of two candidate leading-edge flow laminarization systems applicable to airline service was tested using representative airline operational conditions with respect to air traffic, weather, and airport insect infestation. One of the systems involved a perforated Ti alloy suction surface with about 1 million 0.0025-in. diameter holes drilled by electron beam, as well as a Krueger-type flap that offered protective shielding against insect impingement; the other supplied surface suction through a slotted Ti alloy skin with 27 spanwise slots on the upper and lower surface.

  11. Characteristics of deformable leading edge for high performance helicopter rotor

    NASA Technical Reports Server (NTRS)

    Lee, Soogab; Mcalister, K. W.; Tung, Chee

    1993-01-01

    The deformable leading edge (DLE) concept to improve the blade capability in lift, drag and pitching moments has been investigated for the purpose of meeting new rotor maneuverability and susceptibility requirements. The advantages and disadvantages of this concept have been carefully examined with limited computational and experimental results. This work showed that this concept achieves a substantial improvement in lift capability and also reduces the drag and pitching moment at the same time. Effects of various parameters, such as Reynolds number, reduced frequency, mean angle of oscillation, and airfoil shape, on the performance of these airfoils were also investigated.

  12. Measurements of leading edge vortices in a supersonic stream

    NASA Astrophysics Data System (ADS)

    Milanovic, Ivana Milija

    An experimental investigation of the leading edge vortices from a 75° sweptback, sharp edge delta wing has been carried out in a Mach 2.49 stream. Five-hole conical probe traverses were conducted vertically and horizontally through the primary vortices at the trailing edge and at one half chord downstream station for 7° and 12° angles of attack. The main objective was to determine the Mach number and pressure distributions in the primary vortex and to present comparisons of flow properties at different survey stations. In response to the continued interest in efficient supersonic flight vehicles, particularly in the missile arena, the motivation for this research has been to provide the quantitative details of supersonic leading edge vortices, the understanding of which up to now has been largely based on flow visualizations and presumed similarity to low speed flows. As a prerequisite to the measurement campaign, the employed five-hole conical probe was numerically calibrated using a three-dimensional Thin Layer Navier-Stokes solver in order to circumvent the traditional experimental approach vastly demanding on resources. The pressure readings at the probe orifices were computed for a range of Mach numbers and pitch angles, and subsequently verified in wind tunnel tests. The calibration phase also demonstrated the profound influence of the probe bluntness on the nearby static pressure ports, its relevance to the ultimate modeling strategy and the resulting calibration charts. Flow diagnostics of the leading edge vortices included both qualitative flow visualizations, as well as quantitative measurements. Shadowgraphs provided information regarding the trajectory and relative size of the generated vortices while assuring that no probe-induced vortex breakdown occurred. Surface oil patterns revealed the general spanwise locations of leeward vortices, and confirmed topological similarity to their low speed counterparts. The probe measurements revealed substantial

  13. 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.

  14. Separation Control on a Hydrofoil Using Leading Edge Protuberances

    NASA Astrophysics Data System (ADS)

    Custodio, Derrick; Henoch, Charles; Johari, Hamid

    2007-11-01

    The humpback whale's maneuvarability has been attributed to their use of pectoral flippers, on which protuberances are present along the leading edge. To examine the effects of protuberances on hydrofoil performance, the lift, drag, and pitching moments of two-dimensional hydrofoils with leading edge sinusoidal protuberances were measured in a water tunnel and compared to those of a baseline NACA 63(4)-021 hydrofoil. The amplitude and spanwise wavelengths of the protuberances ranged from 2.5% to 12% and 25% to 50% of the mean chord length respectively. Flow visualizations using tufts and dye, as well as Laser Doppler Velocimetry (LDV) measurements were performed to examine the flow patterns surrounding the hydrofoils. At angles of attack lower than the stall angle of the baseline, the modified foils revealed lower lift and increased drag. However, above this angle the lift generated by the modified foils was up to 50% greater than the baseline foil with little or no drag penalty. The amplitude of the protuberances has a large effect on the performance of the hydrofoils whereas the wavelength has little. Flow topology on the protuberances will be discussed by means of the visualization and measured velocities.

  15. An embedded mesh procedure for leading-edge vortex flows

    NASA Technical Reports Server (NTRS)

    Powell, Kenneth G.; Murman, Earll M.

    1989-01-01

    A cell-vertex scheme is outlined for solving the flow about a delta wing with M (sub infinity) is greater than 1. Embedded regions of mesh refinement allow solutions to be obtained which have much higher resolution than those achieved to date. Effects of mesh refinement and artificial viscosity on the solutions are studied, to determine at what point leading-edge vortex solutions are grid-converged. A macroscale and a microscale for the size of the vortex are defined, and it is shown that the macroscale (which includes the wing surface properties) is converged on a moderately refined grid, while the microscale is very sensitive to grid spacing. The level of numerical diffusion in the core of the vortex is found to be substantial. Comparisons with the experiment are made for two cases which have transonic cross-flow velocities.

  16. Particle rebound characteristics of turbomachinery cascade leading edge geometry

    NASA Astrophysics Data System (ADS)

    Siravuri, Sastri

    The objective of this research work is to investigate and understand the complex phenomena associated with the mechanism of particle impacts on turbomachinery cascade leading edge geometry. At present, there is a need for experimental work in basic and applied research to find out the parameters that are relevant to particle rebound characteristics on turbomachinery blades. In the present work, experiments were conducted with air velocity at 15 m/s (˜50 ft/sec) and at 30 m/s (˜100 ft/sec) using high-speed photography and Laser Doppler Velocimetry (LDV). Silica sand particles of 1000--1500 micron size were used for this study. In the present investigation, particle rebound data was obtained for cylindrical targets with radius of curvature representative of leading edge geometry (cylinder diameter = 4.5mm & 6.5 mm) using LDV. The numerical simulations, which are based on non-linear dynamic analysis, were also performed using the finite element code DYNA3-D. Several different material models viz elastic-elastic, elastic-plastic, elastic-plastic with friction & isotropic-elastic-plastic with dynamic friction and particle rotation were used in the DYNA3-D numerical analysis. The computational results include a time history of the displacement, stress and strain profiles through the particle collision. Numerical results are presented for the rebound conditions of spherical silica sand particle for different pre-collision velocities. The computed particle restitution coefficients, after they reach steady rebound conditions, are compared with experimental results obtained from LDV. A probabilistic model was developed to incorporate the uncertainties in the impact velocity in the numerical model. Histograms and Cumulative Distribution Functions (CDFs) for impact velocity were obtained from experimental LDV data. Ten randomly selected probabilities for each impact angle were used to calculate the impact velocity from cumulative distribution function. This randomly selected

  17. 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.

  18. Managed aquifer recharge: rediscovering nature as a leading edge technology.

    PubMed

    Dillon, P; Toze, S; Page, D; Vanderzalm, J; Bekele, E; Sidhu, J; Rinck-Pfeiffer, S

    2010-01-01

    Use of Managed Aquifer Recharge (MAR) has rapidly increased in Australia, USA, and Europe in recent years as an efficient means of recycling stormwater or treated sewage effluent for non-potable and indirect potable reuse in urban and rural areas. Yet aquifers have been relied on knowingly for water storage and unwittingly for water treatment for millennia. Hence if 'leading edge' is defined as 'the foremost part of a trend; a vanguard', it would be misleading to claim managed aquifer recharge as a leading edge technology. However it has taken a significant investment in scientific research in recent years to demonstrate the effectiveness of aquifers as sustainable treatment systems to enable managed aquifer recharge to be recognised along side engineered treatment systems in water recycling. It is a 'cross-over' technology that is applicable to water and wastewater treatment and makes use of passive low energy processes to spectacularly reduce the energy requirements for water supply. It is robust within limits, has low cost, is suitable from village to city scale supplies, and offers as yet almost untapped opportunities for producing safe drinking water supplies where they do not yet exist. It will have an increasingly valued role in securing water supplies to sustain cities affected by climate change and population growth. However it is not a universal panacea and relies on the presence of suitable aquifers and sources of water together with effective governance to ensure human health and environment protection and water resources planning and management. This paper describes managed aquifer recharge, illustrates its use in Australia, outlining economics, guidelines and policies, and presents some of the knowledge about aquifer treatment processes that are revealing the latent value of aquifers as urban water infrastructure and provide a driver to improving our understanding of urban hydrogeology. PMID:21076220

  19. 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

  20. 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.

  1. Mechanisms of leading edge protrusion in interstitial migration

    NASA Astrophysics Data System (ADS)

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

    2013-12-01

    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.

  2. 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.

  3. Turbulent Coherence Measurements on a Leading Edge Slat

    NASA Astrophysics Data System (ADS)

    Moriarty, Patrick J.; Storms, Bruce L.; Ross, James C.; Horne, W. Clifton; Dougherty, Robert P.

    1997-11-01

    Turbulence spectra have been measured downstream of the gap between the leading-edge slat and the main airfoil of a generic transport aircraft wing model. The model consisted of a NACA 63_2-215 Mod. B main element with a half span Fowler flap and a full span LB-546 slat. Velocity-pressure coherence spectra were determined using signals from a hot-wire anemometer probe in the flow and unsteady-pressure transducers on the wing surface. The coherence coefficient was significant only in a narrow bandwidth, of the order of 15 percent of the peak frequency. Coherence coefficient magnitudes were as large as 0.25. Coherence coefficient magnitude and frequency were found to depend on the flow velocity through the slat gap, which increased with (negative) slat deflection angle. Frequencies and relative strengths of coherence peaks were in agreement with those of radiated noise spectra (measured in a separate experiment). The results demonstrate a close connection between velocity fluctuation in the slat wake and pressure fluctuations on both surfaces of the slat and the upper surface of the main airfoil. Further work is planned to investigate a postulated hydrodynamic-acoustic resonance.

  4. Effects of leading-edge tubercles on wing flutter speeds.

    PubMed

    Ng, B F; New, T H; Palacios, R

    2016-06-01

    The dynamic aeroelastic effects on wings modified with bio-inspired leading-edge (LE) tubercles are examined in this study. We adopt a state-space aeroelastic model via the coupling of unsteady vortex-lattice method and a composite beam to evaluate stability margins as a result of LE tubercles on a generic wing. The unsteady aerodynamics and spanwise mass variations due to LE tubercles have counteracting effects on stability margins with the former having dominant influence. When coupled, flutter speed is observed to be 5% higher, and this is accompanied by close to 6% decrease in reduced frequencies as an indication of lower structural stiffness requirements for wings with LE tubercles. Both tubercle amplitude and wavelength have similar influences over the change in flutter speeds, and such modifications to the LE would have minimal effect on stability margins when concentrated inboard of the wing. Lastly, when used in sweptback wings, LE tubercles are observed to have smaller impacts on stability margins as the sweep angle is increased. PMID:27070824

  5. Unsteady flow phenomena associated with leading-edge vortices

    NASA Astrophysics Data System (ADS)

    Breitsamter, C.

    2008-01-01

    This paper presents selected results from extensive experimental investigations on turbulent flow fields and unsteady surface pressures caused by leading-edge vortices, in particular, for vortex breakdown flow. Such turbulent flows may cause severe dynamic aeroelastic problems like wing and/or fin buffeting on fighter-type aircraft. The wind tunnel models used include a generic delta wing as well as a detailed aircraft configuration of canard-delta wing type. The turbulent flow structures are analyzed by root-mean-square and spectral distributions of velocity and pressure fluctuations. Downstream of bursting local maxima of velocity fluctuations occur in a limited radial range around the vortex center. The corresponding spectra exhibit significant peaks indicating that turbulent kinetic energy is channeled into a narrow band. These quasi-periodic velocity oscillations arise from a helical mode instability of the breakdown flow. Due to vortex bursting there is a characteristic increase in surface pressure fluctuations with increasing angle of attack, especially when the burst location moves closer to the apex. The pressure fluctuations also show dominant frequencies corresponding to those of the velocity fluctuations. Using the measured flow field data, scaling parameters are derived for design purposes. It is shown that a frequency parameter based on the local semi-span and the sinus of angle of attack can be used to estimate the frequencies of dynamic loads evoked by vortex bursting.

  6. Leading edge vortex in a slow-flying passerine

    PubMed Central

    Muijres, Florian T.; Johansson, L. Christoffer; Hedenström, Anders

    2012-01-01

    Most hovering animals, such as insects and hummingbirds, enhance lift by producing leading edge vortices (LEVs) and by using both the downstroke and upstroke for lift production. By contrast, most hovering passerine birds primarily use the downstroke to generate lift. To compensate for the nearly inactive upstroke, weight support during the downstroke needs to be relatively higher in passerines when compared with, e.g. hummingbirds. Here we show, by capturing the airflow around the wing of a freely flying pied flycatcher, that passerines may use LEVs during the downstroke to increase lift. The LEV contributes up to 49 per cent to weight support, which is three times higher than in hummingbirds, suggesting that avian hoverers compensate for the nearly inactive upstroke by generating stronger LEVs. Contrary to other animals, the LEV strength in the flycatcher is lowest near the wing tip, instead of highest. This is correlated with a spanwise reduction of the wing's angle-of-attack, partly owing to upward bending of primary feathers. We suggest that this helps to delay bursting and shedding of the particularly strong LEV in passerines. PMID:22417792

  7. 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.

  8. Forebody and leading edge vortex measurements using planar Doppler velocimetry

    NASA Astrophysics Data System (ADS)

    Beutner, Thomas J.; Elliott, Gregory S.; Williams, Glenn W.; Baust, Henry D.; Crafton, Jim; Carter, Campbell D.

    2001-04-01

    The planar Doppler velocimetry (PDV) technique has been demonstrated by employing it in a large-scale wind tunnel to record velocity fields surrounding a model of a generic fighter plane. The PDV instrument employed here included the following: (i) a frequency monitoring system for measuring the laser frequency corresponding to each set of scattering images; (ii) two detector systems (each composed of two 16-bit CCD cameras), one viewing the model from the top of the wind tunnel and the second from the side; (iii) iodine vapour cells based on the starved-cell design, which eliminated the need for separate temperature control of the iodine reservoir; iv) a vibration-isolated, injection-seeded, Q-switched Nd:YAG laser and (v) custom data acquisition software for linking the four cameras, the laser and the frequency monitor. The PDV instrument was validated by comparing the PDV-derived velocity to the known value in the empty wind tunnel. An error of about 1 m s-1 out of an 18.9 m s-1 velocity component was found; the image noise component (resulting primarily from the speckle effect) was found to be about 1 m s-1. In addition, as a result of laser-sheet impingement on the model surface, velocities near the model surfaces are biased by background scattering effects. Nonetheless, it has been shown that PDV can be used effectively to map velocity fields with high spatial resolution over complex model geometries. Frame-averaged velocity images recorded at four axial stations along the model have shown the formation of forebody and leading-edge vortices and their complex interaction in the presence of the wing flow field.

  9. 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

  10. 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

  11. Trailing edges projected to move faster than leading edges for large pelagic fish habitats under climate change

    NASA Astrophysics Data System (ADS)

    Robinson, L. M.; Hobday, A. J.; Possingham, H. P.; Richardson, A. J.

    2015-03-01

    There is mounting evidence to suggest that many species are shifting their ranges in concordance with the climate velocity of their preferred environmental conditions/habitat. While accelerated rates in species' range shifts have been noted in areas of intense warming, due to climate change, few studies have considered the influence that both spatial temperature gradients and rates of warming (i.e., the two components of climate velocity) could have on rates of movement in species habitats. We compared projected shifts in the core habitat of nine large pelagic fish species (five tuna, two billfish and two shark species) off the east coast of Australia at different spatial points (centre, leading and trailing edges of the core habitat), during different seasons (summer and winter), in the near-(2030) and long-term (2070), using independent species distribution models and habitat suitability models. Model projections incorporated depth integrated temperature data from 11 climate models with a focus on the IPCC SRES A2 general emission scenario. Projections showed a number of consistent patterns: southern (poleward) shifts in all species' core habitats; trailing edges shifted faster than leading edges; shifts were faster by 2070 than 2030; and there was little difference in shifts among species and between seasons. Averaging across all species and climate models, rates of habitat shifts for 2030 were 45-60 km decade-1 at the trailing edge, 40-45 km decade-1 at the centre, and 20-30 km decade-1 at the leading edge. Habitat shifts for 2070 were 60-70 km decade-1 at the trailing edge, 50-55 km decade-1 at the centre, and 30-40 km decade-1 at the leading edge. It is often assumed that the leading edge of a species range will shift faster than the trailing edge, but there are few projections or observations in large pelagic fish to validate this assumption. We found that projected shifts at the trailing edge were greater than at the centre and leading of core habitats in

  12. 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.

  13. Pressure measurements on the leading edge of a swept wing at Mach 2.2

    NASA Technical Reports Server (NTRS)

    Sorrells, R. B., III

    1974-01-01

    Detailed pressure measurements were made on a flat semispan swept wing with a rounded leading edge at Mach number 2.2 through a range of Reynolds numbers. Pressure orifices were distributed in the streamwise direction at five spanwise stations on the leading edge and on the upper and lower surfaces. No significant amount of leading-edge suction was found, but the pressures and integrated normal forces on the upper and lower surfaces indicate the presence of a vortex lift.

  14. A numerical model for the platelet heat-pipe-cooled leading edge of hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Liu, Hongpeng; Liu, Weiqiang

    2016-01-01

    A new design, the platelet heat-pipe-cooled leading edge, is discussed for the thermal management to prevent damage to hypersonic vehicle leading edge component. For calculating the steady state behavior of platelet heat-pipe-cooled leading edge, a numerical model based on the principles of evaporation, convection, and condensation of a working fluid is presented. And then its effectiveness is validated by comparing the wall and vapor temperature against experimental data for a conventional heat pipe. Further investigations indicate that alloy IN718, with sodium as the working fluid is a feasible combination for Mach 8 flight with a 15 mm leading edge radius.

  15. Combined overlay, focus and CD metrology for leading edge lithography

    NASA Astrophysics Data System (ADS)

    Ebert, Martin; Cramer, Hugo; Tel, Wim; Kubis, Michael; Megens, Henry

    2011-04-01

    As leading edge lithography moves to 22-nm design rules, low k1 technologies like double patterning are the new resolution enablers, and system control and setup are the new drivers to meet remarkably tight process requirements. The way of thinking and executing setup and control of lithography scanners is changing in four ways. First, unusually tight process tolerances call for very dense sampling [1], which in effect means measurements at high throughput combined with high order modeling and corrections to compensate for wafer spatial fingerprint. Second, complex interactions between scanner and process no longer allow separation of error sources through traditional metrology approaches, which are based on using one set of metrology tools and methods for setup and another for scanner performance control. Moreover, setup and control of overlay is done independently from CD uniformity, which in effect leads to independent and conflicting adjustments for the scanner. Third, traditional CD setup and control is based on the focus and dose calculated from their CD response and not from measurement of their effect on pattern profile, which allows a clean and orthogonal de-convolution of focus and dose variations across the wafer. Fourth, scanner setup and control has to take into consideration the final goal of lithography, which is the accurate printing of a complex pattern describing a real device layout. To this end we introduce a new setup and control metrology step: measuring-to-match scanner 1D and 2D proximity. In this paper we will describe the strategy for setup and control of overlay, focus, CD and proximity based on the YieldStarTM metrology tool and present the resulting performance. YieldStar-200 is a new, high throughput metrology tool based on a high numerical aperture scatterometer concept. The tool can be used stand-alone as well as integrated in a processing track. It is suitable for determining process offsets in X,Y and Z directions through Overlay

  16. 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.; Van Fossen, G. J.

    1992-01-01

    A study of the effect of spanwise variation on leading edge heat transfer is presented. Experimental and numerical results are given for a circular leading edge and for a 3:1 elliptical leading edge. It is demonstrated that increases in leading edge heat transfer due to spanwise variations in freestream momentum are comparable to those due to freestream turbulence.

  17. Aerothermodynamic Flight Simulation Capabilities for Aerospace Vehicles

    NASA Technical Reports Server (NTRS)

    Miller, Charles G.

    1998-01-01

    Aerothermodynamics, encompassing aerodynamics, aeroheating, and fluid dynamics and physical processes, is the genesis for the design and development of advanced space transportation vehicles and provides crucial information to other disciplines such as structures, materials, propulsion, avionics, and guidance, navigation and control. Sources of aerothermodynamic information are ground-based facilities, Computational Fluid Dynamic (CFD) and engineering computer codes, and flight experiments. Utilization of this aerothermodynamic triad provides the optimum aerothermodynamic design to safely satisfy mission requirements while reducing design conservatism, risk and cost. The iterative aerothermodynamic process for initial screening/assessment of aerospace vehicle concepts, optimization of aerolines to achieve/exceed mission requirements, and benchmark studies for final design and establishment of the flight data book are reviewed. Aerothermodynamic methodology centered on synergism between ground-based testing and CFD predictions is discussed for various flow regimes encountered by a vehicle entering the Earth s atmosphere from low Earth orbit. An overview of the resources/infrastructure required to provide accurate/creditable aerothermodynamic information in a timely manner is presented. Impacts on Langley s aerothermodynamic capabilities due to recent programmatic changes such as Center reorganization, downsizing, outsourcing, industry (as opposed to NASA) led programs, and so forth are discussed. Sample applications of these capabilities to high Agency priority, fast-paced programs such as Reusable Launch Vehicle (RLV)/X-33 Phases I and 11, X-34, Hyper-X and X-38 are presented and lessons learned discussed. Lastly, enhancements in ground-based testing/CFD capabilities necessary to partially/fully satisfy future requirements are addressed.

  18. Preparation and Support of a Tap Test on the Leading Edge Surfaces of the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Bohr, Jerry

    2009-01-01

    This slide presentation reports on a Tap test for the leading edge surfaces of the Space Shuttle. A description of the Wing Leading Edge Impact Detection System (WLEIDS) flight system is given, and the rationale and approach for improving the WLEIDS system. The three phases of the strategy of the test project amd the results of the tests are reviewed.

  19. 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 schools across the…

  20. 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 schools across the…

  1. Aerothermodynamic systems engineering and design

    NASA Astrophysics Data System (ADS)

    A reference source for various aspects of aerothermodynamic systems engineering and design is presented. Air conditioning load analysis is addressed, including physiological requirements, heat and cooling load equations, skin temperature computational methods, cooling loads due to radiation through transparent areas, heating and cooling loads due to internal sources, and practical considerations in the determination of overall heating and cooling loads. Refrigeration system design is considered, including air cycle systems, vapor cycle systems, combined vapor cycle and air cycle systems, and thermoelectric cooling. Heating methods is heating system design and low pressure and high pressure systems in air distribution system design are addressed. Procedures and equations commonly used for aerospace applications of these technologies are included.

  2. Aerothermodynamics of manned Mars missions

    NASA Technical Reports Server (NTRS)

    Park, Chul; Davies, Carol B.

    1989-01-01

    The aerothermodynamic problems associated with the aerobraking of the spacecraft proposed for the manned Mars mission are studied. The propulsive Delta V necessary at departure from earth and Mars and the velocities of the atmospheric entries into the two planets are deduced. It is shown that the propulsive Delta V can be reduced by increasing the entry velocities and that entry velocities up to about 15 km/sec are appropriate at both earth and Mars. L/D values of 0.8 and 2.0 are found to be necessary at earth and Mars, respectively. Density, pressure, and stagnation-point convective-heat-transfer rates are calculated for the typical aerobraking flights. Assuming the shock layer flow to be in equilibrium, the stagnation-point radiative-heat-transfer rates are calculated to be larger than the convective-heat-transfer rates. The possible impact of ablation, turbulence, and nonequilibrium are discussed.

  3. 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.

  4. Aerothermodynamics Overview and Prediction Assessment

    NASA Technical Reports Server (NTRS)

    Heidmann, James D.

    2007-01-01

    An overview of the Aerothermodynamics Discipline within NASA s Subsonic Fixed Wing Project is given. The primary focus of the presentation is on the research efforts conducted in fiscal year 2007. This year (2007), the work primarily consisted of efforts under level 1 (foundational research) and level 2 (tools and technology development). Examples of work under level 1 are large eddy simulation development, advanced turbine cooling concept development, and turbomachinery flow control development. Examples of level 2 research are the development of highly-loaded compressor and turbine test programs and advanced turbomachinery simulation development, including coupled inlet-fan simulations. An overview of the NRA research activity is also provided. This NRA focused on plasma and aspiration flow control for low pressure turbine application. Finally, a status report on the turbomachinery CFD code assessment activity is provided. This activity focuses on the use of several NASA in-house codes for the NASA rotor 37 and stage 35 test cases.

  5. Experimental study of unsteady aerothermodynamic phenomena on shock-tube wall using fast-response temperature-sensitive paints

    NASA Astrophysics Data System (ADS)

    Ozawa, Hiroshi

    2016-04-01

    This paper describes an experimental study that used a fast-response temperature-sensitive paint (TSP) to investigate the unsteady aerothermodynamic phenomena occurring on a shock-tube wall. To understand these phenomena in detail, a fast-response TSP with high temperature sensitivity developed for transient temperature measurement was applied to the wall. The shock-tube experiment was carried out under the over-tailored condition, with a pressure ratio of 110 for test gases of air in driver/driven tubes. The following aspects were clarified using the TSP: (a) the TSP could be used to visualize the unsteady aerothermodynamic phenomena and estimate the quantitative heat flux on the shock-tube wall; (b) an x-t diagram based on the TSP response showed shock-tube wall characteristics that included the incident/reflected shocks, laminar-to-turbulent boundary-layer transition, streaks in the turbulent boundary layer, reflected shock/turbulent boundary layer interaction, and waves reflected from a contact surface; (c) the TSP graphically showed that a transition front from the plate's leading edge and turbulent spots moved with 80% of the free-stream velocity behind the incident shock. In addition, the TSP could track the growth of the turbulent spots on the wall.

  6. Numerical study of the effect of tangential leading edge blowing on delta wing vortical flow

    NASA Technical Reports Server (NTRS)

    Yeh, David T.; Tavella, Domingo A.; Roberts, Leonard; Fujii, Kozo

    1989-01-01

    A numerical simulation of tangential blowing along the leading edge of a delta wing is analyzed as a means of controlling the position and strength of the leading-edge vortices. The computation is done by numerical solutions of the three-dimensional thin-layer Navier-Stokes equations. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of tangential leading-edge blowing at low to moderate angles of attack tends to reduce the pressure peaks associated with leading-edge vortices and to increase the suction peak around the leading edge, such that the integrated value of the surface pressure remains about the same.

  7. A method for computing the leading-edge suction in a higher-order panel method

    NASA Technical Reports Server (NTRS)

    Ehlers, F. E.; Manro, M. E.

    1984-01-01

    Experimental data show that the phenomenon of a separation induced leading edge vortex is influenced by the wing thickness and the shape of the leading edge. Both thickness and leading edge shape (rounded rather than point) delay the formation of a vortex. Existing computer programs used to predict the effect of a leading edge vortex do not include a procedure for determining whether or not a vortex actually exists. Studies under NASA Contract NAS1-15678 have shown that the vortex development can be predicted by using the relationship between the leading edge suction coefficient and the parabolic nose drag. The linear theory FLEXSTAB was used to calculate the leading edge suction coefficient. This report describes the development of a method for calculating leading edge suction using the capabilities of the higher order panel methods (exact boundary conditions). For a two dimensional case, numerical methods were developed using the double strength and downwash distribution along the chord. A Gaussian quadrature formula that directly incorporates the logarithmic singularity in the downwash distribution, at all panel edges, was found to be the best method.

  8. 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.

  9. Introduction: Assessment of aerothermodynamic flight prediction tools through ground and flight experimentation

    NASA Astrophysics Data System (ADS)

    Schmisseur, John D.; Erbland, Peter

    2012-01-01

    This article provides an introduction and overview to the efforts of NATO Research and Technology Organization Task Group AVT-136, Assessment of Aerothermodynamic Flight Prediction Tools through Ground and Flight Experimentation. During the period of 2006-2010, AVT-136 coordinated international contributions to assess the state-of-the-art and research challenges for the prediction of critical aerothermodynamic flight phenomena based on the extrapolation of ground test and numerical simulation. To achieve this goal, efforts were organized around six scientific topic areas: (1) Noses and leading edges, (2) Shock Interactions and Control Surfaces, (3) Shock Layers and Radiation, (4) Boundary Layer Transition, (5) Gas-Surface Interactions, and (6) Base and Afterbody Flows. A key component of the AVT-136 strategy was comparison of state-of-the-art numerical simulations with data to be acquired from planned flight research programs. Although it was recognized from the onset of AVT-136 activities that reliance on flight research data yet to be collected posed a significant risk, the group concluded the substantial benefit to be derived from comparison of computational simulations with flight data warranted pursuit of such a program of work. Unfortunately, program delays and failures in the flight programs contributing to the AVT-136 effort prevented timely access to flight research data. Despite this setback, most of the scientific topic areas developed by the Task Group made significant progress in the assessment of current capabilities. Additionally, the activities of AVT-136 generated substantial interest within the international scientific research community and the work of the Task Group was prominently featured in a total of six invited sessions in European and American technical conferences. In addition to this overview, reviews of the state-of-the-art and research challenges identified by the six research thrusts of AVT-136 are also included in this special

  10. Effects of leading and trailing edge flaps on the aerodynamics of airfoil/vortex interactions

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed A.; Sankar, L. N.; Tadghighi, H.

    1991-01-01

    A numerical procedure based on the unsteady 2D full potential equation is presently used to simulate the effects of leading-edge and trailing-edge flaps on the aerodynamics of airfoil-vortex interactions. Attention is given to unsteady flap-motion effects, which alleviate those interactions at sub- and supercritical onset flows. For subcritical interactions, the results obtained indicate that trailing-edge flaps can be used to alleviate the impulsive loads experienced by the airfoil; for supercritical interactions, a leading- rather than trailing-edge flap must be used to alleviate the interaction.

  11. Generation of instability waves at a leading edge

    NASA Technical Reports Server (NTRS)

    Goldstein, M. E.

    1982-01-01

    Two cases are considered. The first is concerned with mean flows of the Blasius type wherein the instabilities are represented by Tollmien-Schlichting waves. It is shown that the latter are generated fairly far downstream of the edge and are the result of a wave length reduction process that tunes the free stream disturbances to the Tollmien-Schlichting wave length. The other case is concerned with inflectional, uni-directional, transversely sheared mean flows. Such idealized flows provide a fairly good local representation to the nearly parallel flows in jets. They can support inviscid instabilities of the Kelvin-Helmholtz type. The various mathematically permissible mechanisms that can couple these instabilities to the upstream disturbances are discussed.

  12. Overview of the Aerothermodynamics Analysis Conducted in Support of the STS-107 Accident Investigation

    NASA Technical Reports Server (NTRS)

    Campbell, Charles H.

    2004-01-01

    A graphic presentation of the aerothermodynamics analysis conducted in support of the STS-107 accident investigation. Investigation efforts were conducted as part of an integrated AATS team (Aero, Aerothermal, Thermal, Stress) directed by OVEWG. Graphics presented are: STS-107 Entry trajectory and timeline (1st off-nominal event to Post-LOS); Indications from OI telemetry data; Aero/aerothermo/thermal analysis process; Selected STS-107 side fuselage/OMS pod off-nominal temperatures; Leading edge structural subsystem; Relevant forensics evidence; External aerothermal environments; STS-107 Pre-entry EOM3 heating profile; Surface heating and temperatures; Orbiter wing leading edge damage survey; Internal aerothermal environments; Orbiter wing CAD model; Aerodynamic flight reconstruction; Chronology of aerodynamic/aerothermoydynamic contributions; Acreage TPS tile damage; Larger OML perturbations; Missing RCC panel(s); Localized damage to RCC panel/missing T-seal; RCC breach with flow ingestion; and Aero-aerothermal closure. NAIT served as the interface between the CAIB and NASA investigation teams; and CAIB requests for study were addressed.

  13. Leading-edge vortex burst on a low-aspect-ratio rotating flat plate

    NASA Astrophysics Data System (ADS)

    Medina, Albert; Jones, Anya R.

    2016-08-01

    This study experimentally investigates the phenomenon of leading-edge-vortex burst on rotating flat plate wings. An aspect-ratio-2 wing was driven in pure rotation at a Reynolds number of Re=2500 . Of primary interest is the evolution of the leading-edge vortex along the wing span over a single-revolution wing stroke. Direct force measurements of the lift produced by the wing revealed a single global lift maximum relatively early in the wing stroke. Stereoscopic particle image velocimetry was applied to several chordwise planes to quantify the structure and strength of the leading-edge vortex and its effect on lift production. This analysis revealed opposite-sign vorticity entrainment into the core of the leading-edge vortex, originating from a layer of secondary vorticity along the wing surface. Coincident with the lift peak, there emerged both a concentration of opposite vorticity in the leading-edge-vortex core, as well as axial flow stagnation within the leading-edge-vortex core. Planar control volume analysis was performed at the midspan to quantify the contributions of vorticity transport mechanisms to the leading-edge-vortex circulation. The rate of circulation annihilation by opposite-signed vorticity entrainment was found to be minimal during peak lift production, where convection balanced the flux of vorticity resulting in stagnation and eventually reversal of axial flow. Finally, vortex burst was found to be correlated with swirl number, where bursting occurs at a swirl threshold of Sw<0.6 .

  14. 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.

  15. 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.

  16. 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.

  17. Compressibility and Leading-Edge Bluntness Effects for a 65 Deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2004-01-01

    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 compressibility and bluntness effects primarily at a Reynolds number of 6 million from this data set. Emphasis is placed upon on the onset and progression of leading-edge vortex separation, and compressibility is shown to promote this separation. Comparisons with recent publications show that compressibility and Reynolds number have opposite effects on blunt leading edge vortex separation

  18. 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.

  19. Evolution and Control of the Leading Edge Vortex on an Unsteady Wing

    NASA Astrophysics Data System (ADS)

    Akkala, James; Buchholz, James

    2015-11-01

    The development of the leading-edge vortex is investigated on a periodically plunging plate within a uniform free stream. Vortex circulation is governed primarily by the strength of the leading edge shear layer, which provides the primary source of circulation, and a substantial opposite-sign contribution due to the pressure-gradient-driven diffusive flux of vorticity from the suction surface of the plate. The latter has been shown to produce a substantial reduction in leading-edge vortex strength, and leads to the development of a secondary vortex whose evolution influences the interaction between the leading edge vortex and the surface, and thus alters the surface pressure gradients. Suction is applied in the vicinity of the secondary vortex in an attempt to regulate the aerodynamic loads in the presence of the leading-edge vortex. The effect on vorticity transport, leading-edge vortex dynamics, and the resulting aerodynamic loads is discussed. This work was supported, in part, by the Air Force Office of Scientific Research under Grant number FA9550-11-1-0019 and the National Science Foundation under EPSCoR grant EPS1101284.

  20. 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.

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

    PubMed

    Laplante, Caroline; Nilson, Laura A

    2011-01-24

    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

  2. FORTRAN program for calculating leading and trailing-edge geometry of turbomachine blades

    NASA Technical Reports Server (NTRS)

    Schumann, L. F.

    1977-01-01

    A FORTRAN IV program which calculates leading- and trailing-edge circle radii, tangency angles on the leading- and trailing-edge circles, and stagger angles of turbomachinery blade sections using only spline points defining the blade surfaces is described. The program shifts the origin of the blade coordinates to the leading edge of the blade. Required input includes (m, theta) coordinates of a sufficient number of spline points to adequately define the two surfaces of the blade. Other required input are the radii from the axis of rotation of the leading- and trailing-edges. The output from this program is used directly as the geometrical input for a NASA developed program for calculating transonic velocities on a blade-to-blade stream surface of a turbomachine. The program is used for axial, radial, and mixed flow turbomachine blades.

  3. 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.

  4. 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.

  5. 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.

  6. Effect of leading edge roundness on a delta wing in wing-rock motion

    NASA Technical Reports Server (NTRS)

    Ng, T. Terry; Malcolm, Gerald N.

    1990-01-01

    The effect of wing leading-edge roundness on wing rock was investigated using flow visualization in a water tunnel. Eighty degree delta wing models were tested on free-to-roll and forced oscillation rigs. The onset of wing rock was delayed by increasing the roundness of the leading edges. The wing rock amplitude and frequency results suggested that damping was increased at lower angles of attack but reduced at higher angles of attack. Vortex lift-off and vortex breakdown, especially during dynamic situations, were strongly affected by the leading edge roundness. Different forms of wing rock motion could be sustained by combinations of vortex breakdown and vortex lift-off. Behaviors of the wing and vortex motions were explained by the influence of leading edge roundness on the separation location, vortex trajectory, and vortex breakdown.

  7. 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.

  8. 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.

  9. 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.

  10. Vortex structures for flow over a delta wing with sinusoidal leading edge

    NASA Astrophysics Data System (ADS)

    Chen, Huang; Wang, Jin-Jun

    2014-06-01

    The idea of using sinusoidal leading edge as a kind of passive flow control method was inspired by observing the flipper movement of the humpback whale. It was believed that the protuberances along the whale's pectoral fin could delay stall, thus would enhance the maneuverability of the whale. It has also been shown that when equipped with sinusoidal leading edges, the stall of a delta wing could be delayed. In this paper, stereoscopic particle image velocimetry was adopted to study the vortex structures for the flow over a 52° swept delta wing with sinusoidal leading edges. A direct comparison with the flow over a baseline delta wing was made to illustrate the different vortex structures of these two kinds of models. Results have shown that the flow over the baseline delta wing was dominated by dual leading-edge vortices (LEVs), a structure that only existed for flow over nonslender delta wing at certain Reynolds number. On the other hand, the flow over the one with sinusoidal leading edge showed a very different pattern. It has been found in this paper that there were several pairs of LEVs existed on the leeward side of the wing, which might explain the stall-delaying effect of the delta wing with sinusoidal leading edges.

  11. A mechanism of leading-edge protrusion in the absence of Arp2/3 complex

    PubMed Central

    Suraneni, Praveen; Fogelson, Ben; Rubinstein, Boris; Noguera, Philippe; Volkmann, Niels; Hanein, Dorit; Mogilner, Alex; Li, Rong

    2015-01-01

    Cells employ protrusive leading edges to navigate and promote their migration in diverse physiological environments. Classical models of leading-edge protrusion rely on a treadmilling dendritic actin network that undergoes continuous assembly nucleated by the Arp2/3 complex, forming ruffling lamellipodia. Recent work demonstrated, however, that, in the absence of the Arp2/3 complex, fibroblast cells adopt a leading edge with filopodia-like protrusions (FLPs) and maintain an ability to move, albeit with altered responses to different environmental signals. We show that formin-family actin nucleators are required for the extension of FLPs but are insufficient to produce a continuous leading edge in fibroblasts lacking Arp2/3 complex. Myosin II is concentrated in arc-like regions of the leading edge in between FLPs, and its activity is required for coordinated advancement of these regions with formin-generated FLPs. We propose that actomyosin contraction acting against membrane tension advances the web of arcs between FLPs. Predictions of this model are verified experimentally. The dependence of myosin II in leading-edge advancement helps explain the previously reported defect in directional movement in the Arpc3-null fibroblasts. We provide further evidence that this defect is cell autonomous during chemotaxis. PMID:25568333

  12. Analytical model and stability analysis of the leading edge spar of a passively morphing ornithopter wing.

    PubMed

    Wissa, Aimy; Calogero, Joseph; Wereley, Norman; Hubbard, James E; Frecker, Mary

    2015-12-01

    This paper presents the stability analysis of the leading edge spar of a flapping wing unmanned air vehicle with a compliant spine inserted in it. The compliant spine is a mechanism that was designed to be flexible during the upstroke and stiff during the downstroke. Inserting a variable stiffness mechanism into the leading edge spar affects its structural stability. The model for the spar-spine system was formulated in terms of the well-known Mathieu's equation, in which the compliant spine was modeled as a torsional spring with a sinusoidal stiffness function. Experimental data was used to validate the model and results show agreement within 11%. The structural stability of the leading edge spar-spine system was determined analytically and graphically using a phase plane plot and Strutt diagrams. Lastly, a torsional viscous damper was added to the leading edge spar-spine model to investigate the effect of damping on stability. Results show that for the un-damped case, the leading edge spar-spine response was stable and bounded; however, there were areas of instability that appear for a range of spine upstroke and downstroke stiffnesses. Results also show that there exist a damping ratio between 0.2 and 0.5, for which the leading edge spar-spine system was stable for all values of spine upstroke and downstroke stiffnesses. PMID:26502210

  13. An experimental investigation of leading-edge vortex augmentation by blowing

    NASA Technical Reports Server (NTRS)

    Bradley, R. G.; Wray, W. O.; Smith, C. W.

    1974-01-01

    A wind tunnel test was conducted to determine the effects of over-the-wing blowing as a means of augmenting the leading-edge vortex flow of several pointed-tip, sharp-edged planforms. Arrow, delta, and diamond wings with leading-edge sweeps of 30 and 45 degrees were mounted on a body-of-revolution fuselage and tested in a low-speed wind tunnel at a Mach number of 0.2. Nozzle location data, pitch data, and flow-visualization pictures were obtained for a range of blowing rates. Results show pronounced increases in vortex lift due to the blowing.

  14. 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.

  15. Overview of Aerothermodynamic Loads Definition Study

    NASA Technical Reports Server (NTRS)

    Povinelli, L. A.

    1985-01-01

    The Aerothermodynamic Loads Definition were studied to develop methods to more accurately predict the operating environment in the space shuttle main engine (SSME) components. Development of steady and time-dependent, three-dimensional viscous computer codes and experimental verification and engine diagnostic testing are considered. The steady, nonsteady, and transient operating loads are defined to accurately predict powerhead life. Improvements in the structural durability of the SSME turbine drive systems depends on the knowledge of the aerothermodynamic behavior of the flow through the preburner, turbine, turnaround duct, gas manifold, and injector post regions.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Extension of leading-edge-suction analogy to wings with separated flow around the side edges at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Lamar, J. E.

    1974-01-01

    A method for determining the lift, drag, and pitching moment for wings which have separated flow at the leading and side edges with subsequently reattached flow downstream and inboard is presented. Limiting values of the contribution to lift of the side-edge reattached flow are determined for rectangular wings. The general behavior of this contribution is computed for rectangular, cropped-delta, cropped-diamond, and cropped-arrow wings. Comparisons of the results of the method and experiment indicate reasonably good correlation of the lift, drag, and pitching moment for a wide planform range. The agreement of the method with experiment was as good as, or better than, that obtained by other methods. The procedure is computerized and is available from COSMIC as NASA Langley computer program A0313.

  1. Aerothermodynamics at NASA-Langley Research Center

    NASA Technical Reports Server (NTRS)

    Weilmuenster, K. James

    2001-01-01

    The Aerothermodynamics Branch at NASA - Langley Research Center is tasked with developing, assessing and applying aerothermodynamic technologies to enable the development of hypersonic aircraft, launch vehicles, and planetary/earth entry systems. To accomplish this mission, the Branch capitalizes on the synergism between the experimental and computational facilities/tools which reside in the branch and a staff that can draw on five decades of experience in aerothermodynamics. The Aerothermodynamics Branch is staffed by 30 scientists/engineers. The staff, of which two-thirds are less than 40 years old, is split evenly between experimentalists and computationalists. Approximately 90 percent of the staff work on space transportation systems while the remainder work on planetary missions. The Branch manages 5 hypersonic wind tunnels which are staffed by 14 technicians, numerous high end work stations and a SGI Origin 2000 system. The Branch also utilizes other test facilities located at Langley as well as other national and international test sites. Large scale computational requirements are met by access to Agency resources.

  2. Preliminary aerothermodynamic design method for hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Harloff, G. J.; Petrie, S. L.

    1987-01-01

    Preliminary design methods are presented for vehicle aerothermodynamics. Predictions are made for Shuttle orbiter, a Mach 6 transport vehicle and a high-speed missile configuration. Rapid and accurate methods are discussed for obtaining aerodynamic coefficients and heat transfer rates for laminar and turbulent flows for vehicles at high angles of attack and hypersonic Mach numbers.

  3. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Weilmuenster, K. James; Hamilton, H. Harris, II; Olynick, David R.; Venkatapathy, Ethiraj

    2005-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Path finder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  4. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Weilmuenster, K. James; Hamilton, H. Harris, II; Olynick, David R.; Venkatapathy, Ethiraj

    1997-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  5. Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Olynick, David R.; Venkatapathy, Ethiraj

    2004-01-01

    A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

  6. 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.

  7. 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.

  8. The Performance of Finite-span Hydrofoils with Humpback Whale-like Leading Edge Protuberances

    NASA Astrophysics Data System (ADS)

    Custodio, Derrick; Henoch, Charles; Johari, Hamid

    2010-11-01

    The effects of leading edge protuberances on the lift and drag performance of finite-span hydrofoils were examined in a series of water tunnel experiments. The leading edge protuberances are analogous to the tubercles on humpback whale pectoral flippers. The hydrofoils have a rectangular planform and an aspect ratio of 4. The hydrofoil section profile is based on NACA 63(4)-021, and the leading edge has a sinusoidal geometry with constant amplitude and wavelength. The hydrofoil angle of attack was varied up to 30 , and the freestream velocity ranged from 1.8 to 5.4 m/s. Results indicate that the hydrofoils with leading edge protuberances do not stall in the traditional manner. Below 12 lift increased linearly with angle of attack. Beyond this angle, the lift either attained a nearly constant value or increased slowly up to 30 depending on the Reynolds number. Drag increased continuously with the angle of attack, and was not dependent on the Reynolds number. These observations are consistent with our previous infinite span hydrofoil data, and may be explained in terms of the flow modifications created by the leading edge protuberances.

  9. 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.

  10. Transonic flight test of a laminar flow leading edge with surface excrescences

    NASA Technical Reports Server (NTRS)

    Zuniga, Fanny A.; Drake, Aaron; Kennelly, Robert A., Jr.; Koga, Dennis J.; Westphal, Russell V.

    1994-01-01

    A flight experiment, conducted at NASA Dryden Flight Research Center, investigated the effects of surface excrescences, specifically gaps and steps, on boundary-layer transition in the vicinity of a leading edge at transonic flight conditions. A natural laminar flow leading-edge model was designed for this experiment with a spanwise slot manufactured into the leading-edge model to simulate gaps and steps like those present at skin joints of small transonic aircraft wings. The leading-edge model was flown with the flight test fixture, a low-aspect ratio fin mounted beneath an F-104G aircraft. Test points were obtained over a unit Reynolds number range of 1.5 to 2.5 million/ft and a Mach number range of 0.5 to 0.8. Results for a smooth surface showed that laminar flow extended to approximately 12 in. behind the leading edge at Mach number 0.7 over a unit Reynolds number range of 1.5 to 2.0 million/ft. The maximum size of the gap-and-step configuration over which laminar flow was maintained consisted of two 0.06-in. gaps with a 0.02-in. step at a unit Reynolds number of 1.5 million/ft.

  11. The three-dimensional leading-edge vortex of a 'hovering' model hawkmoth

    PubMed Central

    Berg, C. van den; Ellington, C.P.

    1997-01-01

    Recent flow visualisation experiments with the hawkmoth, Manduca sexta, revealed small but clear leading-edge vortex and a pronounced three-dimensional flow. Details of this flow pattern were studied with a scaled-up, robotic insect ('the flapper') that accurately mimicked the wing movements of a hovering hawkmoth. Smoke released from the leading edge of the flapper wing confirmed the existence of a small, strong and stable leading-edge vortex, increasing in size from wingbase to wingtip. Between 25 and 75 per cent of the wing length, its diameter increased approximately from 10 to 50 per cent of the wing chord. The leading-edge vortex had a strong axial flow veolocity, which stabilized it and reduced its diamater. The vortex separated from the wing at approximately 75 per cent of the wing length and thus fed vorticity into a large, tangled tip vortex. If the circulation of the leading-edge vortex were fully used for lift generation, it could support up to two-thirds of the hawkmoth's weight during the downstroke. The growth of this circulation with time and spanwise position clearly identify dynamic stall as the unsteady aerodynamic mechanism responsible for high lift production by hovering hawkmoths and possibly also by many other insect species.

  12. Active concentration of vorticity along the leading edge of a semi-circular wing

    NASA Astrophysics Data System (ADS)

    Williams, David; Collins, Jesse; Colonius, Tim

    2007-11-01

    Leading-edge vorticity concentration plays a key role in lift enhancement for insect flight, swept wings on aircraft, and in unsteady flows through the formation of the dynamic stall vortex. Using 16 spatially localized pulsed-blowing actuators, we are able to concentrate the vorticity at the leading edge of a wing with a semi-circular planform. The experiments are done in a wind tunnel with a model chord Reynolds number of 68,000. Peak vorticity values double those of the unforced case result in an 80 percent increase in lift on the wing relative to the unforced post-stall lift. The semi-circular wing obtains lift coefficients approximately 35 percent larger than a rectangular planform wing with a comparable aspect ratio. The sweep of the wing's leading edge is believed to establish a spanwise transport of vorticity, contributing to the stabilization of the leading edge vortex. Closed-loop control of the wing plunging motion in an unsteady flow stream is demonstrated by modulating the strength of the leading-edge vorticity via a proportional-derivative controller.

  13. 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

  14. 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.

  15. Influence of blade angle distribution along leading edge on cavitation performance of a centrifugal pump

    NASA Astrophysics Data System (ADS)

    Xu, Y.; Tan, L.; Cao, S. L.; Wang, Y. C.; Meng, G.; Qu, W. S.

    2015-01-01

    The influence of blade angle distribution along leading edge on cavitation performance of centrifugal pumps is analysed in the present paper. Three sets of blade angle distribution along leading edge for three blade inlet angles are chosen to design nine centrifugal pump impellers. The RNG k-epsilon turbulence model and the Zwart-Gerber-Belamri cavitation model are employed to simulate the cavitation flows in centrifugal pumps with different impellers and the same volute. The numerical results are compared with the experimental data, and the comparison proves that the numerical simulation can accurately predict the cavitation performance of centrifugal pumps. On the basis of the numerical simulations, the pump head variations with pump inlet pressure, and the flow details in centrifugal pump are revealed to demonstrate the influence of blade angle distribution along leading edge on cavitation performances of centrifugal pumps.

  16. 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.

  17. 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.

  18. Characterizing a burst leading-edge vortex on a rotating flat plate wing

    NASA Astrophysics Data System (ADS)

    Jones, Anya R.; Medina, Albert; Spooner, Hannah; Mulleners, Karen

    2016-04-01

    Identifying, characterizing, and tracking incoherent vortices in highly separated flows is of interest for the development of new low-order models for unsteady lift prediction. The current work examines several methods to identify vortex burst and characterize a burst leading-edge vortex. Time-resolved stereoscopic PIV was performed on a rotating flat plate wing at Re = 2500. The burst process was found to occur at mid-span and is characterized by axial flow reversal, the entrainment of opposite-sign vorticity, and a rapid expansion of vortex size. A POD analysis revealed that variations in certain mode coefficients are indicative of the flow state changes characteristics of burst. During burst, the leading-edge vortex evolves to a region of inhomogeneous vorticity distributed over a large area. Several methods of defining the vortex size and circulation are evaluated and a combination of these can be used to characterize the leading-edge vortex both pre- and post-burst.

  19. 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.

  20. 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.

  1. 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.

  2. Effect of leading-edge porosity on blade-vortex interaction noise

    NASA Technical Reports Server (NTRS)

    Lee, Soogab

    1993-01-01

    The effect of the porous leading-edge of an airfoil on the blade-vortex interaction noise, which dominates far-field acoustic spectrum of the helicopter, is investigated. The thin-layer Navier-Stokes equations are solved with a high-order upwind-biased scheme and a multizonal grid system. The Baldwin-Lomax turbulence model is modified for considering transpiration on the surface. The amplitudes of the propagating acoustic wave in the near-field are calculated directly from the computation. The porosity effect on the surface is modeled. Results show leading-edge transpiration can suppress pressure fluctuations at the leading-edge during BVI, and consequently reduce the amplitude of propagating noise by 30 percent at maximum in the near-field. The effect of porosity factor on the noise level is also investigated.

  3. Leading edge sweep effects in generic three-dimensional sidewall compression scramjet inlets

    NASA Technical Reports Server (NTRS)

    Cozart, Aaron B.; Holland, Scott D.; Trexler, Carl A.; Perkins, John N.

    1992-01-01

    A computational and experimental study of generic 3D sidewall compression inlets is conducted to examine the effects of fore and aft leading edge sweep on the internal shock structure. Inlets with leading edge sweeps of +30 deg and -30 deg with sidewall compression angles of 6 deg were tested in the NASA Langley Mach 4 air tunnel at a geometric contraction ratio of 1.87. The principal difference in performance was determined to be in the mass capture. Spillage was identified as having two components: a pressure induced component and a sweep induced component. It was found that while the direction of the leading edge sweep had a large influence on the spillage, the pressure effects were more important.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Impingement cooling with film coolant extraction in the airfoil leading edge regions

    NASA Astrophysics Data System (ADS)

    Li, Liguo; Li, Zhaohui

    An extensive experimental study is conducted to determine the heat transfer characteristics of arrays of air jets impinging on perforated target surfaces in turbine blade leading edge regions by six large-scale models. The relations of pressure loss and Nusselt number to jet Reynolds number are obtained in a wide range of parameter combinations of interest in cooled airfoil practice for various models, respectively. These parameter combinations are covered in a test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge curvature radius-to-diameter ratio, jet pitch-to-diameter ratio, and jet impingement gap-to-diameter ratio.

  13. Leading-edge tubercles delay stall on humpback whale (Megaptera novaeangliae) flippers

    NASA Astrophysics Data System (ADS)

    Miklosovic, D. S.; Murray, M. M.; Howle, L. E.; Fish, F. E.

    2004-05-01

    The humpback whale (Megaptera novaeangliae) is exceptional among the baleen whales in its ability to undertake acrobatic underwater maneuvers to catch prey. In order to execute these banking and turning maneuvers, humpback whales utilize extremely mobile flippers. The humpback whale flipper is unique because of the presence of large protuberances or tubercles located on the leading edge which gives this surface a scalloped appearance. We show, through wind tunnel measurements, that the addition of leading-edge tubercles to a scale model of an idealized humpback whale flipper delays the stall angle by approximately 40%, while increasing lift and decreasing drag.

  14. 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.

  15. Image analysis tools to quantify cell shape and protein dynamics near the leading edge.

    PubMed

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

    2013-01-01

    We present a set of flexible image analysis tools to analyze dynamics of cell shape and protein concentrations near the leading edge of cells adhered to glass coverslips. Plugins for ImageJ streamline common analyses of microscopic images of cells, including the calculation of leading edge speeds, total and average intensities of fluorescent markers, and retrograde flow rate measurements of fluorescent single-molecule speckles. We also provide automated calculations of auto- and cross-correlation functions between velocity and intensity measurements. The application of the methods is illustrated on images of XTC cells. PMID:23165752

  16. 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.

  17. 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

  18. 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.

  19. On the effect of leading edge blowing on circulation control airfoil aerodynamics

    NASA Technical Reports Server (NTRS)

    Mclachlan, B. G.

    1987-01-01

    In the present context the term circulation control is used to denote a method of lift generation that utilizes tangential jet blowing over the upper surface of a rounded trailing edge airfoil to determine the location of the boundary layer separation points, thus setting an effective Kutta condition. At present little information exists on the flow structure generated by circulation control airfoils under leading edge blowing. Consequently, no theoretical methods exist to predict airfoil performance under such conditions. An experimental study of the flow field generated by a two dimensional circulation control airfoil under steady leading and trailing edge blowing was undertaken. The objective was to fundamentally understand the overall flow structure generated and its relation to airfoil performance. Flow visualization was performed to define the overall flow field structure. Measurements of the airfoil forces were also made to provide a correlation of the observed flow field structure to airfoil performance. Preliminary results are presented, specifically on the effect on the flow field structure of leading edge blowing, alone and in conjunction with trailing edge blowing.

  20. Space shuttle wing leading edge heating environment prediction derived from development flight data

    NASA Technical Reports Server (NTRS)

    Cunningham, J. A.; Haney, J. W., Jr.

    1983-01-01

    An analytical program is in progress at Rockwell International to revise wing leading edge heating predictions in order to improve correlation with STS-1 to -5 flight radiometer data. This paper discusses the methods that have been used to improve agreement between prediction and flight and summarizes the aerodynamic correlations which, when updated, will be used to analyze future orbiter missions.

  1. Spanwise gradients in flow speed help stabilize leading-edge vortices on revolving wings.

    PubMed

    Jardin, T; David, L

    2014-07-01

    While a leading-edge vortex on an infinite translating wing is shed after a short distance of travel, its counterpart on a finite span revolving insect wing or maple seed membrane exhibits robust attachment. The latter explains the aerodynamic lift generated by such biological species. Here we analyze the mechanisms responsible for leading-edge vortex attachment. We compute the Navier-Stokes solution of the flow past a finite span wing (i) embedded in a uniform oncoming flow, (ii) embedded in a spanwise varying oncoming flow, and (iii) revolving about its root. We show that over flapping amplitudes typical of insect flight (ϕ = 120°), the spanwise gradient of the local wing speed may suffice in maintaining leading-edge vortex attachment. We correlate this result with the development of spanwise flow, driven by the spanwise gradient of pressure, and we evaluate the sensitivity of such a mechanism to the Reynolds number. It is noted, however, that leading-edge vortex attachment through the spanwise gradient of the local wing speed does not promote large lift, which ultimately arises from centrifugal and Coriolis effects. PMID:25122373

  2. 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.

  3. 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 work and apply that…

  4. 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 as the…

  5. 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…

  6. Effects of Fin Leading Edge Sweep on Shock-Shock Interaction at Mach 6

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Nowak, Robert J.

    1996-01-01

    The effects of fin leading edge sweep on peak heating rates due to shock-shock interaction have been experimentally examined in the Langley 20-Inch Mach 6 Tunnel. The shock interaction was produced by the intersection of a planar incident shock (16.8 deg shock angle relative to the freestream, generated by a 9 deg wedge) with the bow shock formed around a O.5-inch diameter cylindrical leading edge fin. Heating distributions along the leading edge stagnation line have been obtained using densely spaced thin film resistive-type sensors. Schlieren images were obtained to illustrate the very complex shock-shock interactions. The fin leading edge sweep angle was varied from 15-degrees swept back to 45-degrees swept forward for a freestream unit Reynolds number of 2 x 10(exp 6)/ft. Two models were utilized during the study, one with 0.025-inch spacing between gage centers, and the other 0.015-inch spacing. Gage spatial resolution on the order of 0.015-in appeared to accurately capture the narrow spike in heating. Peak heating due to shock interaction was maximized when the fin was swept forward 15 deg and 25 deg, both promoting augmentations about 7 times the baseline value. The schlieren images for these cases revealed Type 4 and Type 3 interactions, respectively.

  7. Arc Jet Results on Candidate High Temperature Coatings for NASA's NGLT Refractory Composite Leading Edge Task

    NASA Technical Reports Server (NTRS)

    Ohlhorst, C. W.; Vaughn, W. L.; Lewis, R. K.; Milhoan, J. D.

    2004-01-01

    In 2000, arc jet testing was conducted on thirteen material systems for possible use on the nose leading edge of the Hyper-X program s X-43A Mach 10 vehicle. Six material systems survived 3, 130- second cycles. To support NASA s Next Generation Launch Technology Programs (NGLT) need for passive refractory composite leading edges with multiple reuse capability at temperatures up to 3600 F, these six materials were subjected to an expanded arc jet test program. This expanded arc jet test program included three phases. The purpose of the first phase was to generate emissivity data as a function of temperature. The purpose of the second phase was to determine if the material systems had any thermal cycling durability, and the third phase was to determine whether the materials could survive an arc jet test of one hour duration. Some of the coating systems were found to have very low emissivities, suggesting that they would not be good candidates for leading edges coating. Other coating systems survived both the second and third phases of the test program and showed potential for use as an oxidation protection coating for leading edges. This presentation summarizes the test program results.

  8. 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.

  9. A role for actin arcs in the leading edge advance of migrating cells

    PubMed Central

    Burnette, Dylan T.; Manley, Suliana; Sengupta, Prabuddha; Sougrat, Rachid; Davidson, Michael W.; Kachar, Bechara; Lippincott-Schwartz, Jennifer

    2013-01-01

    The migration of epithelial cells requires coordination of two actin modules at the leading edge: one in the lamellipodium and one in the lamella. How the two modules connect mechanistically to regulate directed edge motion is not understood. Using a combination of live-cell imaging and photoactivation approaches, we demonstrate that the actin network of the lamellipodium evolves spatio-temporally into the lamella. This occurs during the retraction phase of edge motion when myosin II redistributes to the cell edge and condenses the lamellipodial-actin into an arc-like bundle (i.e., actin arc) parallel to the edge. The newly formed actin arc moves rearward and couples to focal adhesions as it enters the lamella. We propose net edge extension occurs by nascent focal adhesions advancing the site at which new actin arcs slow down and form the base of the next protrusion event. The actin arc thus serves as a structural element underlying the temporal and spatial connection between the lamellipodium and lamella to drive directed cell motion. PMID:21423177

  10. 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.

  11. 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.

  12. 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.

  13. Estimation of ground range on the sweep frequency backscatter leading edge

    NASA Astrophysics Data System (ADS)

    Su, Fanfan; Zhao, Zhengyu; Deng, Feng; Li, Shipeng

    2011-04-01

    The high frequency management system with backscatter radar supplies the real time ionosphere channel conditions to high frequency users, which leads to the demand for the ground range between the radar location and the scatters on the distant ground. The ionosphere electron density profile is usually inversed to obtain the ground range. An inversion algorithm, with which the ground range on the leading edge of the backscatter ionograms can be obtained without electron density, is presented in this paper. The ray path geometry of the backscatter sounding and the change in the group path on the leading edge with operating frequency are used to derive the ground range. Synthesized backscatter ionogram and experimental backscatter ionograms are processed to validate the algorithm. The results indicate that the algorithm is usable for high frequency management system.

  14. 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

  15. Effects of leading and trailing edge flaps on the aerodynamics of airfoil/vortex interactions

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed A.; Sankar, L. N.; Tadghighi, H.

    1994-01-01

    A numerical procedure has been developed for predicting the two-dimensional parallel interaction between a free convecting vortex and a NACA 0012 airfoil having leading and trailing edge integral-type flaps. Special emphasis is placed on the unsteady flap motion effects which result in alleviating the interaction at subcritical and supercritical onset flows. The numerical procedure described here is based on the implicit finite-difference solutions to the unsteady two-dimensional full potential equation. Vortex-induced effects are computed using the Biot-Savart Law with allowance for a finite core radius. The vortex-induced velocities at the surface of the airfoil are incorporated into the potential flow model via the use of the velocity transpiration approach. Flap motion effects are also modeled using the transpiration approach. For subcritical interactions, our results indicate that trailing edge flaps can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, our results demonstrate the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time-dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented

  16. Single velocity-component modeling of leading edge turbulence interaction noise.

    PubMed

    Gill, J; Zhang, X; Joseph, P

    2015-06-01

    A computational aeroacoustics approach is used to predict leading edge turbulence interaction noise for real airfoils. One-component (transverse), two-component (transverse and streamwise), and three-component (transverse, streamwise, and spanwise) synthesized turbulence disturbances are modeled instead of harmonic transverse gusts, to which previous computational studies of leading edge noise have often been confined. The effects of the inclusion of streamwise and spanwise disturbances on the noise are assessed. It is shown that accurate noise predictions can be made by modeling only transverse disturbances which reduces the computational expense of simulations. The accuracy of using only transverse disturbances is assessed for symmetric and cambered airfoils, and also for airfoils at non-zero angle of attack. PMID:26093411

  17. Integrated production overlay field-by-field control for leading edge technology nodes

    NASA Astrophysics Data System (ADS)

    Chung, Woong Jae; Tristan, John; Gutjahr, Karsten; Subramany, Lokesh; Li, Chen; Sun, Yulei; Yelverton, Mark; Kim, Young Ki; Kim, Jeong Soo; Huang, Chin-Chou Kevin; Pierson, William; Karur-Shanmugam, Ramkumar; Riggs, Brent; Jug, Sven; Robinson, John C.; Yap, Lipkong; Ramanathan, Vidya

    2014-04-01

    As photolithography will continue with 193nm immersion multiple patterning technologies for the leading edge HVM process node, the production overlay requirement for critical layers in logic devices has almost reached the scanner hardware performance limit. To meet the extreme overlay requirements in HVM production environment, this study investigates a new integrated overlay control concept for leading edge technology nodes that combines the run-to-run (R2R) linear or high order control loop, the periodic field-by-field or correction per exposure (CPE) wafer process signature control loop, and the scanner baseline control loop into a single integrated overlay control path through the fab host APC system. The goal is to meet the fab requirements for overlay performance, lower the cost of ownership, and provide freedom of control methodology. In this paper, a detailed implementation of this concept will be discussed, along with some preliminary results.

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

    SciTech Connect

    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 range 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 roughness

  19. Unsteady vortex-dominated flow around wings with oscillating leading-edge flaps

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1991-01-01

    Unsteady vortex-dominated flow around delta wings with oscillating leading-edge flaps represents an important classs of problems for supermaneuverability and flow control of advanced aircraft. The problem is solved using time accurate integration of the unsteady, compressible, thin-layer Navier-Stokes equations in conjunction with the unsteady, linearized, Navier-displacement equations. Starting with an initial configuration of the wing and its flaps, the Navier-Stokes equations are solved on an initial structured grid for the steady flow. The forced oscillation of the flaps is then applied, and the problem is solved accurately in time. The Navier-displacement equations are solved for the grid deformation and the Navier-Stokes equations are solved for the flowfield. Symmetric and anti-symmetric flaps oscillations are presented to study the effect of the flaps oscillation on the leading-edge vortical flow.

  20. Effects of leading-edge flap oscillation on unsteady delta wing flow and rock control

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Salman, Ahmed A.

    1991-01-01

    The isolated and interdisciplinary problems of unsteady fluid dynamics and rigid-body dynamics and control of delta wings with and without leading-edge flap oscillation are considered. For the fluid dynamics problem, the unsteady, compressible, thin-layer Navier-Stokes (NS) equations, which are written relative to a moving frame of reference, are solved along with the unsteady, linearized, Navier-displacement (ND) equations. The NS equations are solved for the flowfield using an implicit finite-volume scheme. The ND equations are solved for the grid deformation, if the leading-edge flaps oscillate, using an ADI scheme. For the dynamics and control problem, the Euler equation of rigid-body rolling motion for a wing and its flaps are solved interactively with the fluid dynamics equations for the wing-rock motion and subsequently for its control. A four-stage Runge-Kutta scheme is used to explicitly integrate the dynamics equation.

  1. 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.

  2. 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.

  3. Heat pipes for wing leading edges of hypersonic vehicles. Final report

    SciTech Connect

    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.

  4. 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.

  5. Effect of a round airfoil nose on leading-edge suction

    NASA Technical Reports Server (NTRS)

    Lan, C. Edward; Su, Ingchung

    1987-01-01

    Kulfan (1979) assumed that the angle of attack for initial vortex separation on a slender wing with rounded leading edges could be obtained by equating the leading-edge suction (LES) and nose drag coefficients. In the present study, this assumption is examined and is shown to predict reasonably well the initial angle of attack at which laminar separation occurs near the airfoil nose. However, the assumption is shown to be slightly less accurate for thick or cambered airfoils. Attainable LES estimated by Kulfan's method seemed to agree well with that obtained from an airfoil aerodynamics code and experimental data on a NACA 64A009 airfoil at M = 0.4 and Re = 0.86 x 10 to the 6th.

  6. Comparison of computations and experimental data for leading edge vortices - Effects of yaw and vortex flaps

    NASA Technical Reports Server (NTRS)

    Murman, E. M.; Powell, K. G.; Miller, D. S.; Wood, R. M.

    1986-01-01

    Computations are presented using the conical Euler equations for swept delta wings with leading edge vortices. All the wings have sharp leading edges swept at 75 degrees to the freestream. In addition to an idealized flat plate model, geometrical features also included are thickness, centerbody, and two vortex flaps. Freestream Mach numbers of 1.7 to 2.8, angles of attack of 10 and 12 degrees, and angles of yaw of 0 and 8 degrees are considered. The computations are compared with pitot pressure traverses for one case. Other calculations are compared with pitot pressure traverses for one case. Other calculations are compared with surface pressure data and vapor screen pictures recently obtained at NASA Langley Research Center. The comparisons indicate that the dominant features of these flows are adequately modeled by the Euler equations, but viscous models are needed for the surface boundary layer and secondary separations.

  7. Comparison of computations and experimental data for leading edge vortices - Effects of yaw and vortex flaps

    NASA Technical Reports Server (NTRS)

    Murman, E. M.; Powell, K. G.; Miller, D. S.; Wood, R. M.

    1986-01-01

    Computations are presented using the conical Euler equations for swept delta wings with leading edge vortices. All the wings have sharp leading edges swept at 75 degrees to the freestream. In addition to an idealized flat plate model, geometrical features also included are thickness, centerbody, and two vortex flaps. Freestream Mach numbers of 1.7 to 2.8, angles of attack of 10 and 12 degrees, and angles of yaw of 0 and 8 degrees are considered. The computations are compared with pitot pressure traverses for one case. Other calculations are compared with surface pressure data and vapor screen pictures recently obtained at NASA Langley Research Center. The comparisons indicate that the dominant features of these flows are adequately modeled by the Euler equations, but viscous models are needed for the surface boundary layer and secondary separations.

  8. Leading edge vortex-flap experiments on a 74 deg delta wing

    NASA Technical Reports Server (NTRS)

    Rao, D. M.

    1979-01-01

    Exploratory wind tunnel tests are reported on a 74 deg. delta wing model. The potential of a vortex flap concept in reducing the subsonic lift dependent drag of highly swept, slender wings is examined. The suction effect of coiled vortices generated through controlled separation over leading edge flap surfaces to produce a thrust component is discussed. A series of vortex-flap configurations were investigated to explore the effect of some primary geometric variables.

  9. Analysis of leading edge separation using a low order panel method

    NASA Technical Reports Server (NTRS)

    Sandlin, Doral R.

    1989-01-01

    An examination of the potential flow computer code VSAERO to model leading edge separation over a delta wing is examined. Recent improvements to the code suggest that it may be capable of predicting pressure coefficients on the body. Investigation showed that although that code does predict the vortex roll-up, the pressure coefficients have significant error. The program is currently unsatisfactory, but with some additional development it may become a useful tool for this application.

  10. The lift of sharp-leading-edged delta wings with blowing

    NASA Technical Reports Server (NTRS)

    Tavella, D. A.

    1985-01-01

    An analysis of the lift augmentation due to a thin jet of air issuing from a slot along the leading edge of a delta wing is presented. The problem is treated with an extension of the method of Brown and Michael, representing the separated flow on the lee side of the wing by a pair of concentrated vortices and corresponding feeding sheets. It is assumed that the jet is not affected by Coanda forces. The analysis produces qualitative agreement with experiments.

  11. A theory for the core of a three-dimensional leading-edge vortex

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    1985-01-01

    A theory is presented for the flow in the core of a separation-induced leading-edge vortex. The theory is based on matching inner and outer representations of the vortex. The inner representation models continuously distributed vorticity and includes an asymptotic viscous subcore. The outer representation models concentrated spiral sheets of vorticity and is fully three dimensional. A parameter is identified which closely tracks the vortex breakdown stability boundary for delta, arrow, and diamond wings.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Experimental investigation of the leading edge vortex on vertical axis wind turbine blades

    NASA Astrophysics Data System (ADS)

    Dunne, Reeve; McKeon, Beverley

    2012-11-01

    A NACA 0018 airfoil is pitched about the leading edge over a large angle of attack range (+/- ~40°) at a chord Reynolds number of 110,000 to simulate the flow over a single blade in a vertical axis wind turbine (VAWT). Particle image velocimetry (PIV) measurements are made to investigate the effects of pitching on leading edge vortex (LEV) development and separation. Time resolved experiments are performed to track vortex formation and convection over the airfoil for sinusoidal pitching motions corresponding to a VAWT trajectory as well as impulsive pitch up and pitch down motions. These results are compared to the wake of steady, post stall, high angle of attack airfoils (α =20° -30°). The characteristics of the leading edge vortex development and subsequent separation from the airfoil are discussed, with a view to characterizing its effect on power generation with VAWTs and future flow control strategies for turbine performance improvement. Funding from the Gordon and Betty Moore Foundation is gratefully acknowledged.

  17. 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.

  18. Actin filament turnover drives leading edge growth during myelin sheath formation in the central nervous system

    PubMed Central

    Schmitt, Sebastian; Snaidero, Nicolas; Mitkovski, Mišo; Velte, Caroline; Brückner, Bastian R.; Alexopoulos, Ioannis; Czopka, Tim; Jung, Sang Y.; Rhee, Jeong S.; Janshoff, Andreas; Witke, Walter; Schaap, Iwan A.T.; Lyons, David A.; Simons, Mikael

    2016-01-01

    Summary During central nervous system development, oligodendrocytes wrap their plasma membrane around axons to generate multi-lamellar myelin sheaths. To drive growth at the leading edge of myelin at the interface with the axon, mechanical forces are necessary, but the underlying mechanisms are not known. Using an interdisciplinary approach that combines morphological, genetic and biophysical analyses, we identified a key role for actin filament network turnover in myelin growth. At the onset of myelin biogenesis, F-actin is redistributed to the leading edge, where its polymerization-based forces push out non-adhesive and motile protrusions. F-actin disassembly converts protrusions into sheets by reducing surface tension and in turn inducing membrane spreading and adhesion. We identified the actin depolymerizing factor ADF/Cofilin1, which mediates high F-actin turnover rates, as essential factor in this process. We propose that F-actin turnover is the driving force in myelin wrapping by regulating repetitive cycles of leading edge protrusion and spreading. PMID:26166299

  19. 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.

  20. A theory for the core flow of leading-edge vortices

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    1986-01-01

    Separation-induced leading-edge vortices can dominate the flow about slender wings at moderate to high angles of attack, often with favorable aerodynamic effects. However, at the high angles of attack which are desirable for takeoff and landing as well as subsonic-transonic maneuver the vortices can breakdown or burst in the vicinity of the aircraft causing many adverse effects; these include lift loss, pitchup, and buffet. The flow in the core of leading-edge vortices is generally affiliated with the vortex breakdown phenomenon. A theory is presented for the flow in the core of separation-induced, leading-edge vortices at practical Reynolds numbers. The theory is based on matching inner and outer representations of the vortex. The inner representation models continuously distributed vorticity and includes an asymptotic viscous subcore. The outer representation models concentrated spiral sheets of vorticity and is fully three dimensional. A parameter is identified which closely tracks the vortex breakdown stability boundary for delta, arrow, and diamond wings.

  1. 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.

  2. Numerical lifting line theory applied to drooped leading-edge wings below and above stall

    NASA Technical Reports Server (NTRS)

    Anderson, J. D., Jr.; Corda, S.; Van Wie, D. M.

    1980-01-01

    A numerical iterative solution to the classical Prandtl lifting-line theory, suitably modified for poststall behavior, is used to study the aerodynamic characteristics of straight rectangular finite wings with and without leading-edge droop. This study is prompted by the use of such leading-edge modifications to inhibit stall/spins in light general aviation aircraft. The results indicate that lifting-line solutions at high angle of attack can be obtained that agree with experimental data to within 20%, and much closer for many cases. Therefore, such solutions give reasonable preliminary engineering results for both drooped and undrooped wings in the poststall region. However, as predicted by von Karman, the lifting-line solutions are not unique when sectional negative lift slopes are encountered. In addition, the present numerical results always yield symmetrical lift distributions along the span, in contrast to the asymmetrical solutions observed by Schairer in the late 1930's. Finally, a series of parametric tests at low angle of attack indicate that the effect of drooped leading edges on aircraft cruise performance is minimal.

  3. Recycling endosome membrane incorporation into the leading edge regulates lamellipodia formation and macrophage migration.

    PubMed

    Veale, Kelly J; Offenhäuser, Carolin; Whittaker, Shane P; Estrella, Ruby P; Murray, Rachael Z

    2010-10-01

    In comparison to our knowledge of the recycling of adhesion receptors and actin assembly, exactly how the cell controls its surface membrane to form a lamellipodium during migration is poorly understood. Here, we show the recycling endosome membrane is incorporated into the leading edge of a migrating cell to expand lamellipodia membrane. We have identified the SNARE complex that is necessary for fusion of the recycling endosome with the cell surface, as consisting of the R-SNARE VAMP3 on the recycling endosome partnering with the surface Q-SNARE Stx4/SNAP23, which was found to translocate and accumulate on the leading edge of migrating cells. Increasing VAMP3-mediated fusion of the recycling endosome with the surface increased membrane ruffling, while inhibition of VAMP3-mediated fusion showed that incorporation of the recycling endosome is necessary for efficient lamellipodia formation. At the same time, insertion of this recycling endosome membrane also delivers its cargo integrin α5β1 to the cell surface. The loss of this extra membrane for lamellipodia expansion and delivery of cargo in cells resulted in macrophages with a diminished capacity to effectively migrate. Thus, the recycling endosome membrane is incorporated into the leading edge and this aids expansion of the lamellipodia and simultaneously delivers integrins necessary for efficient cell migration. PMID:20604897

  4. 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.

  5. 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).

  6. 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 schools across the…

  7. 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 schools across the…

  8. 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 schools across the…

  9. 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 schools across the…

  10. 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 schools across the…

  11. 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 schools across the…

  12. 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 schools across the…

  13. 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.

  14. Exploration in optimal design of an airfoil with a leading edge rotating cylinder

    NASA Astrophysics Data System (ADS)

    Zhang, Yuan-Yuan; Huang, Dian-Gui; Sun, Xiao-Jing; Wu, Guo-Qing

    2010-08-01

    Based on the theory of moving surface boundary layer control (MSBC), a concept of an airfoil having a rotating cylinder at the leading edge has been developed and experimentally proven to have good aerodynamic performance even at large angles of attack. Thus, this research aims to give guidance on optimizing the design of this kind of airfoil with high lift coefficients. Using computational fluid dynamics (CFD) technique, the CFD simulation results have been compared with the experimental results available in the literature, and then the SST two-equation model is selected as the appropriate turbulence model. At a given cylinder surface velocity ratio, the cylinder diameter d, the drop height of trailing edge δ and the curvatures of the pressure and suction surfaces of the airfoil are regarded as the optimal design parameters and the airfoil lift coefficient is considered as the optimization objective function. Therefore, using orthogonal optimization method, we herein develop a new design of airfoil favorable for having a rotating leading edge. It has been numerically proven that the resulting airfoil has good capability of achieving a substantially superior performance when compared to the airfoils of the prior art.

  15. Analysis of the Flow About Delta Wings with Leading Edge Separation at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Nenni, J. P.; Tung, C.

    1973-01-01

    A research program was conducted to develop an improved theoretical flow model for the flow about sharp edge delta wings with leading-edge separation at supersonic speeds. The flow model incorporates a representation of the secondary separation region which occurs just inboard of the leading edge on such wings and is based on a slender-wing theory whereby the full three-dimensional problem is reduced to a quasi two-dimensional problem in the cross-flow plane. The secondary separation region was modeled by a surface distribution of singularities or a linearized type of cavity representation. The primary vortex and separation were modeled by a concentrated vortex and cut in the cross-flow potential which represents its feeding sheet. The cross-flow solutions for the cavity model were obtained, but these solutions have physical significance only in a very restricted range of angle of attack. The reasons for the failure of the flow model are discussed. The analysis is presented so that other interested researchers may critically review the work.

  16. Precocious reproduction increases at the leading edge of a mangrove range expansion.

    PubMed

    Dangremond, Emily M; Feller, Ilka C

    2016-07-01

    Climate change-driven shifts in species ranges are ongoing and expected to increase. However, life-history traits may interact with climate to influence species ranges, potentially accelerating or slowing range shifts in response to climate change. Tropical mangroves have expanded their ranges poleward in the last three decades. Here, we report on a shift at the range edge in life-history traits related to reproduction and dispersal. With a common garden experiment and field observations, we show that Rhizophora mangle individuals from northern populations reproduce at a younger age than those from southern populations. In a common garden at the northern range limit, 38% of individuals from the northernmost population were reproductive by age 2, but less than 10% of individuals from the southernmost population were reproductive by the same age, with intermediate amounts of reproduction from intermediate latitudes. Field observations show a similar pattern of younger reproductive individuals toward the northern range limit. We also demonstrate a shift toward larger propagule size in populations at the leading range edge, which may aid seedling growth. The substantial increase in precocious reproduction at the leading edge of the R. mangle range could accelerate population growth and hasten the expansion of mangroves into salt marshes. PMID:27547335

  17. Lead uptake in diverse plant families: a study applying X-ray absorption near edge spectroscopy.

    PubMed

    Bovenkamp, Gudrun L; Prange, Alexander; Schumacher, Wolfgang; Ham, Kyungmin; Smith, Aaron P; Hormes, Josef

    2013-05-01

    The chemical environment of lead in roots and leaves of plants from four different plant families and a lichen from a former lead mining site in the Eifel Mountains in Germany was determined by Pb L3-edge XANES measurements using solid reference compounds and also aqueous solutions of different ionic strength simulating the plant environment. Pb(2+) ions in the plants were found to have two major coordinations, one with nine oxygen atoms in the first coordination shell similar to outer-sphere complexation and a second coordination with just three oxygen atoms similar to inner-sphere complexation. This can be interpreted assuming that lead is sorbed on the surface of cell walls depending on the concentration of lead in the soil solution. Pb L3-edge XANES spectra of dried and fresh plant samples are very similar because sorption does not change with removal of water but only because of the initial ionic strength. No bonding to biologically important groups (-S, - N) or precipitation (-PO4) was found. PMID:23517351

  18. 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.

  19. An experimental investigation of convective heat transfer at the leading edge of a gas turbine airfoil

    NASA Astrophysics Data System (ADS)

    Gendron, S.; Marchand, N. J.; Korn, C.; Immarigeon, J. P.; Kacprzynski, J. J.

    1992-06-01

    This paper describes the experimental methods used to determine the surface temperatures and heat-transfer coefficients at the leading edge, and elsewhere over the surface, of a specially designed double-edge wedge shell specimen subjected to cyclic heating in a high velocity hot gas stream generated by a burner rig. The methods included temperature measurements with thermocouples (embedded below the surface) as well as surface temperature measurements by optical pyrometry. The experiments were carried out at gas temperatures between 806 to 1323 C and velocities in the range from Mach 0.32 to Mach 0.39. The calibration procedures for each method, the various testing conditions to which the airfoil-like specimen was exposed and the results pertaining to the determination of the surface temperatures and heat-transfer coefficients are described and discussed.

  20. Prediction of unsteady aerodynamic loadings caused by leading edge and trailing edge control surface motions in subsonic compressible flow: Computer program description

    NASA Technical Reports Server (NTRS)

    Redman, M. C.; Rowe, W. S.

    1975-01-01

    A digital computer program has been developed to calculate unsteady loadings caused by motions of lifting surfaces with leading edge or trailing edge controls based on the subsonic kernel function approach. The pressure singularities at hinge line and side edges have been extracted analytically as a preliminary step to solving the integral equation by collocation. The program calculates generalized aerodynamic forces for user supplied deflection modes. Optional intermediate output includes pressure at an array of points, and sectional generalized forces. From one to six controls on the half span can be accommodated.

  1. An improved Woodward's panel method for calculating leading-edge and side-edge suction forces at subsonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

    Woodward's panel method for subsonic and supersonic flow was improved by employing control points determined by exactly matching two-dimensional pressure at a finite number of points. The results show great improvement in the predicted pressure distribution of a flapped airfoil. With the paneling scheme of cosine law in both chordwise and spanwise directions, the method is shown to accurately predict leading edge and side edge suction forces of various configurations in subsonic and supersonic flow.

  2. 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.

  3. 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.

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

    SciTech Connect

    Modlin, J.M.

    1991-01-01

    The feasibility of cooling hypersonic-vehicle leading-edge structures exposed to severe aerodynamic surface heat fluxes was studied, 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. Then the existing heat-pipe model was modified by adding both transpiration and film-cooling options as new surface boundary conditions. The models used to predict the leading-edge surface heat-transfer reduction effects of the transpiration and film cooling were modifications of more-generalized, empirically based models obtained from the literature. It is concluded 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 appears feasible.

  5. Mechanical Properties and Microstructural Characterization of Particulate Reinforced Diboride Composites for High Temperature Leading Edge Applications

    NASA Technical Reports Server (NTRS)

    Ellerby, Donald T.; Johnson, Sylvia M.; Bull, Jeff; Laub, Bernie; Reuther, James; Kinney, David; Kontinos, Dean; Beckman, Sarah; Stuffle, Kevin; Cull, A. D.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Previous work on refractory diboride composites has shown that these systems have the potential for use in high temperature leading edge applications for reusable reentry vehicles. Experiments in reentry environments have shown that these materials have multiple use temperatures greater than 1900 C. The work to be discussed focuses on three compositions: HfB2/SiC, ZrB2/SiC, and ZrB2/C/SiC. These composites have been hot pressed and their mechanical properties measured at room and elevated temperatures. Extensive microstructural characterization has been conducted on polished cross sections and the fracture surfaces have been examined to determine their failure origins.

  6. 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.

  7. 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.

  8. Nondestructive Evaluation Tests Performed on Space Shuttle Leading- Edge Materials Subjected to Impact

    NASA Technical Reports Server (NTRS)

    Roth, Don J.; Martin, Richard E.; Bodis, James R.

    2005-01-01

    In support of the space shuttle Return To Flight efforts at the NASA Glenn Research Center, a series of nondestructive evaluation (NDE) tests were performed on reinforced carbon/carbon (RCC) composite panels subjected to ballistic foam impact. The impact tests were conducted to refine and verify analytical models of an external tank foam strike on the space shuttle leading edge. The NDE tests were conducted to quantify the size and location of the resulting damage zone as well as to identify hidden damage.

  9. Structure of unsteady flows at leading- and trailing-edges: Flow visualization and its interpretation

    NASA Technical Reports Server (NTRS)

    Rockwell, D.; Atta, R.; Kramer, L.; Lawson, R.; Lusseyran, D.; Magness, C.; Sohn, D.; Staubli, T.

    1987-01-01

    Unsteady two- and three-dimensional flow structure at leading and trailing edges of bodies can be characterized effectively using recently developed techniques for acquisition and interpretation of flow visualization. The techniques addressed here include: flow image/surface pressure correlations; 3-D reconstruction of flow structure from flow images; and interactive interpretation of flow images with theoretical simulations. These techniques can be employed in conjunction with: visual correlation and ensemble-averaging, both within a given image and between images; recognition of patterns from images; and estimates of velocity eigenfunctions from images.

  10. The Leading Edge 250: Oblique wing aircraft configuration project, volume 4

    NASA Technical Reports Server (NTRS)

    Schmidt, Andre; Moore, Peri; Nguyen, Dan; Oganesyan, Petros; Palmer, Charles

    1988-01-01

    The design of a high speed transport aircraft using the oblique wing concept as a part of the High Speed Civil Transport (HSCT) aircraft study is the Leading Edge 250 capable of travelling at Mach 4 with 250 passengers and has a 6,500 nautical mile range. Its innovation lies within its use of the unconventional oblique wing to provide efficient flight at any Mach number. Wave drag is kept to a minimum at high speed, while high lift is attained during critical takeoff and landing maneuvers by varying the sweep of the wing.

  11. 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.

  12. Measurement of Leading Edge Vortices from a Delta Wing Using a Three Component Laser Velocimeter

    NASA Technical Reports Server (NTRS)

    Meyers, James F.; Hepner, Timothy E.

    1988-01-01

    A demonstration of the capabilities of a three component laser velocimeter to provide a detailed experimental database of a complex flow field i s presented. The orthogonal three component laser velocimeter was used to measure the leading edge vortex flow field above a 75 degrees delta wing at angles-of-attack of 20.5 degrees and 40.0 degrees. The resulting mean velocity and turbulence intensity measurements are presented. The laser velocimeter is described in detail including a description of the data processing algorithm. A full error analysis was conducted and the results presented.

  13. Influence of a heated leading edge on boundary layer growth, stability, and transition

    SciTech Connect

    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.

  14. Influence of a heated leading edge on boundary layer growth, stability, and transition

    SciTech Connect

    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.

  15. 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.

  16. 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.

  17. 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.

  18. Space environmental effects on LDEF composites: Leading graphite/epoxy panel, selected trailing edge specimens

    NASA Technical Reports Server (NTRS)

    Dursch, Harry; George, Pete; Hill, Sylvester

    1992-01-01

    The composite electronics-module cover for the leading edge (row D9) experiment M0003-8 was fabricated from T300 graphite/934 epoxy unidirectional prepreg tape in a multi-oriented layup. This panel contained thermal control coatings in three of the four quadrants with the fourth quadrant left uncoated as a control. The composite experienced different thermal cycling extremes in each quadrant due to the differing optical properties of the coatings. Results will be presented on microcracking and other Low Earth Orbital (LEO) effects on the coated panel substrate.

  19. 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.

  20. Cascade tests of serrated leading edge blading at high subsonic speeds

    NASA Technical Reports Server (NTRS)

    Smith, E. G.

    1974-01-01

    Cascade tests of two-dimensional fan rotor blade rows were performed to investigate the effects of leading edge serration on acoustic and aerodynamic performance. The test configurations covered a range of serration tooth geometries. Tests were performed to investigate effects of inlet air angle and velocity on performance. Aerodynamic performance was determined by flow surveys at the mid-span of the blade exit. Acoustic performance was determined by wake turbulence surveys and sound measurements in the semireverberent exhaust chamber. Measured acoustic and aerodynamic performance was comparable and indicated that a serration length of about six percent blade chord yields minimum noise generation and minimum total pressure losses.

  1. 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.

  2. 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.

  3. Flow-induced noise control behind bluff bodies with various leading edges using the surface perturbation technique

    NASA Astrophysics Data System (ADS)

    Lu, Z. B.; Halim, D.; Cheng, L.

    2016-05-01

    The present paper is devoted to an investigation on the flow-induced noise control downstream of bluff bodies with various leading edges using the surface perturbation technique. Four typical leading edges used in various engineering applications were studied in this work: the semi-circular, square, 30° symmetric trapezoid and 30° asymmetric trapezoid leading edges. The surface perturbation was created by piezo-ceramic actuators embedded underneath the surface of a bluff body placed in a cross flow. To suppress the flow-induced noise downstream bluff bodies with those leading edges, the surface perturbation technique was implemented. Based on the experiments, a noise reduction in the duct of more than 14.0 dB has been achieved for all leading-edge cases. These results indicated that the vortex shedding and its flow-induced noise have been successfully suppressed by the proposed control scheme. The flow structure alteration around the bluff bodies and the shear layer shift phenomenon observed on the trailing edges were then investigated for interpreting the control mechanism for this flow-induced noise suppression, which were based on the vortex shedding strength suppression and vortex shedding frequency shift phenomenon. The effective control position for various leading edges was also studied for developing optimal control strategies for practical engineering applications.

  4. Low-order phenomenological modeling of leading-edge vortex formation

    NASA Astrophysics Data System (ADS)

    Wang, Chengjie; Eldredge, Jeff D.

    2013-09-01

    A low-order point vortex model for the two-dimensional unsteady aerodynamics of a flat plate wing section is developed. A vortex is released from both the trailing and leading edges of the flat plate, and the strength of each is determined by enforcing the Kutta condition at the edges. The strength of a vortex is frozen when it reaches an extremum, and a new vortex is released from the corresponding edge. The motion of variable-strength vortices is computed in one of two ways. In the first approach, the Brown-Michael equation is used in order to ensure that no spurious force is generated by the branch cut associated with each vortex. In the second approach, we propose a new evolution equation for a vortex by equating the rate of change of its impulse with that of an equivalent surrogate vortex with identical properties but constant strength. This impulse matching approach leads to a model that admits more general criteria for shedding, since the variable-strength vortex can be exchanged for its constant-strength surrogate at any instant. We show that the results of the new model, when applied to a pitching or perching plate, agree better with experiments and high-fidelity simulations than the Brown-Michael model, using fewer than ten degrees of freedom. We also assess the model performance on the impulsive start of a flat plate at various angles of attack. Current limitations of the model and extensions to more general unsteady aerodynamic problems are discussed.

  5. Noise generation by a supersonic leading edge. Part 1: general theory

    NASA Astrophysics Data System (ADS)

    Powles, C. J.

    2004-09-01

    This paper concerns the calculation of the sound fields which are generated when a vortical gust, convected in a supersonic mean flow, strikes the leading edge of a fan blade or aerofoil. Inviscid linear theory is applied, with the blade modelled as an infinite-span flat rigid plate, and a gust of arbitrary form is considered. By application of Fourier transforms the boundary value problem for the velocity potential is solved, leading to an integral expression for the generated sound field. This expression is applicable everywhere inside a Mach wedge. For gusts localized in the span direction, a farfield approximation is derived which is valid inside a Mach cone, and which is of simple enough form to be evaluated analytically for specified gust shapes. The new feature of this analysis is the consideration of an arbitrary gust form: previous authors have only ever considered the properties of specific gusts, focusing principally on harmonic gusts and jets.

  6. 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.

  7. Leading-edge receptivity to a vortical freestream disturbance: A numerical analysis

    NASA Technical Reports Server (NTRS)

    Buter, Thomas A.; Reed, Helen L.

    1991-01-01

    The receptivity to freestream vorticity of the boundary layer over a flat plate with an elliptic leading edge is investigated numerically. The flow is simulated by solving the incompressible Navier-Stokes system in general curvilinear coordinates with the vorticity and stream function as dependent variables. A finite-difference scheme which is second-order accurate in both space and time is used. As a first step, the steady basic-state solution is computed. Then a small amplitude vortical disturbance is introduced at the upstream boundary and the governing equations are solved time-accurately to evaluate the spatial and temporal growth of the perturbations leading to instability waves (Tollmien-Schlichting waves) inside the boundary layer. Preliminary results for a symmetric, 2-D disturbance reveal the presence of Tollmien-Schlichting waves aft of the flat-plate/ellipse juncture.

  8. 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.

  9. Numerical Study of the Effect of the Leading Edge Shape on Cavitation Around Inducer Blade Sections

    NASA Astrophysics Data System (ADS)

    Coutier-Delgosha, Olivier; Reboud, Jean-Luc; Fortes-Patella, Regiane

    A numerical study of the cavitation behaviour of two-dimensional hydrofoils simulating a section of an inducer blade is presented. Two leading edge shapes were chosen to approach rocket engine inducer designs. They were tested with respect to the development of sheet cavitation. The numerical model of cavitating flows is based on the 3D code FINE/TURBOTM, developed by NUMECA International. The cavitation process is taken into account by using a single fluid model, which considers the liquid vapour mixture as a homogeneous fluid whose density varies with respect to the static pressure. Numerical results are compared with experimental ones, obtained in the CREMHyG large cavitation tunnel(1). Pressure distributions along the foil suction side and the tunnel walls were measured for different cavity lengths. Total pressure measurements along the foil suction side allow characterizing the effects of cavitation on the liquid flow. Influence of the leading edge shape on the cavitation behaviour and comparison between experiments and numerical predictions are discussed.

  10. Leading edge film cooling enhancement for an inlet guide vane with fan-shaped holes

    NASA Astrophysics Data System (ADS)

    Liu, Jian-Jun; An, Bai-Tao; Liu, Jie; Zhan, W.

    2010-12-01

    This paper describes the improvement of leading edge film cooling effectiveness for a turbine inlet guide vane by using fan-shaped film cooling holes. The modification details are presented in comparison with the base-line configuration of cylindrical holes. Numerical simulations were carried out for the base-line and modified configurations by using CFX, in which the κ-ɛ turbulence model and scalable wall function were chosen. Contours of adiabatic film cooling effectiveness on the blade surfaces and span-wise distributions of film cooling effectiveness downstream the rows of cooling holes interested for the different cooling configurations were compared and discussed. It is showed that with the use of fan-shaped cooling holes around the leading edge, the adiabatic film cooling effectiveness can be enhanced considerably. In comparison with the cylindrical film cooling holes, up to 40% coolant mass flow can be saved by using fan-shaped cooling holes to obtain the comparable film cooling effectiveness for the studied inlet guide vane.

  11. 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.

  12. 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.

  13. Analog filtering methods improve leading edge timing performance of multiplexed SiPMs.

    PubMed

    Bieniosek, M F; Cates, J W; Grant, A M; Levin, C S

    2016-08-21

    Multiplexing many SiPMs to a single readout channel is an attractive option to reduce the readout complexity of high performance time of flight (TOF) PET systems. However, the additional dark counts and shaping from each SiPM cause significant baseline fluctuations in the output waveform, degrading timing measurements using a leading edge threshold. This work proposes the use of a simple analog filtering network to reduce the baseline fluctuations in highly multiplexed SiPM readouts. With 16 SiPMs multiplexed, the FWHM coincident timing resolution for single [Formula: see text] mm LYSO crystals was improved from 401  ±  4 ps without filtering to 248  ±  5 ps with filtering. With 4 SiPMs multiplexed, using an array of [Formula: see text] mm LFS crystals the mean time resolution was improved from 436  ±  6 ps to 249  ±  2 ps. Position information was acquired with a novel binary positioning network. All experiments were performed at room temperature with no active temperature regulation. These results show a promising technique for the construction of high performance multiplexed TOF PET readout systems using analog leading edge timing pickoff. PMID:27484131

  14. The Influence of Spanwise Flow on Leading-Edge Vortex Growth

    NASA Astrophysics Data System (ADS)

    Wong, Jaime; Kriegseis, Jochen; Rival, David

    2012-11-01

    It has been postulated that a spanwise component velocity through the core of a leading-edge vortex (LEV) can limit its growth and allow the LEV to remain attached to the wing. In the case of a delta wing, spanwise velocity is produced by wing sweep. However, in the case of flapping-wing flight, centripetal and Coriolis accelerations produce spanwise velocities which vary periodically. In order to understand the effect of various spanwise velocity profiles on LEV growth a simple analytical model for vortex growth has been developed. This model is based on the transport of vorticity-containing mass into the LEV through the leading-edge shear layer. By first neglecting spanwise effects, the model has been verified against a nominally two-dimensional plunging profile using Particle Image Velocimetry (PIV). With the addition of a spanwise transport of vorticity-containing mass, swept and flapping spanwise velocity profiles have been modeled and compared with three-dimensional, three-component velocity data collected using Particle Tracking Velocimetry.

  15. 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.

  16. The leading-edge stall of airfoils with various nose shapes

    NASA Astrophysics Data System (ADS)

    Kraljic, Matthew; Rusak, Zvi; Wang, Shixiao

    2015-11-01

    We study the inception of leading-edge stall on stationary, smooth thin airfoils with various nose shapes of the form xa (where 0 < a < 1 / 2) at low to moderately high chord Reynolds number flows. A reduced-order, multi-scale model problem is developed and solved using numerical simulations. The asymptotic theory demonstrates that a subsonic flow about a thin airfoil can be described in terms of an outer region, around most of the airfoil's chord, and an inner region, around the nose, that asymptotically match each other. The flow in the outer region is dominated by the classical thin airfoil theory. Scaled (magnified) coordinates and a modified (smaller) Reynolds number ReM are used to correctly account for the nonlinear behavior and extreme velocity changes in the inner region, where both the near-stagnation and high suction areas occur. The inner region problem is solved numerically to determine the inception of leading-edge stall on the nose. It is found that stall is delayed to higher angles of attack with the decrease of nose parameter a. Specifically, new airfoil shapes are proposed with increased stall angle at subsonic speeds and higher critical Mach numbers at transonic speeds.

  17. 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.

  18. 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.

  19. Leading edge vortex dynamics on a pitching delta wing. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Lemay, Scott P.

    1988-01-01

    The leading edge flow structure was investigated on a 70 deg flat plate delta wing which was pitched about its 1/2 chord position, to increase understanding of the high angle of attack aerodynamics on an unsteady delta wing. The wing was sinusoidally pitched at reduced frequencies ranging from k being identical with 2pi fc/u = 0.05 to 0.30 at chord Reynolds numbers between 90,000 and 350,000, for angle of attack ranges of alpha = 29 to 39 deg and alpha = 0 to 45 deg. The wing was also impulsively pitched at an approximate rate of 0.7 rad/s. During these dynamic motions, visualization of the leading edge vorticies was obtained by entraining titanium tetrachloride into the flow at the model apex. The location of vortex breakdown was recorded using 16mm high speed motion picture photography. When the wing was sinusoidally pitched, a hysteresis was observed in the location of breakdown position. This hysteresis increased with reduced frequency. The velocity of breakdown propagation along the wing, and the phase lag between model motion and breakdown location were also determined. When the wing was impulsively pitched, several convective times were required for the vortex flow to reach a steady state. Detailed information was also obtained on the oscillation of breakdown position in both static and dynamic cases.

  20. Analytical impact models and experimental test validation for the Columbia shuttle wing leading edge panels.

    SciTech Connect

    Lu, Wei-Yang; Metzinger, Kurt Evan; Gwinn, Kenneth West; Antoun, Bonnie R.; Korellis, John S.

    2004-10-01

    This paper describes the analyses and the experimental mechanics program to support the National Aeronautics and Space Administration (NASA) investigation of the Shuttle Columbia accident. A synergism of the analysis and experimental effort is required to insure that the final analysis is valid - the experimental program provides both the material behavior and a basis for validation, while the analysis is required to insure the experimental effort provides behavior in the correct loading regime. Preliminary scoping calculations of foam impact onto the Shuttle Columbia's wing leading edge determined if enough energy was available to damage the leading edge panel. These analyses also determined the strain-rate regimes for various materials to provide the material test conditions. Experimental testing of the reinforced carbon-carbon wing panels then proceeded to provide the material behavior in a variety of configurations and strain-rates for flown or conditioned samples of the material. After determination of the important failure mechanisms of the material, validation experiments were designed to provide a basis of comparison for the analytical effort. Using this basis, the final analyses were used for test configuration, instrumentation location, and calibration definition in support of full-scale testing of the panels in June 2003. These tests subsequently confirmed the accident cause.

  1. Analog filtering methods improve leading edge timing performance of multiplexed SiPMs

    NASA Astrophysics Data System (ADS)

    Bieniosek, M. F.; Cates, J. W.; Grant, A. M.; Levin, C. S.

    2016-08-01

    Multiplexing many SiPMs to a single readout channel is an attractive option to reduce the readout complexity of high performance time of flight (TOF) PET systems. However, the additional dark counts and shaping from each SiPM cause significant baseline fluctuations in the output waveform, degrading timing measurements using a leading edge threshold. This work proposes the use of a simple analog filtering network to reduce the baseline fluctuations in highly multiplexed SiPM readouts. With 16 SiPMs multiplexed, the FWHM coincident timing resolution for single 3~\\text{mm}× 3~\\text{mm}× 20 mm LYSO crystals was improved from 401  ±  4 ps without filtering to 248  ±  5 ps with filtering. With 4 SiPMs multiplexed, using an array of 3~\\text{mm}× 3~\\text{mm}× 20 mm LFS crystals the mean time resolution was improved from 436  ±  6 ps to 249  ±  2 ps. Position information was acquired with a novel binary positioning network. All experiments were performed at room temperature with no active temperature regulation. These results show a promising technique for the construction of high performance multiplexed TOF PET readout systems using analog leading edge timing pickoff.

  2. 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.

  3. 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

  4. 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.

  5. Segregation of leading-edge and uropod components into specific lipid rafts during T cell polarization

    PubMed Central

    Gómez-Moutón, Concepción; Abad, Jose Luis; Mira, Emilia; Lacalle, Rosa Ana; Gallardo, Eduard; Jiménez-Baranda, Sonia; Illa, Isabel; Bernad, Antonio; Mañes, Santos; Martínez-A., Carlos

    2001-01-01

    Redistribution of specialized molecules in migrating cells develops asymmetry between two opposite cell poles, the leading edge and the uropod. We show that acquisition of a motile phenotype in T lymphocytes results in the asymmetric redistribution of ganglioside GM3- and GM1-enriched raft domains to the leading edge and to the uropod, respectively. This segregation to each cell pole parallels the specific redistribution of membrane proteins associated to each raft subfraction. Our data suggest that raft partitioning is a major determinant for protein redistribution in polarized T cells, as ectopic expression of raft-associated proteins results in their asymmetric redistribution, whereas non-raft-partitioned mutants of these proteins are distributed homogeneously in the polarized cell membrane. Both acquisition of a migratory phenotype and SDF-1α-induced chemotaxis are cholesterol depletion-sensitive. Finally, GM3 and GM1 raft redistribution requires an intact actin cytoskeleton, but is insensitive to microtubule disruption. We propose that membrane protein segregation not only between raft and nonraft domains but also between distinct raft subdomains may be an organizational principle that mediates redistribution of specialized molecules needed for T cell migration. PMID:11493690

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. Leading-edge vortex trajectories under the influence of Coriolis acceleration

    NASA Astrophysics Data System (ADS)

    Limacher, Eric; Morton, Chris; Wood, David

    2015-11-01

    Leading-edge vortices (LEVs) can form and remain attached to a rotating wing indefinitely, but the mechanisms of stable attachment are not well understood. Taking for granted that such stable structures do form, a practical question arisesof where an LEV core persists in the body-fixed frame of reference. Noting that span-wise flow exists within the LEV core, it is apparent that a mean streamline aligned with the axis of the LEV must exist. The present work usesthe Navier-Stokes equations alongthis steady, axial streamline in order to consider the accelerations that act in the steamline-normal direction to affect its local curvature. With some simplifying assumptions, an ordinary differential equation is derivedthat describes the trajectory of the axial streamline through the vortex core. Usingempirical values of axial velocity in the vortex core from previous studies, it can be shown that Coriolis and centrifugal forces alone can account for the tilting of the stable LEV into the wake within several chord lengths from the root of rotationin the span-wise direction. This result supports an hypothesisthat LEVs are observed at inboard locations because Coriolis force must actover a finite distance to tilt the stable LEV away from the leading edge.

  11. 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.

  12. 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.

  13. Atmospheric Entry Aerothermodynamics Flight Test on CubeSat Platform

    NASA Astrophysics Data System (ADS)

    Sakraker, I.; Umit, E.; van der Haegen, V.; Chazot, O.

    2014-06-01

    The challenging aerothermochemistry of atmospheric entry is aimed to be experimented on a triple CubeSat platform having ablative TPS in the front unit and ceramic TPS on the side panels. Five aerothermodynamics payloads are presented in this paper.

  14. 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

  15. Formation of Leading-Edge Pinholes in the Space Shuttle Wings Investigated

    NASA Technical Reports Server (NTRS)

    Jacobson, Nathan S.

    2000-01-01

    The space shuttle wing leading edge and nose cap are composed of a carbon/carbon composite that is protected by silicon carbide. The coefficient of thermal expansion mismatch leads to cracks in the silicon carbide. The outer coating of the silicon carbide is a sodium-silicate-based glass that becomes fluid at the shuttles high reentry temperatures and fills these cracks. Small pinholes roughly 0.1 mm in diameter have been observed on these materials after 12 or more flights. These pinholes have been investigated by researchers at the NASA Johnson Space Center, Rockwell International, the Boeing Company, Lockheed Martin Corporation, and the NASA Glenn Research Center at Lewis Field to determine the possible sources and the extent of damage. A typical pinhole is illustrated in the photomicrographs. These pinholes are found primarily on the wing leading edges and not on the nose cap, which is covered when the orbiter is on the launch pad. The pinholes are generally associated with a bead of zincrich glass. Examination of the orbiter and launch structure indicates that weathering paint on the launch structure leads to deposits of zinc-containing paint flakes on the wing leading edge. These may become embedded in the crevices of the wing leading edge and form the observed zinc-rich glass. Laboratory experiments indicate that zinc oxide reacts vigorously with the glass coating on the silicon carbide. Thus, it is likely that this is the reaction that leads to pinhole formation (Christensen, S.V.: Reinforced Carbon/Carbon Pin Hole Formation Through Zinc Oxide Attack. Rockwell International Internal Letter, RDW 96 057, May 1996). Cross-sectional examination of pinholes suggests that they are enlarged thermal expansion mismatch cracks. This is illustrated in the photomicrographs. A careful microstructural analysis indicates that the pinhole walls consist of layers of zinc-containing glass. Thus, pinholes are likely formed by zinc oxide particles lodging in crevices and

  16. 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.

  17. Shuttle Tethered Aerothermodynamics Research Facilty (STARFAC) instrumentation requirements

    NASA Technical Reports Server (NTRS)

    Wood, G. M.; Siemers, P. M.; Carlomagno, G. M.; Hoffman, J.

    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.

  18. Shuttle entry technology payloads. [for aerothermodynamic research

    NASA Technical Reports Server (NTRS)

    Siemers, P. M., III

    1975-01-01

    The flight frequency of the Space Transportation System (STS) coupled with its large payload-carrying capability will provide an unprecedented opportunity for conducting aerothermodynamic/entry technology research. This STS research opportunity can be characterized into two distinct categories: (1) that research which will utilize the STS orbiter as the test vehicle, and (2) that research which will utilize a vehicle launched from the orbiter for entry. To date, on-going studies have defined experiments as well as the support systems required for the shuttle launched research program. The proposed Entry Technology Program will provide a flight data base from which accurate correlations can be performed relative to ground test and analysis data. These correlations will result in optimized designs for future flight systems.

  19. 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.

  20. 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.

  1. Experimental Pressure Distributions over Wing Tips at Mach Number 1.9 I : Wing Tip with Subsonic Leading Edge

    NASA Technical Reports Server (NTRS)

    Jagger, James M; Mirels, Harold

    1949-01-01

    An investigation was conducted at a Mach number of 1.91 to determine spanwise pressure distribution over a wing tip in a region influenced by a sharp subsonic leading edge swept back at 70 degrees. Except for pressure distribution on the top surface in the immediate vicinity of the subsonic leading edge, the maximum difference between linearized theory and experimental data was 2 1/2 percent (of free-stream dynamic pressure) for angles of attack up to 4 degrees and 7 percent for angles of attack up to 8 degrees. Pressures on the top surface nearest the subsonic edge indicated local expansions beyond values predicted by linearized theory.

  2. Aerodynamic Forces and Loadings on Symmetrical Circular-Arc Airfoils with Plain Leading-Edge and Plain Trailing-Edge Flaps

    NASA Technical Reports Server (NTRS)

    Cahill, Jones F; Underwood, William J; Nuber, Robert J; Cheesman, Gail A

    1953-01-01

    An investigation has been made in the Langley two-dimensional low-turbulence tunnel and in the Langley two-dimensional low-pressure tunnel of 6- and 10-percent-thick symmetrical circular-arc airfoil sections at low Mach numbers and several Reynolds numbers. The airfoils were equipped with 0.15-chord plain leading-edge flaps and 0.20-chord plan trailing-edge flaps. The section lift and pitching-moment characteristics were determined for both airfoils with the flaps deflected individually and in combination. The section drag characteristics were obtained for the 6-percent-thick airfoil with the flaps partly deflected as low-drag-control flaps and for airfoils with the flaps neutral. Surface pressures were measured on the 6-percent-thick airfoil section with the flaps deflected either individually or in appropriate combination to furnish flap load and hinge-moment data applicable to the structural design of the airfoil. A generalized method is developed that permits the determination of the chordwise pressure distribution over sharp-edge airfoils with plain leading-edge flaps and plain trailing-edge flaps of arbitrary size and deflection.

  3. 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

  4. 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.

  5. Open-Type Separation on Delta Wings for Leading-Edge Bluntness

    NASA Astrophysics Data System (ADS)

    Chiba, Kazuhisa; Obayashi, Shigeru; Nakahashi, Kazuhiro

    A numerical simulation was carried out corresponding to recent experiments using delta wings with sharp and blunt leading-edges, which indicates the second primary vortex, at NASA Langley Research Center. However, the experimental data did not reveal the detailed physical phenomena regarding the second primary vortex, because the experiment used only on-the-body-surface data. In the present study, the physical phenomena were revealed using Reynolds-averaged Navier-Stokes computations with three one-equation turbulence models on an unstructured hybrid mesh. The adaptive mesh refinement method in the vicinity of the vortex center was also applied to have more mesh resolution. Consequently, the result quantitatively revealed that appropriate modeling regarding turbulent kinematic viscosity was significant. Moreover, the three-dimensional visualization of the computational fluid dynamics results suggested that the second primary vortex was a developing shear layer merging into an open-type separation generated late by the primary vortex.

  6. An experimental study of turbine vane heat transfer with leading edge and downstream film cooling

    NASA Technical Reports Server (NTRS)

    Nirmalan, V.; Hylton, L. D.

    1989-01-01

    This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine-vane external heat transfer. Steady-state experimental measurements were made in a three-vane linear two-dimensional cascade. The principal independent parameters were maintained over ranges consistent with actual engine conditions. 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. The data obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil.

  7. 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

  8. An experimental study of turbine vane heat transfer with leading edge and downstream film cooling

    NASA Astrophysics Data System (ADS)

    Nirmalan, V.; Hylton, L. D.

    1989-06-01

    This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine-vane external heat transfer. Steady-state experimental measurements were made in a three-vane linear two-dimensional cascade. The principal independent parameters were maintained over ranges consistent with actual engine conditions. 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. The data obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil.

  9. Multiple leading edge vortices of unexpected strength in freely flying hawkmoth

    NASA Astrophysics Data System (ADS)

    Johansson, L. Christoffer; Engel, Sophia; Kelber, Almut; Heerenbrink, Marco Klein; Hedenström, Anders

    2013-11-01

    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.

  10. Leading edge serrations which reduce the noise of low-speed rotors

    NASA Technical Reports Server (NTRS)

    Soderman, P. T.

    1973-01-01

    Acoustic effects of serrated brass strips attached near the leading edges of two different size rotors were investigated. The two bladed rotors were tested in hover. Rotor rotational speed, blade angle, serration shape, and serration position were varied. The serrations were more effective as noise suppressors at rotor tip speeds less than 135 m/sec (444 ft/sec) than at higher speeds. high frequency noise was reduced but the low frequency rotational noise was little affected. Noise reductions from 4 to 8 db overall sound pressure level and 3 to 17 db in the upper octave bands were achieved on the 1.52 m (5.0 ft) diameter rotor. Noise reductions up to 4 db overall sound pressure level were measured for the 2.59 m (8.5 ft) diameter rotor at some conditions.

  11. Fabrication of a graphite/epoxy composite leading edge for laminar flow control

    NASA Technical Reports Server (NTRS)

    Beall, R. T.

    1980-01-01

    Lockheed, under NASA contract, has recently completed the first phase of a program to evaluate laminar flow control concepts for transport aircraft. Achievement of laminar flow over a wing surface requires a system of slots, metering holes, ducts and pumps to be used to remove the turbulent air adjacent to the surface. This requirement poses severe restrictions on conventional metallic structure. Graphite/epoxy composite with its unique properties appears to be the material that might solve the very complex structural problems associated with a laminar flow control aircraft. A six-foot span graphite/epoxy test article incorporating provisions for leading edge cleaning, deicing and laminar flow control was designed, fabricated and tested.

  12. An experimental study of the flow in a wing with a partial span drooped leading edge

    NASA Technical Reports Server (NTRS)

    Winkelmann, A. E.; Tsao, C. P.

    1981-01-01

    The flow field produced by a low aspect ratio wing (AR = 3.0) with a partial span leading edge droop was investigated in a series of low speed wind tunnel tests (Reynolds number based on 17.8 cm chord = 560,000). Photographs were obtained of surface oil flow patterns over an angle of attack range of alpha = 0 to 29 deg. Flow field surveys of the partially stalled wing at alpha = 25 deg were completed using a hot-wire probe, a split-film probe and a Conrad probe. The flow field survey data was presented using a color video display. The data indicated regions of apparent reversed flow in the separation region behind the wing and indicated the general cross-sectional shape of the separated wake flow.

  13. Application of superplastically formed and diffusion bonded aluminum to a laminar flow control leading edge

    NASA Technical Reports Server (NTRS)

    Goodyear, M. D.

    1987-01-01

    NASA sponsored the Aircraft Energy Efficiency (ACEE) program in 1976 to develop technologies to improve fuel efficiency. Laminar flow control was one such technology. Two approaches for achieving laminar flow were designed and manufactured under NASA sponsored programs: the perforated skin concept used at McDonnell Douglas and the slotted design used at Lockheed-Georgia. Both achieved laminar flow, with the slotted design to a lesser degree (JetStar flight test program). The latter design had several fabrication problems concerning springback and adhesive flow clogging the air flow passages. The Lockheed-Georgia Company accomplishments is documented in designing and fabricating a small section of a leading edge article addressing a simpler fabrication method to overcome the previous program's manufacturing problems, i.e., design and fabrication using advanced technologies such as diffusion bonding of aluminum, which has not been used on aerospace structures to date, and the superplastic forming of aluminum.

  14. 3D characterization of leading-edge vortex formation and growth

    NASA Astrophysics Data System (ADS)

    Onoue, Kyohei; Breuer, Kenneth

    2015-11-01

    We examine the vorticity transport mechanisms responsible for regulating the stability and strength of the leading-edge vortex (LEV) on rapidly pitching plates with different planforms (swept vs. rectangular) in a uniform airflow. All experiments are carried out using a cyber-physical experimental setup (Onoue et al., 2015, JFS vol. 55) and synchronized 3D PIV measurements. In the case of a swept wing, two distinct regions of intense spanwise flow are observed around the LEV centroid--a feature conspicuously absent on a rectangular pitching plate. The interaction between these spanwise flows and the LEV core seems to play a role in prolonging the LEV residence time at the cost of the vortex circulation growth rate and magnitude. Detailed control volume analysis is performed to elucidate the flow physics at work. This research is funded by the Air Force Office of Scientific Research (AFOSR).

  15. 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.

  16. 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.

  17. Leading-edge transition and relaminarization phenomena on a subsonic high-lift system

    NASA Technical Reports Server (NTRS)

    Van Dam, C. P.; Vijgen, P. M. H. W.; Yip, L. P.; Potter, R. C.

    1993-01-01

    Boundary-layer transition and relaminarization may have a critical effect on the flow development about multielement high-lift systems of subsonic transport aircraft with swept wings. The purpose of this paper is a study of transition phenomena in the leading-edge region of the various elements of a high-lift system. The flow phenomena studied include transition of the attachment-line flow, relaminarization, and crossflow instability. The calculations are based on pressure distributions measured in flight on the NASA Transport Systems Research Vehicle (Boeing 737-100) at a wing station where the flow approximated infinite swept wing conditions. The results indicate that significant regions of laminar flow can exist on all flap elements in flight. In future flight experiments the extent of these regions will be measured, and the transition mechanisms and the effect of laminar flow on the high-lift characteristics of the multi-element system will be further explored.

  18. Effect of leading edge sweep on shock-shock interference at Mach 8

    NASA Technical Reports Server (NTRS)

    Glass, Christopher E.; Wieting, Allan R.; Holden, Michael S.

    1989-01-01

    These Mach 8 experimental results are applicable to the details of a shock-shock interference that may occur on an engine inlet of a hypersonic vehicle from a swept forebody shock interacting with a swept cowl leading edge bow shock or from a swept splitter plate shock interacting with a swept fuel injection strut bow shock. 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 deg and 30 deg swept results with the 0 deg swept results shows 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.

  19. Performance Evaluation of Leading Edge Slats on Rigid Wing Sail Catamarans

    NASA Astrophysics Data System (ADS)

    Johnson, Chelsea; O'Neill, Charles

    2015-11-01

    Rigid wing sails have created the fastest catamarans in history, however with the addition of a leading edge slat higher lift and faster speeds may be achieved. Slats are currently used on airplane wings to increase lift, but have not been implemented on a rigid wing sail catamaran. Using 3D modeling and computational fluid dynamics software, this research investigates the effect that slats have on the performance of rigid wing sail catamarans. Aerodynamics and hydrodynamics form the basis of the research. The preliminary results show an increase in the coefficient of lift for sail models with slats over sail models without slats, allowing the catamaran to perform at higher speeds. The ability of the slat to rotate has also been identified as a key factor in increasing the benefit of the slat. This work was supported by NSF site award 1358991.

  20. 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.

  1. Visualization and flow surveys of the leading edge vortex structure on delta wing planforms

    NASA Technical Reports Server (NTRS)

    Payne, F. M.; Ng, T. T.; Nelson, R. C.; Schiff, L. B.

    1986-01-01

    In the present experimental investigation of thin delta wing vortex breakdown, for the cases of sweep angles of 70, 75, 80, and 85 deg, and smoke flow visualization/laser light sheet technique is used to obtain cross sectional views of the leading edge vortices as they break down. A combination of lateral and longitudinal cross sectional views furnishes data on the three-dimensional character of the vortex before, during, and after breakdown. Velocity measurements conducted with a laser Doppler anemometer on the 70 deg sweep delta, at 30 deg angle-of-attack, indicate that when breakdown occurs the core flow is transformed from a jet-like to a wake-like flow.

  2. A mechanism for mitigation of blade-vortex interaction using leading edge blowing flow control

    NASA Astrophysics Data System (ADS)

    Weiland, Chris; Vlachos, Pavlos P.

    2009-09-01

    The interaction of a vortical unsteady flow with structures is often encountered in engineering applications. Such flow structure interactions (FSI) can be responsible for generating significant loads and can have many detrimental structural and acoustic side effects, such as structural fatigue, radiated noise and even catastrophic results. Amongst the different types of FSI, the parallel blade-vortex interaction (BVI) is the most common, often encountered in helicopters and propulsors. In this work, we report on the implementation of leading edge blowing (LEB) active flow control for successfully minimizing the parallel BVI. Our results show reduction of the airfoil vibrations up to 38% based on the root-mean-square of the vibration velocity amplitude. This technique is based on displacing an incident vortex using a jet issued from the leading edge of a sharp airfoil effectively increasing the stand-off distance of the vortex from the body. The effectiveness of the method was experimentally analyzed using time-resolved digital particle image velocimetry (TRDPIV) recorded at an 800 Hz rate, which is sufficient to resolve the spatio-temporal dynamics of the flow field and it was combined with simultaneous accelerometer measurements of the airfoil, which was free to oscillate in a direction perpendicular to the freestream. Analysis of the flow field spectra and a Proper Orthogonal Decomposition (POD) of the TRDPIV data of the temporally resolved planar flow fields indicate that the LEB effectively modified the flow field surrounding the airfoil and increased the convecting vortices stand-off distance for over half of the airfoil chord length. It is shown that LEB also causes a redistribution of the flow field spectral energy over a larger range of frequencies.

  3. 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.

  4. An Attached Flow Design of a Noninterferring Leading Edge Extension to a Thick Delta Wing

    NASA Technical Reports Server (NTRS)

    Lamar, John E.; Ghaffari, Farhad

    1985-01-01

    An analytical procedure for the determination of the shape of a Leading-Edge Extension (LEE) which satisfies design criteria, including especially noninterference at the wing design point, has been developed for thick delta wings. The LEE device best satisfying all criteria is designed to be mounted on a wing along a dividing stream surface associated with an attached flow design lift coefficient (C(sub L,d)) of greater than zero. This device is intended to improve the aerodynamic performance of transonic aircraft at C(sub L) greater than C(sub L,d) system emanating from the LEE leading edge. In order to quantify this process a twisted and cambered thick delta wing was chosen for the initial application of this design procedure. Appropriate computer codes representing potential and vortex flows were employed to determine the dividing stream surface at C(sub L,d) and an optimized LEE planform shape at C(sub L) greater than C(sub L,d), respectively. To aid in the LEE selection, the aerodynamic effectiveness of 36 planforms was investigated at C(sub L) greater than C(sub L,d). This study showed that reducing the span of the candidate LEEs has the most detrimental effect on overall aerodynamic efficiency, regardless of the shape or area. Furthermore, for a fixed area, constant-chord LEE candidates were relatively more efficient than those with sweep less than the wing. At C(sub L,d), the presence of the LEE planform best satisfying the design criteria was found to have no effect on the wing alone aerodynamic performance.

  5. Vibration and sound of an elastic wing actuated at its leading edge

    NASA Astrophysics Data System (ADS)

    Manela, A.

    2012-01-01

    The motion and sound of a thin elastic plate, subject to uniform low-Mach flow and actuated at its leading edge, is studied. The linearized response to arbitrary small-amplitude translation and rotation is analyzed using Fourier decomposition of the forcing signal. Both periodic (sinusoidal) and non-periodic ("step-jump") actuations are investigated. When the frequency spectrum of the forcing signal contains an eigenfrequency Ωres of the unforced system, a resonance motion is excited and the plate oscillates at the corresponding eigenmode. The dynamical description is applied to formulate the acoustic problem, where the sources of sound include the plate velocity and fluid vorticity. Acoustic radiation of a dipole type is calculated and discussed in the limit where the plate is acoustically compact. In the case of sinusoidal excitation, plate elasticity has two opposite effects on sound radiation, depending on the forcing frequency: at frequencies close to Ωres, the near-resonance motion results in the generation of high sound levels; however, at frequencies far from Ωres, plate elasticity reduces the amplitude of plate deflection (compared to that of a rigid plate), leading to noise reduction. In the case of non-periodic actuation, the plate-fluid system amplifies those frequencies that are closest to Ωres, which, in turn, dominate the acoustic signature. The results identify the trailing edge noise as the main source of sound, dominating the sound generated by direct plate motion. We suggest the present theory as a preliminary tool for examining the acoustic signature of flapping flight, common in insects and flapping micro-air-vehicles.

  6. Control of the separated flow around an airfoil using a wavy leading edge inspired by humpback whale flippers

    NASA Astrophysics Data System (ADS)

    Favier, Julien; Pinelli, Alfredo; Piomelli, Ugo

    2012-01-01

    The influence of spanwise geometrical undulations of the leading edge of an infinite wing is investigated numerically at low Reynolds number, in the context of passive separation control and focusing on the physical mechanisms involved. Inspired by the tubercles of the humpback whale flippers, the wavy leading edge is modeled using a spanwise sinusoidal function whose amplitude and wavelength constitute the parameters of control. A direct numerical simulation is performed on a NACA0020 wing profile in a deep stall configuration ( α=20°), with and without the presence of the leading edge waviness. The complex solid boundaries obtained by varying the sinusoidal shape of the leading edge are modeled using an immersed boundary method (IBM) recently developed by the authors [Pinelli et al., J. Comput. Phys. 229 (2010) 9073-9091]. A particular set of wave parameters is found to change drastically the topology of the separated zone, which becomes dominated by streamwise vortices generated from the sides of the leading edge bumps. A physical analysis is carried out to explain the mechanism leading to the generation of these coherent vortical structures. The role they play in the control of boundary layer separation is also investigated, in the context of the modifications of the hydrodynamic performances which have been put forward in the literature in the last decade.

  7. 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.

  8. 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.

  9. 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.

  10. Prospore membrane formation linked to the leading edge protein (LEP) coat assembly

    PubMed Central

    Moreno-Borchart, Alexandra C.; Strasser, Katrin; Finkbeiner, Martin G.; Shevchenko, Anna; Shevchenko, Andrej; Knop, Michael

    2001-01-01

    In yeast, the differentiation process at the end of meiosis generates four daughter cells inside the boundaries of the mother cell. A meiosis-specific plaque (MP) at the spindle pole bodies (SPBs) serves as the starting site for the formation of the prospore membranes (PSMs) that are destined to encapsulate the post-meiotic nuclei. Here we report the identification of Ady3p and Ssp1p, which are functional components of the leading edge protein (LEP) coat, that covers the ring-shaped opening of the PSMs. Ssp1p is required for the assembly of the LEP coat, which consists of at least three proteins (Ssp1p, Ady3p and Don1p). The assembly of the LEP coat starts with the formation of cytosolic precursors, which then bind in an Ady3p-dependent manner to the SPBs. Subsequent processes at the SPBs leading to functional LEP coats require Ssp1p and the MP components. During growth of the PSMs, the LEP coat functions in formation of the cup-shaped membrane structure that is indispensable for the regulated cellularization of the cytoplasm around the post-meiotic nuclei. PMID:11742972

  11. 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.

  12. Spin-dependent transport for armchair-edge graphene nanoribbons between ferromagnetic leads.

    PubMed

    Zhou, Benhu; Chen, Xiongwen; Zhou, Benliang; Ding, Kai-He; Zhou, Guanghui

    2011-04-01

    We theoretically investigate the spin-dependent transport for the system of an armchair-edge graphene nanoribbon (AGNR) between two ferromagnetic (FM) leads with arbitrary polarization directions at low temperatures, where a magnetic insulator is deposited on the AGNR to induce an exchange splitting between spin-up and -down carriers. By using the standard nonequilibrium Green's function (NGF) technique, it is demonstrated that the spin-resolved transport property for the system depends sensitively on both the width of AGNR and the polarization strength of FM leads. The tunneling magnetoresistance (TMR) around zero bias voltage possesses a pronounced plateau structure for a system with semiconducting 7-AGNR or metallic 8-AGNR in the absence of exchange splitting, but this plateau structure for the 8-AGNR system is remarkably broader than that for the 7-AGNR one. Interestingly, an increase of the exchange splitting Δ suppresses the amplitude of the structure for the 7-AGNR system. However, the TMR is much enhanced for the 8-AGNR system under a bias amplitude comparable to the splitting strength. Further, the current-induced spin-transfer torque (STT) for the 7-AGNR system is systematically larger than that for the 8-AGNR one. The findings here suggest the design of GNR-based spintronic devices by using a metallic AGNR, but it is more favorable to fabricate a current-controlled magnetic memory element by using a semiconducting AGNR. PMID:21415476

  13. 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.

  14. 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.

  15. A simplified method for thermal analysis of a cowl leading edge subject to intense local shock-wave-interference heating

    NASA Technical Reports Server (NTRS)

    Mcgowan, David M.; Camarda, Charles J.; Scotti, Stephen J.

    1992-01-01

    Type IV shock wave interference heating on a blunt body causes extremely intense heating over a very localized region of the body. An analytical solution is presented to a heat transfer problem that approximates the shock wave interference heating of an engine cowl leading edge of the National Aero-Space Plane. The problem uses a simplified geometry to represent the leading edge. An analytical solution is developed that provides a means for approximating maximum temperature differences between the outer and inner surface temperatures of the leading edge. The solution is computationally efficient and, as a result, is well suited for conceptual and preliminary design or trade studies. Transient and steady state analyses are conducted, and results obtained from the analytical solution are compared with results of 2-D thermal finite element analyses over a wide range of design parameters. Isotropic materials as well as laminated composite materials are studied. Results of parametric studies are presented to indicate the effects of the thickness of the cowl leading edge and the width of the region heated by the shock wave interference on the thermal response of the leading edge.

  16. 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

  17. 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.

  18. 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.

  19. Aero-Thermo-Dynamic Mass Analysis

    PubMed Central

    Shiba, Kota; Yoshikawa, Genki

    2016-01-01

    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. PMID:27412335

  20. Aeroassist flight experiment aerodynamics and aerothermodynamics

    NASA Technical Reports Server (NTRS)

    Brewer, Edwin B.

    1989-01-01

    The problem is to determine the transitional flow aerodynamics and aerothermodynamics, including the base flow characteristics, of the Aeroassist Flight Experiment (AFE). The justification for the computational fluid dynamic (CFD) Application stems from MSFC's system integration responsibility for the AFE. To insure that the AFE objectives are met, MSFC must understand the limitations and uncertainties of the design data. Perhaps the only method capable of handling the complex physics of the rarefied high energy AFE trajectory is Bird's Direct Simulation Monte Carlo (DSMC) technique. The 3-D code used in this analysis is applicable only to the AFE geometry. It uses the Variable Hard Sphere (VHS) collision model and five specie chemistry model available from Langley Research Center. The code is benchmarked against the AFE flight data and used as an Aeroassisted Space Transfer Vehicle (ASTV) design tool. The code is being used to understand the AFE flow field and verify or modify existing design data. Continued application to lower altitudes is testing the capability of the Numerical Aerodynamic Simulation Facility (NASF) to handle 3-D DSMC and its practicality as an ASTV/AFE design tool.

  1. 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.

  2. Aero-Thermo-Dynamic Mass Analysis.

    PubMed

    Shiba, Kota; Yoshikawa, Genki

    2016-01-01

    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. PMID:27412335

  3. Preservation of wing leading edge suction at the plane of symmetry as a factor in wing-fuselage design

    NASA Technical Reports Server (NTRS)

    Larrabee, E. E.

    1975-01-01

    Most fuselage geometries cover a portion of the wing leading edge near the plane of symmetry, and it seems reasonable to expect that a large fraction of the leading edge suction which would be developed by the covered wing at high angles of attack is not developed on the fuselage. This is one of the reasons that the Oswald span efficiency factor for the wing body combination fails to approach the value predicted by lifting line theory for the isolated wing. Some traditional and recent literature on wing-body interference is discussed and high Reynolds number data on wing-body-nacelle drag are reviewed. An exposed central leading edge geometry has been developed for a sailplane configuration. Low Reynolds number tests have not validated the design concept.

  4. Influence of high-intensity turbulence on laminar boundary layer development on a cylindrical leading edge: Enhancement to eddy diffusivity

    NASA Astrophysics Data System (ADS)

    Pearson, Juli K.

    The growing demand for increased efficiency in turbine engine designs has sparked a growing interest for research of air flow around curved surfaces. The turbine's operating conditions result in material property constraints, especially in the first stage turbine vanes and blades. These turbine vane components experience extreme loading conditions of both high temperature and high turbulence intensities exiting the combustor. The surface of the turbine blades has cylindrical leading edges that promote stabilizing flow accelerations. These convex surfaces can cause a reduced eddy diffusivity across the boundary layer. This thesis reviews measurements of velocity and turbulence intensities taken just shy of the thirty degrees offset from the stagnation line of a two-dimensional cylindrical leading edge under a wide range of turbulence and flow conditions flow conditions. Flow conditions and velocity measurements were gathered with respect to the distance to the surface. The length of the measurements extended from the surface to beyond the boundary layer's edge. The instrumentation used to collect data was a single wire driven by a constant temperature anemometer bridge. The hot wire is specially modified to measure data near the cylindrical leading edges curved surface. The traversing system allowed the acquisition of high-resolution boundary layer data. The traversing system was installed internally to the cylindrical leading edge to reduce probe blockage.

  5. 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

  6. Drag Coefficient of Water Droplets Approaching the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida Garcia

    2013-01-01

    This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Results are presented and discussed for drag coefficients of droplets with diameters in the range of 300 to 1800 micrometers, and airfoil velocities of 50, 70 and 90 meters/second. The effect of droplet oscillation on the drag coefficient is discussed.

  7. Pressure-Velocity Correlations in the Cove of a Leading Edge Slat

    NASA Astrophysics Data System (ADS)

    Wilkins, Stephen; Richard, Patrick; Hall, Joseph

    2015-11-01

    One of the major sources of aircraft airframe noise is related to the deployment of high-lift devices, such as leading-edge slats, particularly when the aircraft is preparing to land. As the engines are throttled back, the noise produced by the airframe itself is of great concern, as the aircraft is low enough for the noise to impact civilian populations. In order to reduce the aeroacoustic noise sources associated with these high lift devices for the next generation of aircraft an experimental investigation of the correlation between multi-point surface-mounted fluctuating pressures measured via flush-mounted microphones and the simultaneously measured two-component velocity field measured via Particle Image Velocimetry (PIV) is studied. The development of the resulting shear-layer within the slat cove is studied for Re =80,000, based on the wing chord. For low Mach number flows in air, the major acoustic source is a dipole acoustic source tied to fluctuating surface pressures on solid boundaries, such as the underside of the slat itself. Regions of high correlations between the pressure and velocity field near the surface will likely indicate a strong acoustic dipole source. In order to study the underlying physical mechanisms and understand their role in the development of aeroacoustic noise, Proper Orthogonal Decomposition (POD) by the method of snapshots is employed on the velocity field. The correlation between low-order reconstructions and the surface-pressure measurements are also studied.

  8. Control of Flow Structure on Low Swept Delta Wing with Steady Leading Edge Blowing

    NASA Astrophysics Data System (ADS)

    Ozturk, Ilhan; Zharfa, Mohammadreza; Yavuz, Mehmet Metin

    2014-11-01

    Interest in unmanned combat air vehicles (UCAVs) and micro air vehicles (MAVs) has stimulated investigation of the flow structure, as well as its control, on delta wings having low and moderate values of sweep angle. In the present study, the flow structure is characterized on a delta wing of low sweep 35-degree angle, which is subjected to steady leading edge blowing. The techniques of laser illuminated smoke visualization, laser Doppler anemometry (LDA), and surface pressure measurements are employed to investigate the steady and unsteady nature of the flow structure on delta wing, in relation to the dimensionless magnitude of the blowing coefficient. Using statistics and spectral analysis, unsteadiness of the flow structure is studied in detail. Different injection locations are utilized to apply different blowing patterns in order to identify the most efficient control, which provides the upmost change in the flow structure with the minimum energy input. The study aims to find the optimum flow control strategy to delay or to prevent the stall and possibly to reduce the buffeting on the wing surface. Since the blowing set-up is computer controlled, the unsteady blowing patterns compared to the present steady blowing patterns will be studied next. This project was supported by the Scientific and Technological Research Council of Turkey (Project Number: 3501 111M732).

  9. Suppression of dynamic stall with a leading-edge slat on a VR-7 airfoil

    NASA Technical Reports Server (NTRS)

    Mcalister, K. W.; Tung, C.

    1993-01-01

    The VR-7 airfoil was experimentally studied with and without a leading-edge slat at fixed angles of attack from 0 deg to 30 deg at Re = 200,000 and for unsteady pitching motions described by alpha equals alpha(sub m) + 10 deg(sin(wt)). The models were two dimensional, and the test was performed in a water tunnel at Ames Research Center. The unsteady conditions ranged over Re equals 100,000 to 250,000, k equals 0.001 to 0.2, and alpha(sub m) = 10 deg to 20 deg. Unsteady lift, drag, and pitching-moment measurements were obtained along with fluorescent-dye flow visualizations. The addition of the slat was found to delay the static-drag and static-moment stall by about 5 degrees and to eliminate completely the development of a dynamic-stall vortex during unsteady motions that reached angles as high as 25 degrees. In all of the unsteady cases studied, the slat caused a significant reduction in the force and moment hysteresis amplitudes. The reduced frequency was found to have the greatest effect on the results, whereas the Reynolds number had little effect on the behavior of either the basic or the slatted airfoil. The slat caused a slight drag penalty at low angles of attack, but generally increased the lift/drag ratio when averaged over the full cycle of oscillation.

  10. Thermostructural Evaluation of Joggle Region on the Shuttle Orbiter's Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Walker, Sandra P.; Warren, Jerry E.

    2012-01-01

    An investigation was initiated to determine the cause of coating spallation occurring on the Shuttle Orbiter's wing leading edge panels in the slip-side joggle region. The coating spallation events were observed, post flight, on differing panels on different missions. As part of the investigation, the high re-entry heating occurring on the joggles was considered here as a possible cause. Thus, a thermostructural evaluation was conducted to determine the detailed state-of-stress in the joggle region during re-entry and the feasibility of a laboratory test on a local joggle specimen to replicate this state-of-stress. A detailed three-dimensional finite element model of a panel slip-side joggle region was developed. Parametric and sensitivity studies revealed significant stresses occur in the joggle during peak heating. A critical interlaminar normal stress concentration was predicted in the substrate at the coating interface and was confined to the curved joggle region. Specifically, the high interlaminar normal stress is identified to be the cause for the occurrence of failure in the form of local subsurface material separation occurring in the slip-side joggle. The predicted critical stresses are coincident with material separations that had been observed with microscopy in joggle specimens obtained from flight panels.

  11. 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.

  12. Effect of Impact Location on the Response of Shuttle Wing Leading Edge Panel 9

    NASA Technical Reports Server (NTRS)

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

    2005-01-01

    The objective of this paper is to compare the results of several simulations performed to determine the worst-case location for a foam impact on the Space Shuttle wing leading edge. The simulations were performed using the commercial non-linear transient dynamic finite element code, LS-DYNA. These simulations represent the first in a series of parametric studies performed to support the selection of the worst-case impact scenario. Panel 9 was selected for this study to enable comparisons with previous simulations performed during the Columbia Accident Investigation. The projectile for this study is a 5.5-in cube of typical external tank foam weighing 0.23 lb. Seven locations spanning the panel surface were impacted with the foam cube. For each of these cases, the foam was traveling at 1000 ft/s directly aft, along the orbiter X-axis. Results compared from the parametric studies included strains, contact forces, and material energies for various simulations. The results show that the worst case impact location was on the top surface, near the apex.

  13. Wing Leading Edge RCC Rapid Response Damage Prediction Tool (IMPACT2)

    NASA Technical Reports Server (NTRS)

    Clark, Robert; Cottter, Paul; Michalopoulos, Constantine

    2013-01-01

    This rapid response computer program predicts Orbiter Wing Leading Edge (WLE) damage caused by ice or foam impact during a Space Shuttle launch (Program "IMPACT2"). The program was developed after the Columbia accident in order to assess quickly WLE damage due to ice, foam, or metal impact (if any) during a Shuttle launch. IMPACT2 simulates an impact event in a few minutes for foam impactors, and in seconds for ice and metal impactors. The damage criterion is derived from results obtained from one sophisticated commercial program, which requires hours to carry out simulations of the same impact events. The program was designed to run much faster than the commercial program with prediction of projectile threshold velocities within 10 to 15% of commercial-program values. The mathematical model involves coupling of Orbiter wing normal modes of vibration to nonlinear or linear springmass models. IMPACT2 solves nonlinear or linear impact problems using classical normal modes of vibration of a target, and nonlinear/ linear time-domain equations for the projectile. Impact loads and stresses developed in the target are computed as functions of time. This model is novel because of its speed of execution. A typical model of foam, or other projectile characterized by material nonlinearities, impacting an RCC panel is executed in minutes instead of hours needed by the commercial programs. Target damage due to impact can be assessed quickly, provided that target vibration modes and allowable stress are known.

  14. Numerical Predictions of Sonic Boom Signatures for a Straight Line Segmented Leading Edge Model

    NASA Technical Reports Server (NTRS)

    Elmiligui, Alaa A.; Wilcox, Floyd J.; Cliff, Susan; Thomas, Scott

    2012-01-01

    A sonic boom wind tunnel test was conducted on a straight-line segmented leading edge (SLSLE) model in the NASA Langley 4- by 4- Foot Unitary Plan Wind Tunnel (UPWT). The purpose of the test was to determine whether accurate sonic boom measurements could be obtained while continuously moving the SLSLE model past a conical pressure probe. Sonic boom signatures were also obtained using the conventional move-pause data acquisition method for comparison. The continuous data acquisition approach allows for accurate signatures approximately 15 times faster than a move-pause technique. These successful results provide an incentive for future testing with greatly increased efficiency using the continuous model translation technique with the single probe to measure sonic boom signatures. Two widely used NASA codes, USM3D (Navier-Stokes) and CART3D-AERO (Euler, adjoint-based adaptive mesh), were used to compute off-body sonic boom pressure signatures of the SLSLE model at several different altitudes below the model at Mach 2.0. The computed pressure signatures compared well with wind tunnel data. The effect of the different altitude for signature extraction was evaluated by extrapolating the near field signatures to the ground and comparing pressure signatures and sonic boom loudness levels.

  15. 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.

  16. 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.

  17. Dkk-1 Inhibits Intestinal Epithelial Cell Migration by Attenuating Directional Polarization of Leading Edge Cells

    PubMed Central

    Koch, Stefan; Capaldo, Christopher T.; Samarin, Stanislav; Nava, Porfirio; Neumaier, Irmgard; Skerra, Arne; Sacks, David B.; Parkos, Charles A.

    2009-01-01

    Wnt signaling pathways regulate proliferation, motility, and survival in a variety of human cell types. Dickkopf-1 (Dkk-1) is a secreted Wnt antagonist that has been proposed to regulate tissue homeostasis in the intestine. In this report, we show that Dkk-1 is secreted by intestinal epithelial cells after wounding and that it inhibits cell migration by attenuating the directional orientation of migrating epithelial cells. Dkk-1 exposure induced mislocalized activation of Cdc42 in migrating cells, which coincided with a displacement of the polarity protein Par6 from the leading edge. Consequently, the relocation of the microtubule organizing center and the Golgi apparatus in the direction of migration was significantly and persistently inhibited in the presence of Dkk-1. Small interfering RNA-induced down-regulation of Dkk-1 confirmed that extracellular exposure to Dkk-1 was required for this effect. Together, these data demonstrate a novel role of Dkk-1 in the regulation of directional polarization of migrating intestinal epithelial cells, which contributes to the effect of Dkk-1 on wound closure in vivo. PMID:19776352

  18. GMF promotes leading edge dynamics and collective cell migration in vivo

    PubMed Central

    Poukkula, Minna; Hakala, Markku; Pentinmikko, Nalle; Sweeney, Meredith O.; Jansen, Silvia; Mattila, Jaakko; Hietakangas, Ville; Goode, Bruce L.; Lappalainen, Pekka

    2014-01-01

    SUMMARY Lamellipodia are dynamic actin-rich cellular extensions, which drive advancement of the leading edge during cell migration [1–3]. Lamellipodia undergo periodic extension/retraction cycles [4–8], but the molecular mechanisms underlying these dynamics and their role in cell migration have remained obscure. We show that gliamaturation factor (GMF), which is an Arp2/3 complex inhibitor and actin filament debranching factor [9, 10], regulates lamellipodial protrusion dynamics in living cells. In cultured S2R+ cells, GMF silencing resulted in an increase in the width of lamellipodial actin filament arrays. Importantly, live-imaging of mutant Drosophila egg chambers revealed that the dynamics of actin-rich protrusions in migrating border cells are diminished in the absence of GMF. Consequently, velocity of border cell clusters undergoing guided migration was reduced in GMF mutant flies. Furthermore, genetic studies demonstrated that GMF cooperates with the Drosophila homologue of Aip1 (flare) in promoting disassembly of Arp2/3-nucleated actin filament networks and driving border cell migration. These data suggest that GMF functions in vivo to promote the disassembly of Arp2/3-nucleated actin filament arrays, making an important contribution to cell migration within a three-dimensional tissue environment. PMID:25308079

  19. Mechanism of Water Droplet Breakup Near the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida, Garcia

    2012-01-01

    This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. The airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Droplet deformation is defined and studied against main parameters. The high speed imaging allowed observation of the actual mechanism of breakup and identification of the sequence of configurations from the initiation of the breakup to the disintegration of the droplet. Results and comparisons are presented for droplets of diameters in the range of 500 to 1800 microns, and airfoil velocities of 70 and 90 m/sec.

  20. Mechanism of Water Droplet Breakup near the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida Garcia

    2012-01-01

    This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de T cnica Aeroespacial (INTA) in Madrid, Spain. The airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Droplet deformation is defined and studied against main parameters. The high speed imaging allowed observation of the actual mechanism of breakup and identification of the sequence of configurations from the initiation of the breakup to the disintegration of the droplet. Results and comparisons are presented for droplets of diameters in the range of 500 to 1800 micrometers, and airfoil velocities of 70 and 90 meters/second.

  1. 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.

  2. Tropomyosin Promotes Lamellipodial Persistence by Collaborating with Arp2/3 at the Leading Edge.

    PubMed

    Brayford, Simon; Bryce, Nicole S; Schevzov, Galina; Haynes, Elizabeth M; Bear, James E; Hardeman, Edna C; Gunning, Peter W

    2016-05-23

    At the leading edge of migrating cells, protrusion of the lamellipodium is driven by Arp2/3-mediated polymerization of actin filaments [1]. This dense, branched actin network is promoted and stabilized by cortactin [2, 3]. In order to drive filament turnover, Arp2/3 networks are remodeled by proteins such as GMF, which blocks the actin-Arp2/3 interaction [4, 5], and coronin 1B, which acts by directing SSH1L to the lamellipodium where it activates the actin-severing protein cofilin [6, 7]. It has been shown in vitro that cofilin-mediated severing of Arp2/3 actin networks results in the generation of new pointed ends to which the actin-stabilizing protein tropomyosin (Tpm) can bind [8]. The presence of Tpm in lamellipodia, however, is disputed in the literature [9-19]. Here, we report that the Tpm isoforms 1.8/9 are enriched in the lamellipodium of fibroblasts as detected with a novel isoform-specific monoclonal antibody. RNAi-mediated silencing of Tpm1.8/9 led to an increase of Arp2/3 accumulation at the cell periphery and a decrease in the persistence of lamellipodia and cell motility, a phenotype consistent with cortactin- and coronin 1B-deficient cells [2, 7]. In the absence of coronin 1B or cofilin, Tpm1.8/9 protein levels are reduced while, conversely, inhibition of Arp2/3 with CK666 leads to an increase in Tpm1.8/9 protein. These findings establish a novel regulatory mechanism within the lamellipodium whereby Tpm collaborates with Arp2/3 to promote lamellipodial-based cell migration. PMID:27112294

  3. 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.

  4. 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.

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

    NASA Technical Reports Server (NTRS)

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

  6. 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.

  7. 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.

  8. 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.

  9. Prediction of unsteady aerodynamic loadings caused by leading edge and trailing edge control surface motions in subsonic compressible flow: Analysis and results

    NASA Technical Reports Server (NTRS)

    Rowe, W. S.; Redman, M. C.; Ehlers, F. E.; Sebastian, J. D.

    1975-01-01

    A theoretical analysis and computer program was developed for the prediction of unsteady lifting surface loadings caused by motions of leading edge and trailing edge control surfaces having sealed gaps. The final form of the downwash integral equation was formulated by isolating the singularities from the nonsingular terms and using a preferred solution process to remove and evaluate the downwash discontinuities in a systematic manner. Comparisons of theoretical and experimental pressure data are made for several control surface configurations. The comparisons indicate that reasonably accurate theoretical pressure distributions and generalized forces may be obtained for a wide variety of control surface configurations. Spanwise symmetry or antisymmetry of motion, and up to six control surfaces on each half span can be accommodated.

  10. 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.

  11. Data acquisition electronics for gamma ray emission tomography using width-modulated leading-edge discriminators.

    PubMed

    Lage, E; Tapias, G; Villena, J; Desco, M; Vaquero, J J

    2010-08-01

    We present a new high-performance and low-cost approach for implementing radiation detection acquisition systems. The basic elements used are charge-integrating ADCs and a set of components encapsulated in an HDL (hardware definition language) library which makes it possible to implement several acquisition tasks such as time pickoff and coincidence detection using a new and simple trigger technique that we name WMLET (width-modulated leading-edge timing). As proof of concept, a 32-channel hybrid PET/SPECT acquisition system based on these elements was developed and tested. This demonstrator consists of a master module responsible for the generation and distribution of trigger signals, 2 x 16-channel ADC cards (12-bit resolution) for data digitization and a 32-bit digital I/O PCI card for handling data transmission to a personal computer. System characteristics such as linearity, maximum transmission rates or timing resolution in coincidence mode were evaluated with test and real detector signals. Imaging capabilities of the prototype were also evaluated using different detector configurations. The performance tests showed that this implementation is able to handle data rates in excess of 600k events s(-1) when acquiring simultaneously 32 channels (96-byte events). ADC channel linearity is >98.5% in energy quantification. Time resolution in PET mode for the tested configurations ranges from 3.64 ns FWHM to 7.88 ns FWHM when signals from LYSO-based detectors are used. The measured energy resolution matched the expected values for the detectors evaluated and single elements of crystal matrices can be neatly separated in the acquired flood histograms. PMID:20647602

  12. Data acquisition electronics for gamma ray emission tomography using width-modulated leading-edge discriminators

    NASA Astrophysics Data System (ADS)

    Lage, E.; Tapias, G.; Villena, J.; Desco, M.; Vaquero, J. J.

    2010-08-01

    We present a new high-performance and low-cost approach for implementing radiation detection acquisition systems. The basic elements used are charge-integrating ADCs and a set of components encapsulated in an HDL (hardware definition language) library which makes it possible to implement several acquisition tasks such as time pickoff and coincidence detection using a new and simple trigger technique that we name WMLET (width-modulated leading-edge timing). As proof of concept, a 32-channel hybrid PET/SPECT acquisition system based on these elements was developed and tested. This demonstrator consists of a master module responsible for the generation and distribution of trigger signals, 2 × 16-channel ADC cards (12-bit resolution) for data digitization and a 32-bit digital I/O PCI card for handling data transmission to a personal computer. System characteristics such as linearity, maximum transmission rates or timing resolution in coincidence mode were evaluated with test and real detector signals. Imaging capabilities of the prototype were also evaluated using different detector configurations. The performance tests showed that this implementation is able to handle data rates in excess of 600k events s-1 when acquiring simultaneously 32 channels (96-byte events). ADC channel linearity is >98.5% in energy quantification. Time resolution in PET mode for the tested configurations ranges from 3.64 ns FWHM to 7.88 ns FWHM when signals from LYSO-based detectors are used. The measured energy resolution matched the expected values for the detectors evaluated and single elements of crystal matrices can be neatly separated in the acquired flood histograms.

  13. A numerical and experimental study of the effects of dynamic roughness on laminar leading edge separation

    NASA Astrophysics Data System (ADS)

    Gall, Peter D.

    The aircraft industry, as a whole, has been deeply concerned with improving the aerodynamic efficiency of current and future flight vehicles, particularly in the commercial and military markets. However, of particular interest to the field of aerodynamics is the elusive concept of a workable flow control mechanism. Effective flow control is a concept which if properly applied can increase aerodynamic efficiency. Various concepts and ideas to obtain successful flow control have been studied in an attempt to reap these rewards. Some examples include boundary layer blowing (steady and periodic), suction, and compliant walls for laminar flow control. The overall goal of flow control is to increase performance by increasing lift, reducing drag, and delaying or eliminating leading edge separation. The specific objectives of flow control are to (1) delay or eliminate flow separation, (2) delay boundary layer transition, and (3) and reduce skin friction drag. The purpose of this research is to investigate dynamic roughness as a novel method of flow control technology for external boundary layer flows. As opposed to standard surface roughness, dynamic roughness incorporates small time dependent perturbations to the surface of the airfoil. These surface perturbations are actual humps and/or ridges on the surface of the airfoil that are on the scale of the laminar boundary, and oscillate with an unsteady motion. Research has shown that this can provide a means to modify the instantaneous and mean velocity profile near the wall and favorably control the existing state of the boundary layer. Several flow control parameters were studied including dynamic roughness frequency, amplitude, and geometry. The results of this study have shown, both numerically and experimentally, that dynamic roughness can provide an effective means for eliminating both a short and long laminar separation bubble and possibly prove a viable alternative in effective flow control, hence reaping some of

  14. Space environmental effects on LDEF composites: A leading edge coated graphite epoxy panel

    NASA Technical Reports Server (NTRS)

    George, Pete E.; Dursch, Harry W.; Hill, Sylvester G.

    1993-01-01

    The electronics module cover for the leading edge (Row D 9) experiment M0003-8 was fabricated from T300 graphite/934 epoxy unidirectional prepreg tape in a (O(sub 2), +/- 45, O(sub 2), +/- 45, 90, 0)(sub s) layup. This 11.75 in x 16.75 in panel was covered with thermal control coatings in three of the four quadrants with the fourth quadrant uncoated. The composite panel experienced different thermal cycling extremes in each quadrant due to the different optical properties of the coatings and bare composite. The panel also experienced ultraviolet (UV) and atomic oxygen (AO) attack as well as micrometeoroid and space debris impacts. An AO reactivity of 0.99 x 10(exp -24) cm(sup 3)/atom was calculated for the bare composite based on thickness loss. The white urethane thermal control coatings (A276 and BMS 1060) prevented AO attack of the composite substrate. However, the black urethane thermal control coating (Z306) was severely eroded by AO, allowing some AO attack of the composite substrate. An interesting banding pattern on the AO eroded bare composite surface was investigated and found to match the dimensions of the graphite fiber tow widths as prepregged. Also, erosion depths were greater in the darker bands. Five micrometeoroid/space debris impacts were cross sectioned to investigate possible structural damage as well as impact/AO interactions. Local crushing and delaminations were found to some extent in all of the impacts. No signs of coating undercutting were observed despite the extensive AO erosion patterns seen in the exposed composite material at the impact sites. An extensive microcrack study was performed on the panel along with modeling of the thermal environment to estimate temperature extremes and thermal shock. The white coated composite substrate displayed almost no microcracking while the black coated and bare composite showed extensive microcracking. Significant AO erosion was seen in many of the cracks in the bare composite.

  15. 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.

  16. Design fabrication, testing, and delivery of shuttle heat pipe leading edge test modules. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The construction of two test modules is presented for a feasibility demonstration of a concept for reusable space shuttle wing leading edge surfaces. In this leading edge concept high temperature heat pipes were incorporated into the structure to cool the stagnation region, allowing the use of super-alloys in place of refractory metal, ablator protected, or carbon-carbon structures. The program included the analysis and design of the heat pipes, their integration into the test module structure, heat pipe development testing, construction of the test modules and a facility adapter, and formulation of recommended testing conditions. The results of the heat pipe and leading edge module thermal analyses indicate the test modules will meet the design goal; reducing the leading edge temperature at the stagnation line from 1315 C (2400 F) to less than 1010 C (1850 F). The development tests demonstrated that the module assembly could be brazed with active heat pipes, as was borne out by the subsequent successful brazing of both modules with active heat pipes loaded with sodium.

  17. "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 a series…

  18. 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.

  19. Turbulent Vortex-Flow Simulation Over a 65 deg Sharp and Blunt Leading-Edge Delta Wing at Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Ghaffari, Farhad

    2005-01-01

    Turbulent thin-layer, Reynolds-Averaged Navier-Stokes solutions, based on a multi-block structured grid, are presented for a 65 deg delta wing having either a sharp leading edge (SLE) or blunt leading edge (BLE) geometry. The primary objective of the study is to assess the prediction capability of the method for simulating the leading-edge flow separation and the ensuing vortex flow characteristics. Computational results are obtained for two angles of attack of approximately 13 and 20 deg, at free-stream Mach number of 0.40 and Reynolds number of 6 million based on the wing mean aerodynamic chord. The effects of two turbulence models of Baldwin-Lomax with Degani-Schiff (BL/DS) and the Spalart-Allmaras (SA) on the numerical results are also discussed. The computations also explore the effects of two numerical flux-splitting schemes, i.e., flux difference splitting (fds) and flux vector splitting (fvs), on the solution development and convergence characteristics. The resulting trends in solution sensitivity to grid resolution for the selected leading-edge geometries, angles of attack, turbulence models and flux splitting schemes are also presented. The validity of the numerical results is evaluated against a unique set of experimental wind-tunnel data that was obtained in the National Transonic Facility at the NASA Langley Research Center.

  20. 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.

  1. A general aerodynamic approach to the problem of decaying or growing vibrations of thin, flexible wings with supersonic leading and trailing edges and no side edges

    NASA Technical Reports Server (NTRS)

    Warner, R. W.

    1975-01-01

    Indicial aerodynamic influence coefficients were evaluated from potential theory for a thin, flexible wing with supersonic leading and trailing edges only. The analysis is based on the use of small surface areas in which the downwash is assumed uniform. Within this limitation, the results are exact except for the restriction of linearized theory. The areas are not restricted either to square boxes or Mach boxes. A given area may be any rectangle or square which may or may not be cut by the Mach forecone, and any area can be used anywhere in the forecone without loss of accuracy.

  2. 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.

  3. 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.

  4. Stochastic Model of Integrin-Mediated Signaling and Adhesion Dynamics at the Leading Edges of Migrating Cells

    PubMed Central

    Cirit, Murat; Krajcovic, Matej; Choi, Colin K.; Welf, Erik S.; Horwitz, Alan F.; Haugh, Jason M.

    2010-01-01

    Productive cell migration requires the spatiotemporal coordination of cell adhesion, membrane protrusion, and actomyosin-mediated contraction. Integrins, engaged by the extracellular matrix (ECM), nucleate the formation of adhesive contacts at the cell's leading edge(s), and maturation of nascent adhesions to form stable focal adhesions constitutes a functional switch between protrusive and contractile activities. To shed additional light on the coupling between integrin-mediated adhesion and membrane protrusion, we have formulated a quantitative model of leading edge dynamics combining mechanistic and phenomenological elements and studied its features through classical bifurcation analysis and stochastic simulation. The model describes in mathematical terms the feedback loops driving, on the one hand, Rac-mediated membrane protrusion and rapid turnover of nascent adhesions, and on the other, myosin-dependent maturation of adhesions that inhibit protrusion at high ECM density. Our results show that the qualitative behavior of the model is most sensitive to parameters characterizing the influence of stable adhesions and myosin. The major predictions of the model, which we subsequently confirmed, are that persistent leading edge protrusion is optimal at an intermediate ECM density, whereas depletion of myosin IIA relieves the repression of protrusion at higher ECM density. PMID:20195494

  5. Subscale, hydrogen-burning, airframe-integrated-scramjet: Experimental and theoretical evaluation of a water cooled strut airframe-integrated-scramjet: Experimental leading edge

    NASA Technical Reports Server (NTRS)

    Pinckney, S. Z.; Guy, R. W.; Beach, H. L., Jr.; Rogers, R. C.

    1975-01-01

    A water-cooled leading-edge design for an engine/airframe integrated scramjet model strut leading edge was evaluated. The cooling design employs a copper cooling tube brazed just downstream of the leading edge of a wedge-shaped strut which is constructed of oxygen-free copper. The survival of the strut leading edge during a series of tests at stagnation point heating rates confirms the practicality of the cooling design. A finite difference thermal model of the strut was also proven valid by the reasonable agreement of calculated and measured values of surface temperature and cooling-water heat transfer.

  6. 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.

  7. 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.

  8. An experimental investigation of the shear-layer and acoustic sources produced by a leading edge slat

    NASA Astrophysics Data System (ADS)

    Wilkins, Stephen; Richard, Patrick; Hall, Joseph; Turbulence; Flow Noise Laboratory Team

    2013-11-01

    Leading edge slats are a common addition to airfoils as part of a high lift configuration employed during take-off and landing; the unsteady flow caused by these slats is a major contributor to the overal airframe noise. As the next generation of aircraft seeks to reduce these noise concerns, a better understanding of the sources of aeroacoustic noise generation is sought. Particle Image Velocimetry (PIV) and simultaneous multipoint measurements of the unsteady surface pressure are used herein to investigate the unsteady flow around a leading edge slat coupled with an airfoil for several different configurations and a range of Reynolds numbers (Re = 156 , 000 to Re = 1 . 2 million based on the wing chord). Shear-layer development off the slat cusp and the related unsteady vortex structures are examined in detail to better establish and understand the mechanisms responsible for the generation of aeroacoustic slat noise. The authors are grateful for the support provided by GARDN.

  9. Calculation of the 3-D viscous flow at the endwall leading edge region of an axial annular turbine cascade

    NASA Technical Reports Server (NTRS)

    Walitt, L.

    1984-01-01

    A three-dimensional viscous computer code (VANS/MD) was employed to calculate the turbulent flow field at the end wall leading edge region of a 20 inch axial annular turbine cascade. The initial boundary layer roll-up and formation of the end wall vortices were computed at the vane leading edge. The calculated flow field was found to be periodic with a frequency of approximately 1600 Hz. The calculated size of the separation region for the hub endwall vortex compared favorably with measured endwall oil traces. In an effort to determine the effects of the turbulence model on the calculated unsteadiness, a laminar calculation was made. The periodic nature of the calculated flow field persisted with the frequency essentially unchanged.

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

    SciTech Connect

    Roussel-Dupre, R.; Symbalisty, E.; Fox, C.; and Vanderlinde, O.

    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 the 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).

  11. 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.

  12. 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.

  13. Investigation of Porous Gas-Heated Leading-Edge Section for Icing Protection of a Delta Wing

    NASA Technical Reports Server (NTRS)

    Bowden, Dean T.

    1955-01-01

    A tip section of a delta wing having an NACA 0004-65 airfoil section and a 600 leading-edge sweepback was equipped with a porous leading-edge section through which hot gas was 'bled for anti-icing. Heating rates for anti-icing were determined for a wide range of icing conditions. The effects of gas flow through the porous leading-edge section on airfoil pressure distribution and drag in dry air were investigated. The drag increase caused by an ice formation on the unheated airfoil was measured for several icing conditions. Experimental porous surface- to free-stream convective heat-transfer coefficients were obtained in dry air and compared with theory. Adequate icing protection was obtained at all icing conditions investigated. Savings in total gas-flow rate up to 42 percent may be obtained with no loss in anti-icing effectiveness by sealing half the upper-surface porous area. Gas flow through the leading-edge section had no appreciable effect on airfoil pressure distribution. The airfoil section drag increased slightly (5-percent average) with gas flow through the porous surface. A heavy glaze-ice formation produced after 10 minutes of icing caused an increase in section drag coefficient of 240 percent. Experimental convective heat-transfer coefficients obtained with hot-gas flow through the porous area in dry air and turbulent flow were 20 to 30 percent lower than the theoretical values for a solid surface under similar conditions. The transition region from laminar to turbulent flow moved forward as the ratio of gas velocity through the porous surface to air-stream velocity was increased.

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

    NASA Astrophysics Data System (ADS)

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

    2015-05-01

    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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. Modeling Creep-Induced Stress Relaxation at the Leading Edge of SiC/SiC Airfoils

    NASA Technical Reports Server (NTRS)

    Lang, Jerry; DiCarlo, James A.

    2007-01-01

    Anticipating the implementation of advanced SiC/SiC composites into internally cooled airfoil components within the turbine section of future aero-propulsion engines, the primary objective of this study was to develop physics-based analytical and finite-element modeling tools to predict the effects of composite creep and stress relaxation at the airfoil leading edges, which will generally experience large thermal gradients at high temperatures. A second objective was to examine how some advanced NASA-developed SiC/SiC systems coated with typical EBC materials would behave as leading edge materials in terms of long-term steady-state operating temperatures. Because of the complexities introduced by mechanical stresses inherent in internally cooled airfoils, a simple cylindrical thin-walled tube model subjected to thermal stresses only is employed for the leading edge, thereby obtaining a best-case scenario for the material behavior. In addition, the SiC/SiC composite materials are assumed to behave as isotropic materials with temperature-dependent viscoelastic creep behavior as measured in-plane on thin-walled panels. Key findings include: (1) without mechanical stresses and for typical airfoil geometries, as heat flux is increased through the leading edge, life-limiting tensile crack formation will occur first in the hoop direction on the inside wall of the leading edge; (2) thermal gradients through all current SiC/SiC systems should be kept below approx.300 F at high temperatures to avoid this cracking; (3) at temperatures near the maximum operating temperatures of advanced SiC/SiC systems, thermal stresses induced by the thermal gradients will beneficially relax with time due to creep; (4) although stress relaxation occurs, the maximum gradient should still not exceed 300oF because of residual tensile stress buildup on the airfoil outer wall during cool-down; and (5) without film cooling and mechanical stresses, the NASA-developed N26 SiC/SiC system with thru

  20. 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.

  1. The definition of the Shuttle Tethered Aerothermodynamic Research Facility

    NASA Technical Reports Server (NTRS)

    Siemers, P. M., III; Wood, G. M., Jr.; Wolf, H.; Flanagan, P. F.; Henry, M. W.

    1985-01-01

    Studies have been conducted to define the feasibility and practical limitations of the Shuttle Orbiter Tethered 'wind-tunnel' concept. This concept, referred to as the Shuttle Tethered Aerothermodynamic Research Facility (STARFAC), is proposed to provide researchers access to altitudes above 90 km to accomplish aerothermodynamic research in the rarefied upper atmosphere. Determining the feasibility and limitations of the concept has required the enhancement and/or development of mission simulation analytical techniques and control laws; the accomplishment of candidate mission simulations; the definition of instrumentation requirements, both for science and engineering; and the establishment of tether and satellite design requirements to meet STARFAC objectives. The results of the study, to date, indicate that such a concept is both feasible and practical. Representative results are presented, as are recommendations for continued studies which would result in program implementation.

  2. The tethered satellite system for low density aerothermodynamics studies

    NASA Technical Reports Server (NTRS)

    Carlomagno, Giovanni M.; De Luca, Luigi; Siemers, P. M., III; Wood, George M., Jr.

    1986-01-01

    The feasibility of the operation of the Tethered Satellite System (TSS) as a continuous open wind tunnel for low-density aerothermodynamic studies (applicable to the design of hypersonic space vehicles including STARFAC, AOTV, and ERV) is considered. The Shuttle Continuous Open Wind Tunnel (SCOWT) program, for the study of the energy and momentum transfer between the tethered satellite and its environmental medium during the TSS/2 mission, is described. Instrumentation and TSS design requirements to meet SCOWT objectives are also considered. SCOWT will provide information on the gasdynamic processes occurring downstream of the bow wave standing in front of the TS, the chemistry and physics of the upper atmosphere related to satellite aerothermodynamics, and TSS's overall experimental envelope of operation.

  3. 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.

  4. Downward-deployed tethered platforms for high enthalpy aerothermodynamic research

    NASA Technical Reports Server (NTRS)

    Wood, George M.; Siemers, Paul M.; Squires, R. Kenneth; Wolf, Henry; Carlomagno, Giovanni M.

    1988-01-01

    The data on aerothermodynamic and aerodynamic interactions at altitudes above 50 km is extremely limited because of the relative inaccessibility of the region to research vehicles of any sort. This paper addresses the practicability of using downward deployed satellites tethered to an orbiting host vehicle in order to obtain steady-state data in the upper reaches of the region above 80 or 90 km.

  5. Shuttle entry aerothermodynamic flight research - The Orbiter Experiments (OEX) Program

    NASA Technical Reports Server (NTRS)

    Throckmorton, David A.

    1992-01-01

    Results of the OEX program are summarized with emphasis on the information on entry aerothermodynamic phenomena derived from Space Shuttle operations. The discussion focuses on OEX experiment complement and operational history, freestream environment and vehicle attitude data, aerodynamic force and moment data, aerodynamic surface data, and vehicle configuration data. Attention is also given to orbiter aerodynamic performance, stability and control, high-altitude atmospheric density variability, direct simulation Monte Carlo validation, orbital drag variation, and computational fluid dynamic technique validation.

  6. 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.

  7. An Euler aerodynamic method for leading-edge vortex flow simulation

    NASA Technical Reports Server (NTRS)

    Raj, P.; Long, L. N.

    1986-01-01

    The current capabilities and the future plans for a three dimensional Euler Aerodynamic Method are described. The basic solution algorithm is based on the finite volume, Runge-Kutta pseudo-time-stepping scheme of FLO-57. Several modifications to improve accuracy and computational efficiency were incorporated and others are being investigated. The computer code is used to analyze a cropped delta wing at 0.6 Mach number and an arrow wing at 0.85 Mach number. Computed aerodynamic parameters are compared with experimental data. In all cases, the configuration is impulsively started and no Kutta condition is applied at sharp edges. The results indicate that with additional development and validation, the present method will be a useful tool for engineering analysis of high speed aircraft.

  8. Characterization of Invading Glioma Cells Using Molecular Analysis of Leading-Edge Tissue

    PubMed Central

    Kim, Cheol-Soo; Jung, Tae-Young; Jang, Woo-Youl; Sun, Heung-Suk; Ryu, Hyang-Hwa

    2011-01-01

    Objective We have introduced a method of characterization of invading glioma cells by using molecular analysis of marginal invading tumor cells and molecular profiles of glioma tumor margin. Methods Each of tumor core and marginal tissues was obtained in 22 glioma patients. Tumor core cells and marginal cells from each glial tumor were collected by laser capture microdissection or intraoperative microdissection under the operating microscope. Expression of MMP-2, MMP-9, CD44 and RHAMM mRNA by invading glioma cells compared with tumor core was confirmed by realtime-PCR of twenty-four glioma specimens. Clinical data also were reviewed for invasion and recurrence pattern of the gliomas radiologically and invasive rim pattern microscopically. Results Overall results of the molecular analysis showed that relative overexpression of MMP-2, MMP-9 and RHAMM were noted at the invasive edge of human glioma specimens comparing to the tumor core but CD44 was highly expressed in the tumor core comparing to the margin. High marginal expression of MMP-2 and MMP-9 were noted in poorly ill-defined margin on the pathological finding. High marginal expression of CD44 and MMP-2 were demonstrated in the midline cross group on the radiological review, and that of RHAMM and MMP-2 were showed in the aggressive recurrence group. High expression of MMP-2 seems to be involved in the various invasion-related phenomenons. Conclusion Up-regulation of MMP-2, MMP-9, CD44 and RHAMM was noted in invasive edge of gliomas according to the various clinical situations. PMID:22102942

  9. 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.

  10. Aerothermal investigations of mixing flow phenomena in case of radially inclined ejection holes at the leading edge

    SciTech Connect

    Bohn, D.E.; Kusterer, K.A.

    2000-04-01

    A leading edge cooling configuration is investigated numerically by application of a three-dimensional conjugate fluid flow and heat transfer solver, CHT-flow. The code has been developed at the Institute of Steam and Gas Turbines, Aachen University of Technology. It works on the basis of an implicit finite volume method combined with a multi-block technique. The cooling configuration is an axial turbine blade cascade with leading edge ejection through two rows of cooling holes. The rows are located in the vicinity of the stagnation line, one row on the suction side, the other row is on the pressure side. the cooling holes have a radial ejection angle of 45 degrees. This configuration has been investigated experimentally by other authors and the results have been documented as a test case for numerical calculations of ejection flow phenomena. The numerical investigations focus on the aerothermal mixing process in the cooling jets and the impact on the temperature distribution on the blade surface. The radial ejection angles lead to a fully three-dimensional and asymmetric jet flow field. Within a secondary flow analysis, the cooling fluid jets are investigated in detail. The secondary flow fields include asymmetric kidney vortex systems with one dominating vortex on the back side of the jets. The numerical and experimental data show a respectable agreement concerning the vortex development.

  11. 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.

  12. Effects of Mach Number, Leading-Edge Bluntness, and Sweep on Boundary-Layer Transition on a Flat Plate

    NASA Technical Reports Server (NTRS)

    Jillie, Don W.; Hopkins, Edward J.

    1961-01-01

    The effects of leading-edge bluntness and sweep on boundary-layer transition on flat plate models were investigated at Mach numbers of 2.00, 2.50, 3.00, and 4.00. The effect of sweep on transition was also determined on a flat plate model equipped with an elliptical nose at a Mach number of 0.27. Models used for the supersonic investigation had leading-edge radii varying from 0.0005 to 0.040 inch. The free-stream unit Reynolds number was held constant at 15 million per foot for the supersonic tests and the angle of attack was 0 deg. Surface flow conditions were determined by visual observation and recorded photographically. The sublimation technique was used to indicate transition, and the fluorescent-oil technique was used to indicate flow separation. Measured Mach number and sweep effects on transition are compared with those predicted from shock-loss considerations as described in NACA Rep. 1312. For the models with the blunter leading edges, the transition Reynolds number (based on free-stream flow conditions) was approximately doubled by an increase in Mach number from 2.50 to 4.00; and nearly the same result was predicted from shock-loss considerations. At all super- sonic Mach numbers, increases in sweep reduced the transition Reynolds number and the amount of reduction increased with increases in bluntness. The shock-loss method considerably underestimated- the sweep effects, possibly because of the existence of crossflow instability associated with swept wings. At a Mach number of 0.27, no reduction in the transition Reynolds number with sweep was measured (as would be expected with no shock loss) until the sweep angle was attained where crossflow instability appeared.

  13. 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.

  14. 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.

  15. Cloud particle effects on laminar flow in the NASA LEFT program - Preliminary results. [Leading Edge Flight Test

    NASA Technical Reports Server (NTRS)

    Davis, R. E.; Fischer, M. C.; Fisher, D. F.; Young, R.

    1986-01-01

    Laminar flow offers the promise of significant fuel savings on future commercial transport aircraft, but laminar flow can be lost while encountering clouds or haze at cruise conditions. To quantify the effect of cloud particles on laminar flow during typical airline operating conditions, and evaluate candidate cloud particle detection instrument concepts for future laminar flow aircraft, two types of cloud particle detectors are being flown aboard a NASA JetStar aircraft in the Leading Edge Flight Test (LEFT) program. The instrumentation is described, and preliminary results and conclusions are presented.

  16. 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.

  17. Hypersonic High-Enthalpy Flow in a Leading-Edge Separation

    NASA Astrophysics Data System (ADS)

    Deepak, N. R.; Gai, S. L.; O'Byrne, Sean; Moss, J. N.

    Flow separation in/over a hypersonic space vehicle is an important phenomenon which occurs due to flow interaction with various geometric elements of the vehicle. This however can lead to adverse pressure gradient and localised intense heating resulting in detrimental consequences for the successful performance of the vehicle.

  18. 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.

  19. Leading Edge Flame Detachment: Effect on Burning Rate of Ammonium Perchlorate Propellants

    NASA Technical Reports Server (NTRS)

    Price, E. W.; Chakravarthy, S. R.; Sambamurthi, J. K.; Sigman, R. K.

    1997-01-01

    Inner details of the transition from premixed to diffusion controlled burning are examined by considering the transition at the individual particle flamelet level. These considerations lead to prediction of observable singular burning rate behavior of bimodal AP formulations. Burning rate measurements verify the predictions, supporting the view that particle flamelets become detached and retreat to more remote premixed locations at definite (particles size dependent) pressures, revealed by the burning rate tests.

  20. Young Children of Immigrants: The Leading Edge of America's Future. Brief No. 3

    ERIC Educational Resources Information Center

    Fortuny, Karina; Hernandez, Donald J.; Chaudry, Ajay

    2010-01-01

    Children of immigrants have nearly doubled as a share of pre-K to 3rd grade students since 1990. The share of children under age 8 with immigrant parents stood at 24 percent in 2008, up from 13 percent in 1990. Young children of immigrants account for more than 30 percent of children in seven states, with California leading the nation at 50…

  1. A numerical investigation into the effects of Reynolds number on the flow mechanism induced by a tubercled leading edge

    NASA Astrophysics Data System (ADS)

    Rostamzadeh, Nikan; Kelso, Richard M.; Dally, Bassam

    2016-05-01

    Leading-edge modifications based on designs inspired by the protrusions on the pectoral flippers of the humpback whale (tubercles) have been the subject of research for the past decade primarily due to their flow control potential in ameliorating stall characteristics. Previous studies have demonstrated that, in the transitional flow regime, full-span wings with tubercled leading edges outperform unmodified wings at high attack angles. The flow mechanism associated with such enhanced loading traits is, however, still being investigated. Also, the performance of full-span tubercled wings in the turbulent regime is largely unexplored. The present study aims to investigate Reynolds number effects on the flow mechanism induced by a full-span tubercled wing with the NACA-0021 cross-sectional profile in the transitional and near-turbulent regimes using computational fluid dynamics. The analysis of the flow field suggests that, with the exception of a few different flow features, the same underlying flow mechanism, involving the presence of transverse and streamwise vorticity, is at play in both cases. With regard to lift-generation characteristics, the numerical simulation results indicate that in contrast to the transitional flow regime, where the unmodified NACA-0021 undergoes a sudden loss of lift, in the turbulent regime, the baseline foil experiences gradual stall and produces more lift than the tubercled foil. This observation highlights the importance of considerations regarding the Reynolds number effects and the stall characteristics of the baseline foil, in the industrial applications of tubercled lifting bodies.

  2. 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.

  3. Design and longitudinal dynamic stability analysis of a slender delta kite for high altitudes using leading edge suction analogy

    NASA Astrophysics Data System (ADS)

    Madduri, Bharath

    In this thesis, the longitudinal dynamic stability modes, namely Phugoid and Short-period of delta kite with single tether are examined, for different aspect ratios (A) and flow conditions. The equations of motion, of kite are solved in polar-inertial wind frame and the tether is approximated by straight line elements. The vortex lift and induced drag due to leading edge vortices are calculated using Polhamus leading edge suction analogy. The Polhamus proportionality constants (Kp, Kv) are used to estimate the overall coefficient of lift and drag (C L, CD) and are computed using Multhopp lifting surface theory. The values of total coefficient of lift and drag (CL, CD) are examined for a wide variety of aspect ratio of delta kite and are validated by comparing with the experimental data. Linear stability analysis is performed for the chosen design variables to ensure the nominal design has stable longitudinal dynamics. A plot of the root locus of the system matrix for longitudinal dynamics as a function of geometry and flight conditions, provided an intuitive understanding of the flight modes of the kite, with respect to design parameters of interest.

  4. Improving Self-Guiding of an Ultra-Intense Laser by Tailoring the Longitudinal Profile of the Leading Edge

    NASA Astrophysics Data System (ADS)

    Mori, Warren; Tzoufras, Michail; Tsung, Frank; Sahai, Aakash

    2013-10-01

    Self-guiding of an ultra-intense laser pulse requires the refractive index to build up rapidly to a sufficient value before the main body of the pulse passes by. For short single-frequency pulses this occurs within a plasma period and a large portion of the leading edge is subject to diffraction. Nevertheless, if the body of the pulse survives long enough, the concomitant changes in its spectral content result in highly localized absorption, such that a large amount of the energy of the leading edge of the pulse is absorbed before it can diffract. To illustrate these mechanisms and optimize laser wakefield accelerators we propose a pulse profile with a ``bulbous bow,'' that is a lower-intensity low-energy precursor, that can produce the necessary buildup for the index of refraction to guide the body of the laser. The wake-field behind such a pulse is more stable, contains more energy, is sustained longer, and the corresponding de-phasing length is extended.

  5. DENND2B activates Rab13 at the leading edge of migrating cells and promotes metastatic behavior

    PubMed Central

    Ioannou, Maria S.; Bell, Emily S.; Girard, Martine; Chaineau, Mathilde; Hamlin, Jason N.R.; Daubaras, Mark; Monast, Anie; Park, Morag; Hodgson, Louis

    2015-01-01

    The small guanosine triphosphatase Rab13 functions in exocytic vesicle trafficking in epithelial cells. Alterations in Rab13 activity have been observed in human cancers, yet the mechanism of Rab13 activation and its role in cancer progression remain unclear. In this paper, we identify the DENN domain protein DENND2B as the guanine nucleotide exchange factor for Rab13 and develop a novel Förster resonance energy transfer–based Rab biosensor to reveal activation of Rab13 by DENND2B at the leading edge of migrating cells. DENND2B interacts with the Rab13 effector MICAL-L2 at the cell periphery, and this interaction is required for the dynamic remodeling of the cell’s leading edge. Disruption of Rab13-mediated trafficking dramatically limits the invasive behavior of epithelial cells in vitro and the growth and migration of highly invasive cancer cells in vivo. Thus, blocking Rab13 activation by DENND2B may provide a novel target to limit the spread of epithelial cancers. PMID:25713415

  6. 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.

  7. Application of an aerodynamic analysis method including attainable thrust estimates to low speed leading-edge flap design for supersonic cruise vehicles

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.

    1982-01-01

    A study of low speed leading-edge flap design for supersonic cruise vehicle was conducted. Wings with flaps were analyzed with the aid of a newly developed subsonic wing program which provides estimates of attainable leading-edge thrust. Results indicate that the thrust actually attainable can have a significant influence on the design and that the resultant flaps can be smaller and simpler than those resulting from more conventional approaches.

  8. 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.

  9. Predictive process simulation of cryogenic implants for leading edge transistor design

    SciTech Connect

    Gossmann, Hans-Joachim; Zographos, Nikolas; Park, Hugh; Colombeau, Benjamin; Parrill, Thomas; Khasgiwale, Niranjan; Borges, Ricardo; Gull, Ronald; Erokhin, Yuri

    2012-11-06

    Two cryogenic implant TCAD-modules have been developed: (i) A continuum-based compact model targeted towards a TCAD production environment calibrated against an extensive data-set for all common dopants. Ion-specific calibration parameters related to damage generation and dynamic annealing were used and resulted in excellent fits to the calibration data-set. (ii) A Kinetic Monte Carlo (kMC) model including the full time dependence of ion-exposure that a particular spot on the wafer experiences, as well as the resulting temperature vs. time profile of this spot. It was calibrated by adjusting damage generation and dynamic annealing parameters. The kMC simulations clearly demonstrate the importance of the time-structure of the beam for the amorphization process: Assuming an average dose-rate does not capture all of the physics and may lead to incorrect conclusions. The model enables optimization of the amorphization process through tool parameters such as scan speed or beam height.

  10. 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.

  11. The formation mechanism and impact of streamwise vortices on NACA 0021 airfoil's performance with undulating leading edge modification

    NASA Astrophysics Data System (ADS)

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

    2014-10-01

    Wings with tubercles have been shown to display advantageous loading behavior at high attack angles compared to their unmodified counterparts. In an earlier study by the authors, it was shown that an undulating leading-edge configuration, including but not limited to a tubercled model, induces a cyclic variation in circulation along the span that gives rise to the formation of counter-rotating streamwise vortices. While the aerodynamic benefits of full-span tubercled wings have been associated with the presence of such vortices, their formation mechanism and influence on wing performance are still in question. In the present work, experimental and numerical tests were conducted to further investigate the effect of tubercles on the flow structure over full-span modified wings based on the NACA 0021 profile, in the transitional flow regime. It is found that a skew-induced mechanism accounts for the formation of streamwise vortices whose development is accompanied by flow separation in delta-shaped regions near the trailing edge. The presence of vortices is detrimental to the performance of full-span wings pre-stall, however renders benefits post-stall as demonstrated by wind tunnel pressure measurement tests. Finally, primary and secondary vortices are identified post-stall that produce an enhanced momentum transfer effect that reduces flow separation, thus increasing the generated amount of lift.

  12. 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. PMID:27538989

  13. 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.

  14. ALES, the multi­mission Adaptive Leading Edge Sub­Waveform Retracker, design and validation

    NASA Astrophysics Data System (ADS)

    Passaro, Marcello; Benveniste, Jérôme; Vignudelli, Stefano; Cipollini, Paolo; Quartly, Graham; Snaith, Helen

    Satellite altimetry has revolutionized our understanding of ocean dynamics thanks to high repetition rate and global coverage. Nevertheless, coastal data has been flagged as unreliable due to land and calm water interference in the altimeter and radiometer footprint and high frequency tidal and atmospheric forcing. Our study addresses the first issue, i.e. retracking, presenting ALES, the Adaptive Leading Edge Subwaveform Retracker. ALES is potentially applicable to all the pulse­limited altimetry altimetry missions and its aim is to retrack with the same precision both open ocean and coastal data with the same algorithm. ALES selects part of each returned echo and models it with a classic ‘open ocean’ Brown functional form, by means of least square estimation whose convergence is found through the Nelder­Mead nonlinear optimization technique. By avoiding echoes from bright targets along the trailing edge, it is capable of retrieving the majority of coastal waveform up to 2 to 3 Km from the coasts. By adapting the estimation window to the significant wave height, it aims at preserving the precision of the standard data both in open ocean and in the coastal strip. ALES is validated against tide gauges in the Adriatic Sea and in the Greater Agulhas System for three different missions: Envisat, Jason­1 and Jason­2. Considerations on noise and biases provide a further verification of the strategy.

  15. 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.

  16. Air forces and moments on triangular and related wings with subsonic leading edges oscillating in supersonic potential flow

    NASA Technical Reports Server (NTRS)

    Watkins, Charles E; Berman, Julian N

    1952-01-01

    This analysis treats the air forces and moments in supersonic potential flow on oscillating triangular wings and a series of sweptback and arrow wings with subsonic leading edges and supersonic trailing edges. For the wings undergoing sinusoidal torsional oscillations simultaneously with vertical translations, the linearized velocity potential is derived in the form of a power series in terms of a frequency parameter. This method can be useful for treatment of similar problems for other plan forms and for wings undergoing other sinusoidal motions. For triangular wings, as many terms of such a series expansion as may be derived can be determined; however, the terms after the first few become very cumbersome. Closed expressions that include the reduced frequency to the fifth power, an order which is sufficient for a large class of practical application, are given for the velocity potential and for the components of chordwise section force and moment coefficients. These wings are found to exhibit the possibility of undamped torsional oscillations for certain ranges of Mach number and locations of the axis of rotation. The ranges of these parameters are delineated for triangular wings.

  17. AMELIA CESTOL Test: Acoustic Characteristics of Circulation Control Wing with Leading-and Trailing-Edge Slot Blowing

    NASA Technical Reports Server (NTRS)

    Horne, Clifton; Burnside, Nathan J.

    2013-01-01

    Aeroacoustic measurements of the 11 % scale full-span AMELIA CESTOL model with leading- and trailing-edge slot blowing circulation control (CCW) wing were obtained during a recent test in the Arnold Engineering Development Center 40- by 80-Ft. Wind Tunnel at NASA Ames Research Center, Sound levels and spectra were acquired with seven in-flow microphones and a 48-element phased microphone array for a variety of vehicle configurations, CCW slot flow rates, and forward speeds, Corrections to the measurements and processing are in progress, however the data from selected configurations presented in this report confirm good measurement quality and dynamic range over the test conditions, Array beamform maps at 40 kts tunnel speed show that the trailing edge flap source is dominant for most frequencies at flap angles of 0deg and 60deg, The overall sound level for the 60deg flap was similar to the 0deg flap for most slot blowing rates forward of 90deg incidence, but was louder by up to 6 dB for downstream angles, At 100 kts, the in-flow microphone levels were louder than the sensor self-noise for the higher blowing rates, while passive and active background noise suppression methods for the microphone array revealed source levels as much as 20 dB lower than observed with the in-flow microphones,

  18. 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.

  19. Advanced X-Ray Inspection of Reinforced Carbon Composite Materials on the Orbiter Leading Edge Structural Subsystem (LESS)

    NASA Technical Reports Server (NTRS)

    Hernandez, Jose M.; Berry, Robert F.; Osborn, Robin; Bueno, Clifford; Osterlitz, Mark; Mills, Richard; Morris, Philip; Phalen, Robert; McNab, Jim; Thibodeaux, Tahanie; Thompson, Kyle

    2004-01-01

    The post return-to-flight (RTF) inspection methodology for the Orbiter Leading Edge Structural Subsystem (LESS) is currently being defined. Numerous NDT modalities and techniques are being explored to perform the flight-to-flight inspections of the reinforced carbon/carbon (RCC) composite material for impact damage, general loss of mass in the bulk layers, or other anomalous conditions that would pose risk to safe return upon re-entry. It is possible to have an impact upon ascent that is not visually observable on the surface, yet causes internal damage. Radiographic testing may be a useful NDT technique for such occurrences. The authors have performed radiographic tests on full-sized mock samples of LESS hardware with embedded image quality phantoms. Digitized radiographic film, computed radiography and flat panel digital real-time radiography was acquired using a GE Eresco 200 x-ray tube, and Se-75 and Yb-169 radioisotopes.

  20. 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.

  1. The novel RacE-binding protein GflB sharpens Ras activity at the leading edge of migrating cells

    PubMed Central

    Senoo, Hiroshi; Cai, Huaqing; Wang, Yu; Sesaki, Hiromi; Iijima, Miho

    2016-01-01

    Directional sensing, a process in which cells convert an external chemical gradient into internal signaling events, is essential in chemotaxis. We previously showed that a Rho GTPase, RacE, regulates gradient sensing in Dictyostelium cells. Here, using affinity purification and mass spectrometry, we identify a novel RacE-binding protein, GflB, which contains a Ras GEF domain and a Rho GAP domain. Using biochemical and gene knockout approaches, we show that GflB balances the activation of Ras and Rho GTPases, which enables cells to precisely orient signaling events toward higher concentrations of chemoattractants. Furthermore, we find that GflB is located at the leading edge of migrating cells, and this localization is regulated by the actin cytoskeleton and phosphatidylserine. Our findings provide a new molecular mechanism that connects directional sensing and morphological polarization. PMID:27009206

  2. The novel RacE-binding protein GflB sharpens Ras activity at the leading edge of migrating cells.

    PubMed

    Senoo, Hiroshi; Cai, Huaqing; Wang, Yu; Sesaki, Hiromi; Iijima, Miho

    2016-05-15

    Directional sensing, a process in which cells convert an external chemical gradient into internal signaling events, is essential in chemotaxis. We previously showed that a Rho GTPase, RacE, regulates gradient sensing in Dictyostelium cells. Here, using affinity purification and mass spectrometry, we identify a novel RacE-binding protein, GflB, which contains a Ras GEF domain and a Rho GAP domain. Using biochemical and gene knockout approaches, we show that GflB balances the activation of Ras and Rho GTPases, which enables cells to precisely orient signaling events toward higher concentrations of chemoattractants. Furthermore, we find that GflB is located at the leading edge of migrating cells, and this localization is regulated by the actin cytoskeleton and phosphatidylserine. Our findings provide a new molecular mechanism that connects directional sensing and morphological polarization. PMID:27009206

  3. Icing tunnel tests of a glycol-exuding porous leading edge ice protection system on a general aviation airfoil

    NASA Technical Reports Server (NTRS)

    Kohlman, D. L.; Schweikhard, W. G.; Albright, A. E.; Evanich, P.

    1981-01-01

    A glycol-exuding porous leading edge ice protection system was tested. Results show that the system is very effective in preventing ice accretion (anti-ice mode) or removing ice from an airfoil. Minimum glycol flow rates required for anti-icing are a function of velocity, liquid water content in the air, ambient temperature, and droplet size. Large ice caps were removed in only a few minutes using anti-ice flow rates. It was found that the shed time is a function of the type of ice, size of the ice cap, angle of attack, and glycol flow rate. Wake survey measurements show that there is no significant drag penalty for the installation or operation of the system tested.

  4. 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.

  5. Turbine vane gas film cooling with injection in the leading edge region from a single row of spanwise angled holes

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

    An experimental study of gas film cooling was conducted on a 3X size model turbine vane. Injection in the leading edge region was from a single row of holes angled in a spanwise direction. Measurements of the local heat flux downstream from the row of coolant holes, both with and without film coolant flow, were used to determine the film cooling performance presented in terms of the Stanton number ratio. Results for a range of coolant blowing ratio, M = 0 to 2.0, indicate a reduction in heat flux of up to 15 to 30 percent at a point 10 to 11 hole diameters downstream from injection. An optimum coolant blowing ratio corresponds to a coolant-to-freestream velocity ratio in the range of 0.5. The shallow injection angle resulted in superior cooling performance for injection closest to stagnation, while the effect of injection angle was insignificant for injection further from stagnation.

  6. 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.

  7. Test-Analysis Correlation for Space Shuttle External Tank Foam Impacting RCC Wing Leading Edge Component Panels

    NASA Technical Reports Server (NTRS)

    Lyle, Karen H.

    2008-01-01

    The Space Shuttle Columbia Accident Investigation Board recommended that NASA develop, validate, and maintain a modeling tool capable of predicting the damage threshold for debris impacts on the Space Shuttle Reinforced Carbon-Carbon (RCC) wing leading edge and nosecap assembly. The results presented in this paper are one part of a multi-level approach that supported the development of the predictive tool used to recertify the shuttle for flight following the Columbia Accident. The assessment of predictive capability was largely based on test analysis comparisons for simpler component structures. This paper provides comparisons of finite element simulations with test data for external tank foam debris impacts onto 6-in. square RCC flat panels. Both quantitative displacement and qualitative damage assessment correlations are provided. The comparisons show good agreement and provided the Space Shuttle Program with confidence in the predictive tool.

  8. Aerodynamic forces and flow structures of the leading edge vortex on a flapping wing considering ground effect.

    PubMed

    Van Truong, Tien; Byun, Doyoung; Kim, Min Jun; Yoon, Kwang Joon; Park, Hoon Cheol

    2013-09-01

    The aim of this work is to provide an insight into the aerodynamic performance of the beetle during takeoff, which has been estimated in previous investigations. We employed a scaled-up electromechanical model flapping wing to measure the aerodynamic forces and the three-dimensional flow structures on the flapping wing. The ground effect on the unsteady forces and flow structures were also characterized. The dynamically scaled wing model could replicate the general stroke pattern of the beetle's hind wing kinematics during takeoff flight. Two wing kinematic models have been studied to examine the influences of wing kinematics on unsteady aerodynamic forces. In the first model, the angle of attack is asymmetric and varies during the translational motion, which is the flapping motion of the beetle's hind wing. In the second model, the angle of attack is constant during the translational motion. The instantaneous aerodynamic forces were measured for four strokes during the beetle's takeoff by the force sensor attached at the wing base. Flow visualization provided a general picture of the evolution of the three-dimensional leading edge vortex (LEV) on the beetle hind wing model. The LEV is stable during each stroke, and increases radically from the root to the tip, forming a leading-edge spiral vortex. The force measurement results show that the vertical force generated by the hind wing is large enough to lift the beetle. For the beetle hind wing kinematics, the total vertical force production increases 18.4% and 8.6% for the first and second strokes, respectively, due to the ground effect. However, for the model with a constant angle of attack during translation, the vertical force is reduced during the first stroke. During the third and fourth strokes, the ground effect is negligible for both wing kinematic patterns. This finding suggests that the beetle's flapping mechanism induces a ground effect that can efficiently lift its body from the ground during takeoff

  9. An experimental analysis of critical factors involved in the breakdown process of leading edge vortex flows. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Visser, Kenneth D.

    1991-01-01

    Experimental crosswire measurements of the flowfield above a 70 and 75 degree flat plate delta wing were performed at a Reynolds number of 250,000. Survey grids were taken normal to the platform at a series of chordwise locations for angles of attack of 20 and 30 degrees. Axial and azimuthal vorticity distributions were derived from the velocity fields. The dependence of circulation on distance from the vortex core as well as on chordwise location was examined. The effects of nondimensionalization in comparison with other experimental data was made. The circulation distribution scales with the local semispan and grows approximately linearly in the chordwise direction. For regions of the flow outside of the vortex subcore, the circulation at any chordwise station was observed to vary logarithmically with distance from the vortex axis. The circulation was also found to increase linearly with angle of incidence at a given chordwise station. A reduction in the local circulation about the vortex axis occurred at breakdown. The spanwise distribution of axial vorticity was severely altered through the breakdown region and the spanwise distribution of axial vorticity present appeared to reach a maximum immediately preceding breakdown. The local concentration of axial vorticity about the vortex axis was reduced while the magnitude of the azimuthal vorticity decreased throughout the breakdown zone. The axial vorticity components with a negative sense, found in the secondary vortex, remained unaffected by changes in wing sweep or angle of attack, in direct contrast to the positive components. The inclusion of the local wing geometry into a previously derived correlation parameter indicated that the circulation of growing leading edge vortex flows were similar at corresponding radii from the vortex axis. It was concluded that the flow over a delta wing, upstream of the breakdown regions and away from the apex and trailing edge regions, is conical. In addition, the dominating

  10. Rate and topography of peptidoglycan synthesis during cell division in Escherichia coli: Concept of a leading edge

    SciTech Connect

    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 surface 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.

  11. 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.

  12. 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.

  13. 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.

  14. ESA Intermediate Experimental Vehicle. Independent Aerothermodynamic Characterization And Aerodatabase Development

    NASA Astrophysics Data System (ADS)

    Rufolo, Giuseppe C.; Di Benedetto, Sara; Walpot, Louis; Roncioni, Pietro; Marini, Marco

    2011-05-01

    In the frame of the Intermediate eXperimental Vehicle (IXV) project, the European Space Agency (ESA) is coordinating a series of technical assistance activities aimed at verifying and supporting the IXV industrial design and development process. The technical assistance is operated with the support of the Italian Space Agency (ASI), by means of the Italian Aerospace Research Center (CIRA), and the European Space Research and Technology Centre (ESTEC) under the super visioning and coordination of ESA IXV team. One of the purposes of the activity is to develop an independent capability for the assessment and verification of the industrial results with respect to the aerothermodynamic characterization of the IXV vehicle. To this aim CIRA is developing and independent AeroThermodynamics DataBase (ATDB), intended as a tool generating in output the time histories of local quantities (heat flux, pressure, skin friction) for each point of the IXV vehicle and for each trajectory (in a pre-defined envelope), together with an uncertainties model. The reference Computational Fluid Dynamics (CFD) solutions needed for the development of the tool have been provided by ESA-ESTEC (with the CFD code LORE) and CIRA (with the CFD code H3NS).

  15. 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.

  16. 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.

  17. Effects of Outboard Thickened and Blunted Leading Edges on the Wave Drag of a 45 Degree Swept-Wing and Body Combination

    NASA Technical Reports Server (NTRS)

    Holdaway, George H.; Lazzeroni, Frank A.; Hatfield, Elaine W.

    1959-01-01

    An investigation to evaluate the effects of thickened and blunted leading-edge modifications on the wave drag of a swept wing has been made at Mach numbers from 0.65 to 2.20 and at a Reynolds number of 2,580,000 based on the mean aerodynamic chord of the basic wing. Two leading-edge designs were investigated and they are referred to as the thickened and the blunted modifications although both sections had equally large leading-edge radii. The thickened leading edge was formed by increasing the thickness over the forward 40 percent of the basic wing section. The blunted modification was formed by reducing the wing chords about 1 percent and by increasing the section thickness slightly over the forward 6 percent of the basic section in a manner to keep the wing sweep and volume essentially equal to the respective values for the basic wing. The basic wing had an aspect ratio of 3, a leading-edge sweep of 45 deg., a taper ratio of 0.4, and NACA 64AO06 sections perpendicular to a line swept back 39.45 deg., the quarter-chord line of these sections. Test results indicated that the thickened modification resulted in an increase in zero-lift drag coefficient of from 0.0040 to 0.0060 over values for the basic model at Mach numbers at which the wing leading edge was sonic or supersonic. Although drag coefficients of both the basic and thickened models were reduced at all test Mach numbers by body indentations designed for the range of Mach numbers from 1.00 to 2.00, the greater drag of the thickened model relative to that of the basic model was not reduced. The blunted model, however, had less than one quarter of the drag penalty of the thickened model relative to the basic model at supersonic leading-edge conditions (M greater or equal to root-2).

  18. 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.

  19. 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.

  20. Icing tunnel tests of a composite porous leading edge for use with a liquid anti-ice system. [Lewis icing research tunnel

    NASA Technical Reports Server (NTRS)

    Kohlman, D. L.

    1981-01-01

    The efficacy of liquid ice protection systems which distribute a glycol-water solution onto leading edge surfaces through a porous skin was demonstrated in tests conducted in the NASA Lewis icing research tunnel using a composite porous leading edge panels. The data obtained were compared with the performance of previously tested stainless steel leading edge with the same geometry. Results show: (1) anti-ice protection of a composite leading edge is possible for all the simulated conditions tested; (2) the glycol flow rates required to achieve anti-ice protection were generally much higher than those required for a stainless steel panel; (3) the low reservoir pressures of the glycol during test runs indicates that more uniform distribution of glycol, and therefore lower glycol flow rates, can probably be achieved by decreasing the porosity of the panel; and (4) significant weight savings can be achieved in fluid ice protection systems with composite porous leading edges. The resistance of composite panels to abrasion and erosion must yet be determined before they can be incorporated in production systems.

  1. 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.

  2. 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.

  3. Local heat transfer in internally cooled turbine airfoil leading edge regions. I - Impingement cooling without film coolant extraction. II - Impingement cooling with film coolant extraction

    NASA Astrophysics Data System (ADS)

    Bunker, R. S.; Metzger, D. E.

    The highly localized internal heat transfer characteristics of large-scale models of impingement-cooled turbine blade leading edge regions presently studied derives its cooling from a single line of equally-spaced multiple jets aimed at the leading-edge apex, and exiting the leading-edge region in the opposite or chordwise direction. Detailed two-dimensional local surface Nusselt number distributions have been obtained with temperature-indicating coatings. Results indicate generally increasing heat transfer with the 0.6 power of jet Reynolds number. In the second part of this study, in which the same cooling process is used in conjunction with the extraction of the coolant fluid, the results obtained indicate that heat transfer is primarily dependent on jet Reynolds number, with smaller influences from the flow-extraction rate.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Measurements of heat transfer coefficients and friction factors in rib-roughened channels simulating leading-edge cavities of a modern turbine blade

    SciTech Connect

    Taslim, M.E.; Li, T.; Spring, S.D.

    1997-07-01

    Leading edge cooling cavities in modern gas turbine blades play an important role in maintaining the leading edge temperature at levels consistent with air foil design life. These cavities often have a complex cross-sectional shape to be compatible with the external contour of the blade at the leading edge. A survey of many existing geometries shows that, for analytical as well as experimental analyses, such cavities can be simplified in shape by a four-sided polygon with one curved side similar to the leading edge curvature, a rectangle with one semicircular side (often the smaller side) or a trapezoid, the smaller base of which is replaced by a semicircle. Furthermore, to enhance the heat transfer coefficient in these cavities, they are mostly roughened on three sides with ribs of different geometries. Experimental data on friction factors and heat transfer coefficients in such cavities are rare if not nonexistent. A liquid crystal technique was used in this experimental investigation to measure heat transfer coefficients in six test sections representing the leading-edge cooling cavities. Both straight and tapered ribs were configured on the two opposite sidewalls in a staggered arrangement with angles of attack to the mainstream flow, {alpha}, of 60 and 90 deg. The ribs on the curved surface were of constant cross section with an angle of attack 90 deg to the flow. Heat transfer measurements were performed on the straight sidewalls, as well as on the round surface adjacent to the blade leading edge. Effects such as rib angle of attack to the mainstream flow and constant versus tapered rib cross-sectional areas were also investigated. Nusselt numbers, friction factors, and thermal performances are reported for nine rib geometries in six test sections.

  9. Experimental Observations on the Deformation and Breakup of Water Droplets Near the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Feo, Alex

    2011-01-01

    This work presents the results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model placed at the end of the rotating arm was moved at speeds of 50 to 90 m/sec. A monosize droplet generator was employed to produce droplets that were allowed to fall from above, perpendicular to the path of the airfoil at a given location. High speed imaging was employed to observe the interaction between the droplets and the airfoil. The high speed imaging allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. A tracking software program was used to measure from the high speed movies the horizontal and vertical displacement of the droplet against time. The velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of a given droplet from beginning of deformation to breakup and/or hitting the airfoil. Results are presented for droplets with a diameter of 490 micrometers at airfoil speeds of 50, 60, 70, 80 and 90 m/sec

  10. Goertler vortices in growing boundary layers: The leading edge receptivity problem, linear growth and the nonlinear breakdown stage

    NASA Technical Reports Server (NTRS)

    Hall, Philip

    1989-01-01

    Goertler vortices are thought to be the cause of transition in many fluid flows of practical importance. A review of the different stages of vortex growth is given. In the linear regime, nonparallel effects completely govern this growth, and parallel flow theories do not capture the essential features of the development of the vortices. A detailed comparison between the parallel and nonparallel theories is given and it is shown that at small vortex wavelengths, the parallel flow theories have some validity; otherwise nonparallel effects are dominant. New results for the receptivity problem for Goertler vortices are given; in particular vortices induced by free stream perturbations impinging on the leading edge of the walls are considered. It is found that the most dangerous mode of this type can be isolated and it's neutral curve is determined. This curve agrees very closely with the available experimental data. A discussion of the different regimes of growth of nonlinear vortices is also given. Again it is shown that, unless the vortex wavelength is small, nonparallel effects are dominant. Some new results for nonlinear vortices of 0(1) wavelengths are given and compared to experimental observations.

  11. Leading edge vortices in lesser long-nosed bats occurring at slow but not fast flight speeds.

    PubMed

    Muijres, Florian T; Christoffer Johansson, L; Winter, York; Hedenström, Anders

    2014-06-01

    Slow and hovering animal flight creates high demands on the lift production of animal wings. Steady state aerodynamics is unable to explain the forces required and the most commonly used mechanism to enhance the lift production is a leading edge vortex (LEV). Although LEVs increase the lift, they come at the cost of high drag. Here we determine the flow above the wing of lesser long-nosed bats at slow and cruising speed using particle image velocimetry (PIV). We find that a prominent LEV is present during the downstroke at slow speed, but not at cruising speed. Comparison with previously published LEV data from a robotic flapper inspired by lesser long-nosed bats suggests that bats should be able to generate LEVs at cruising speeds, but that they avoid doing so, probably to increase flight efficiency. In addition, at slow flight speeds we find LEVs of opposite spin at the inner and outer wing during the upstroke, potentially providing a control challenge to the animal. We also note that the LEV stays attached to the wing throughout the downstoke and does not show the complex structures found in insects. This suggests that bats are able to control the development of the LEV and potential control mechanisms are discussed. PMID:24855067

  12. A numerical investigation of new film cooling hole configuration at the leading edge of asymmetrical turbine blade: part A

    NASA Astrophysics Data System (ADS)

    Benabed, Mustapha

    2013-04-01

    The focus of the first part of this numerical study is to investigate the effects of two new configurations: (1) slot with cylindrical end and (2) slot with median cylindrical hole, generated by the combination between two film cooling configurations: cylindrical hole and uniform slot. Computational results are presented for a row of coolant injection holes on each side of an asymmetrical turbine blade model near the leading edge. For each configuration, three values of the radius are taken: R = 0.4, R = 0.8 and R = 1.2. The six cases simulations, thus obtained, are conducted for the same density ratio of 1.0 and the same inlet plenum pressure. A new parameter, Rc, is defined to measure the rate of blade coverage by the film cooling. Results show that, at the pressure side; for the two new configurations, the six studied cases exceed the case baseline in cooling effectiveness term with the best result obtained for R = 0.8 (case 2). For the suction side, only configurations with R = 0.4 (cases 1 and 4) provide an increase of film effectiveness compared to the case baseline. The following configuration: Cases 1 or 4 at the suction side and case 2 at the pressure side, gets the best thermal protection because of their higher coverage and strong cooling effectiveness.

  13. Do rapid 'superbug' tests pay off? Balance the costs and benefits of leading-edge technology. Interview by Alan Joch.

    PubMed

    Morgan, Margie Ann

    2009-02-01

    As hospitals become increasingly sensitive to the health and financial consequences of health care-associated infections (HAIs), a new generation of molecular-based testing technologies promises to significantly shorten the time required to identify "superbugs" and other bacterial infections. The leading-edge techniques promise to reduce costs by helping hospitals quickly determine which patients to isolate because they carry active methicillin-resistant Staphylococcus aureus (MRSA) infections, for example, or which ones to release from prophylactic isolation because they ultimately tested negative for a dangerous infection. But diagnostic speed comes at a price--the costs to perform molecular tests are significantly higher than conventional methods. This challenges hospitals to balance health care expenses with medical efficacy, says molecular testing veteran Margie Morgan, Ph.D., scientific director at Cedars-Sinai Medical Center, Los Angeles. "The rapid methods can be extreme time savers and possibly help a great deal with the isolation of patients. But some of the tests may cost five times what manual methods might be, so there is a price for seeing so much of a reduction in time," she says. PMID:19288674

  14. A Search for Compositional Differences in Slow Solar Wind at the Leading and Trailing Edges of Stream Interfaces

    NASA Astrophysics Data System (ADS)

    Kahler, Stephen

    2010-03-01

    The elemental and ionic composition of the slow (~400 km/s) solar wind (SW) differs from that of the fast (>600 km/s) SW streams. In particular, the first ionization potential (FIP) effect and the O7/O6 ratios are higher in the slow SW than in the fast SW. While this fundamental difference between fast and slow SW has long been appreciated, there has been no reason to expect or effort to search for systematic variations within given fast or slow SW streams. If present, however, such systematic variations could be diagnostic of the solar processes at the sources of the SW. We suggest two reasons to expect compositional differences between leading and trailing edges of slow SW streams. After selecting two long-lived fast SW streams from 2005 and 2006 and determining their slow-fast (SF) and fast-slow (FS) stream interfaces (SIs), ACE SWICS (version 3) compositional data were obtained for the 1-day periods of the preceding SF and following FS slow SW streams. The statistics were limited, but no compositional differences between the preceding and following slow SW regions were found for either stream sequence.

  15. Modeling the Nonlinear, Strain Rate Dependent Deformation of Shuttle Leading Edge Materials with Hydrostatic Stress Effects Included

    NASA Technical Reports Server (NTRS)

    Goldberg, Robert K.; Carney, Kelly S.

    2004-01-01

    An analysis method based on a deformation (as opposed to damage) approach has been developed to model the strain rate dependent, nonlinear deformation of woven ceramic matrix composites, such as the Reinforced Carbon Carbon (RCC) material used on the leading edges of the Space Shuttle. In the developed model, the differences in the tension and compression deformation behaviors have also been accounted for. State variable viscoplastic equations originally developed for metals have been modified to analyze the ceramic matrix composites. To account for the tension/compression asymmetry in the material, the effective stress and effective inelastic strain definitions have been modified. The equations have also been modified to account for the fact that in an orthotropic composite the in-plane shear response is independent of the stiffness in the normal directions. The developed equations have been implemented into LS-DYNA through the use of user defined subroutines (UMATs). Several sample qualitative calculations have been conducted, which demonstrate the ability of the model to qualitatively capture the features of the deformation response present in woven ceramic matrix composites.

  16. Partners in Science: A Model Cooperative Program Introducing High School Teachers and Students to Leading-Edge Pharmaceutical Science

    NASA Astrophysics Data System (ADS)

    Woska, Joseph R., Jr.; Collins, Danielle M.; Canney, Brian J.; Arcario, Erin L.; Reilly, Patricia L.

    2005-12-01

    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 a series of lectures, tours, and demonstrations given by scientists within our research and development division (R&D). Phase 2 involves the selection of a small group of participants to intern for the summer in a research laboratory, working side by side with a scientist within R&D. In this manuscript, the specific aims, goals, and development of the Partners in Science program are described, as well as the syllabus/agenda, the logistics surrounding the operation of the program, and our shared personal experiences with students and teachers who have participated. Some of the pitfalls/problems associated with the program will be presented, and finally, the future direction of the program including areas of improvement and expansion are described.

  17. Endocytic membrane turnover at the leading edge is driven by a transient interaction between Cdc42 and GRAF1

    PubMed Central

    Francis, Monika K.; Holst, Mikkel R.; Vidal-Quadras, Maite; Henriksson, Sara; Santarella-Mellwig, Rachel; Sandblad, Linda; Lundmark, Richard

    2015-01-01

    ABSTRACT Changes in cell morphology require coordination of plasma membrane turnover and cytoskeleton dynamics, processes that are regulated by Rho GTPases. Here, we describe how a direct interaction between the Rho GTPase Cdc42 and the GTPase-activating protein (GAP) GRAF1 (also known as ARHGAP26), facilitates rapid cell surface turnover at the leading edge. Both Cdc42 and GRAF1 were required for fluid-phase uptake and regulated the generation of transient GRAF1-coated endocytic carriers, which were distinct from clathrin-coated vesicles. GRAF1 was found to transiently assemble at discrete Cdc42-enriched punctae at the plasma membrane, resulting in a corresponding decrease in the microdomain association of Cdc42. However, Cdc42 captured in its active state was, through a GAP-domain-mediated interaction, localised together with GRAF1 on accumulated internal structures derived from the cell surface. Correlative fluorescence and electron tomography microscopy revealed that these structures were clusters of small membrane carriers with defective endosomal processing. We conclude that a transient interaction between Cdc42 and GRAF1 drives endocytic turnover and controls the transition essential for endosomal maturation of plasma membrane internalised by this mechanism. PMID:26446261

  18. Effects of Segmented Slot Blowing at the Leading Edge of a Finite Span Cavity in Supersonic Flow

    NASA Astrophysics Data System (ADS)

    George, Benjamin; Ukeiley, Lawrence; Cattafesta, Louis; Taira, Kunihiko

    2015-11-01

    In this investigation, the effects of employing segmented slot blowing at the leading edge of a finite span cavity in Mach 1.4 flow are studied. The rectangular cavity under consideration has a length to depth ratio of 6 and width to depth ratio of 2 with an approaching turbulent boundary layer. Qualitative surface flow visualization results reveal changes in the flow characteristics due to the introduction of the sidewalls and multiple slot blowing configurations, as has been previously shown. Quantitatively, unsteady surface pressure measurements and particle image velocimetry (PIV) were utilized to characterize the mechanisms for suppressing surface pressure fluctuations in a three-dimensional flow field. Joint time-frequency analysis using wavelet transformations highlight changes in the tonal and broadband surface pressure fluctuations as a function of time with the different slot configurations. PIV data results from the baseline finite span case were compared with the slot blowing cases to illustrate their effects on the mean flow field properties in the shear layer and recirculation region. Finally, the finite span cavity experimental results are compared with previously acquired data for the full span cavity case to gain some insight into the flow field modifications as the cavity span is altered.

  19. Aerothermodynamic design feasibility of a Mars aerocapture/aeromaneuver vehicle

    NASA Technical Reports Server (NTRS)

    Florence, D. E.

    1981-01-01

    Lifting aerodynamic configurations have been screened and selected for the Mars aerocapture mission that (1) meet the geometric packaging requirements of the various payloads and the Space Shuttle cargo bay and (2) provide the aerodynamic performance characteristics required to obtain the atmospheric exit steering accuracy and the parachute deployment conditions desired. Hypersonic heat transfer and aerodynamic loads to the vehicle in the CO2 atmosphere are evaluated. Contemporary low density ablative thermal protection materials were selected that meet all the atmospheric entry requirements and provide a minimum mass solution. Results are presented of the aerodynamic configuration and thermal protection materials screening and selection. It is concluded that the aerothermodynamic design of this concept is feasible using state-of-the-art technology.

  20. Aerothermodynamic methods for a Mars environmental survey Mars entry

    NASA Technical Reports Server (NTRS)

    Mitcheltree, R. A.

    1994-01-01

    Computational fluid dynamics models for the thermodynamics and transport properties used in an equilibrium version of the Langley aerothermodynamics upwind relaxation algorithm (LAURA) for Mars atmospheric entries are described. In addition, the physical models used in a nonequilibrium version of LAURA for Mars-entry flows are described. Uncertainties exist in defining constants used in the transport properties for the equilibrium model and in many of the physical models for the nonequilibrium version. Solutions from the two codes using the best available constants are examined at the Mars-entry conditions characteristics of the Mars environmental survey mission. While the flowfields are near thermal equilibrium, chemical nonequilibrium effects are present in the entry cases examined. Convective heating at the stagnation point for these flows (assuming fully catalytic wall boundary conditions) is approximately 100 W/cm(exp 2). Radiative heating is negligible.

  1. 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

  2. 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.

  3. Aerothermodynamic Facilities And Measurement: Flow Characterization in Shock Tunnels

    NASA Technical Reports Server (NTRS)

    Cavolowsky, John A.; Edwards, Thomas A. (Technical Monitor)

    1995-01-01

    This presentation will examine the key performance aspects of shock tunnels as they relate to their use as aerothermodynamic flow simulation facilities. Assessment of shock tube reservoir conditions and flow contaminants generated in the shock tube will be presented along with their limiting impact on viable test envelopes, Facility nozzle performance as it pertains to test time assessment and nozzle exit flow quality (survey of pressure, temperature, and species) will be addressed. Also included will be a discussion of free stream flow diagnostics, both intrusive and nonintrusive, for measurement of critical flow properties not directly inferred from surface mounted transducers. The use of computational fluid dynamics for purposes of validating experimental measurements as well as predicting performance in regimes where measurements are not feasible or possible will be discussed. The use of CFD for facility research and design will also be presented.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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 linearly at the trailing edge

    NASA Technical Reports Server (NTRS)

    Hague, D. S.; Werz, A. W.

    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 airfoils. The additional thickness distribution had the form of a continuous mathematical function which disappears at both the leading edge and the trailing edge. Results were obtained at a Mach number of 0.2 with an angle of attack of 6 deg. All calculations employed the full potential flow equations for two dimensional flow. The relaxation method of Jameson was used for solution of the potential flow equations. It is shown that increasing the thickness and variations in shape increases the lift and the adverse pitching moment coefficients.

  9. Salinization of Freshwater-Dependent Coastal Ecosystems: Understanding Landscapes in Transition Along the Leading Edge of Climate Change

    NASA Astrophysics Data System (ADS)

    Emanuel, R. E.; Bernhardt, E. S.; Ardón, M.; Wright, J. P.; BenDor, T.; Bhattachan, A.

    2015-12-01

    Climate change is transforming the outer edge of the Southern US coastal plain. Lower-lying parts of this region, characterized by extensive freshwater-dependent ecosystems, will be largely inundated by gradual sea level rise by the end of this century. In the interim, however, ocean waters are already penetrating and influencing freshwater-dependent coastal landscapes due to a combination of human and natural factors. This landward movement of salinity from the coast onto the coastal plain or "saltwater intrusion" is a critical water resource issue representing the leading edge of climate change for many coastal areas. The salinization of surface waters and adjacent lands has implications for crop and timber yields in managed ecosystems, ecosystem carbon sequestration in unmanaged ecosystems, and degradation of coastal water quality due to extraction of soil nutrients by seasalts. With this in mind, we seek to understand more broadly how vulnerability of coastal landscapes to saltwater intrusion shapes and is shaped by both natural and anthropogenic processes. We present a novel framework that couples intensive, in situ monitoring of hydrological and ecological conditions with a geospatial saltwater intrusion vulnerability index (SIVI). We discuss application of this framework to the Albemarle-Pamlico region of coastal North Carolina, where we are learning how climate, natural landscape structure, and human activities interact to mediate or exacerbate the vulnerability of freshwater-dependent lands to saltwater intrusion. We discuss the involvement of stakeholders and local knowledge in the research process as well. This work advances understanding of vulnerability to climate change in coastal regions by moving beyond simple inundation models to gain a more sophisticated understanding of the human and natural processes influencing salinization of surface waters and adjacent lands. As the Albemarle-Pamlico and similar regions worldwide transform in response to and

  10. Aerothermodynamic Characteristics in the Hypersonic Continuum-Rarefied Transitional Regime

    NASA Technical Reports Server (NTRS)

    Glass, Christopher, E.; Moss, James N.

    2001-01-01

    Stagnation pressure and heating for cylinders and spheres representative of wing leading edges and nose sections for small-scale space transportation vehicles are presented for various Knudsen numbers at Earth entry conditions that bridge from the free molecular to continuum regimes. CFD, DSMC, and collisionless DSMC are used to span Knudsen numbers from 0.001 to 100 with the CFD/DSMC overlap at a Knudsen number of 0.01 and the DSMC/collisionless DSMC overlap at 10. The present results include density contours, density and temperature along the stagnation streamline for all cases, and stagnation pressure and heat transfer coefficient as a function of Knudsen number.

  11. 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

  12. 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

  13. The importance of leading edge vortices under simplified flapping flight conditions at the size scale of birds.

    PubMed

    Hubel, Tatjana Y; Tropea, Cameron

    2010-06-01

    Over the last decade, interest in animal flight has grown, in part due to the possible use of flapping propulsion for micro air vehicles. The importance of unsteady lift-enhancing mechanisms in insect flight has been recognized, but unsteady effects were generally thought to be absent for the flapping flight of larger animals. Only recently has the existence of LEVs (leading edge vortices) in small vertebrates such as swifts, small bats and hummingbirds been confirmed. To study the relevance of unsteady effects at the scale of large birds [reduced frequency k between 0.05 and 0.3, k=(pifc)/U(infinity); f is wingbeat frequency, U(infinity) is free-stream velocity, and c is the average wing chord], and the consequences of the lack of kinematic and morphological refinements, we have designed a simplified goose-sized flapping model for wind tunnel testing. The 2-D flow patterns along the wing span were quantitatively visualized using particle image velocimetry (PIV), and a three-component balance was used to measure the forces generated by the wings. The flow visualization on the wing showed the appearance of LEVs, which is typically associated with a delayed stall effect, and the transition into flow separation. Also, the influence of the delayed stall and flow separation was clearly visible in measurements of instantaneous net force over the wingbeat cycle. Here, we show that, even at reduced frequencies as low as those of large bird flight, unsteady effects are present and non-negligible and have to be addressed by kinematic and morphological adaptations. PMID:20472780

  14. Prediction of STS-107 Hypervelocity Flow Fields about the Shuttle Orbiter with Various Wing Leading Edge Damage

    NASA Technical Reports Server (NTRS)

    Pulsonetti, Maria V.; Thompson, Richard A.; Alter, Stephen J.

    2004-01-01

    Computations were performed for damaged configurations of the Shuttle Orbiter in support of the STS-107 Columbia accident investigation. Two configurations with missing wing leading-edge reinforced carbon-carbon (RCC) panels were evaluated at conditions just prior to the peak heating trajectory point. The initial configuration modeled the Orbiter with an approximate missing RCC panel 6 to determine whether this damage could result in anomalous temperatures measured during the STS-107 reentry. This missing RCC panel 6 computation was found to produce heating augmentation factors of 5 times the nominal heating rates on the side fuselage with lesser heat increases on the front of the OMS pod. This is consistent with the thermocouple and resistance temperature detector sensors from the STS-107 re-entry which observed off nominal high early in the re-entry trajectory. A second damaged configuration modeled the Orbiter with missing RCC panel 9 and included ingestion of the flow into the outboard RCC channel. This computation lowered the level (only 2 times nominal) and moved the location of the heating augmentation on the leeside fuselage relative to the missing RCC panel 6 configuration. The lesser heating augmentation for missing RCC panel 9 was confined near the wing fuselage juncture. Near nominal heating was predicted on the remainder of the side fuselage with some lower than nominal heating on the front surface of the OMS pod. These results for missing RCC panel 9 are consistent with data from the STS-107 re-entry where the heating augmentation was observed to move off the side fuselage and OMS pod sensors at later times in the trajectory. As this solution requires supersonic mass ingestion into the RCC channel, it is probably not an appropriate model prior to penetration of the flow through the spar into the wing structure. It may, however, be representative of the conditions at later times and could account for the movement of the heating signature on the side

  15. Human phosphatase CDC14A is recruited to the cell leading edge to regulate cell migration and adhesion.

    PubMed

    Chen, Nan-Peng; Uddin, Borhan; Voit, Renate; Schiebel, Elmar

    2016-01-26

    Cell adhesion and migration are highly dynamic biological processes that play important roles in organ development and cancer metastasis. Their tight regulation by small GTPases and protein phosphorylation make interrogation of these key processes of great importance. We now show that the conserved dual-specificity phosphatase human cell-division cycle 14A (hCDC14A) associates with the actin cytoskeleton of human cells. To understand hCDC14A function at this location, we manipulated native loci to ablate hCDC14A phosphatase activity (hCDC14A(PD)) in untransformed hTERT-RPE1 and colorectal cancer (HCT116) cell lines and expressed the phosphatase in HeLa FRT T-Rex cells. Ectopic expression of hCDC14A induced stress fiber formation, whereas stress fibers were diminished in hCDC14A(PD) cells. hCDC14A(PD) cells displayed faster cell migration and less adhesion than wild-type controls. hCDC14A colocalized with the hCDC14A substrate kidney- and brain-expressed protein (KIBRA) at the cell leading edge and overexpression of KIBRA was able to reverse the phenotypes of hCDC14A(PD) cells. Finally, we show that ablation of hCDC14A activity increased the aggressive nature of cells in an in vitro tumor formation assay. Consistently, hCDC14A is down-regulated in many tumor tissues and reduced hCDC14A expression is correlated with poorer survival of patients with cancer, to suggest that hCDC14A may directly contribute to the metastatic potential of tumors. Thus, we have uncovered an unanticipated role for hCDC14A in cell migration and adhesion that is clearly distinct from the mitotic and cytokinesis functions of Cdc14/Flp1 in budding and fission yeast. PMID:26747605

  16. 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.

    1983-01-01

    A proposed method of analytically predicting the minimum fluid flow rate required to provide anti-ice protection with a porous leading edge system on a wing under a given set of flight conditions is presented. Results of the proposed method are compared with the actual results of an icing test of a real wing section in the NASA Lewis Icing Research Tunnel.

  17. 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.

  18. 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.

  19. 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.

  20. 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.