Optimal operation of turbo blowers serially connected using inlet vanes

NASA Astrophysics Data System (ADS)

Jang, Choon-Man

2011-03-01

Optimal operation of the turbo blowers having an inlet vane has been studied to understand the blowers' operating performance. To analyze three-dimensional flow field in the turbo blowers serially connected, general analysis code, CFX, is introduced in the present work. SST turbulence model is employed to estimate the eddy viscosity. Throughout the numerical analysis, it is found that the flow rates of the turbo blowers can be controlled at the vane angle between 90 (full open condition) degrees and 60 degrees effectively, because pressure loss rapidly increases below 60 degree of a vane angle. Efficiency also has almost the same values from 90 degrees to 60 degrees of a vane angle. It is noted that the distorted inlet velocity generated in the small vane angle makes performance deterioration of the turbo blowers due to the local leading edge separation and the following non-uniform blade loading.

Effects of a ceramic coating on metal temperatures of an air-cooled turbine vane

NASA Astrophysics Data System (ADS)

Gladden, H. J.; Liebert, C. H.

1980-02-01

The metal temperatures of air cooled turbine vanes both uncoated and coated with the NASA thermal barrier system were studied experimentally. Current and advanced gas turbine engine conditions were simulated at reduced temperatures and pressures. Airfoil metal temperatures were significantly reduced, both locally and on the average, by use of the the coating. However, at low gas Reynolds number, the ceramic coating tripped a laminar boundary layer on the suction surface, and the resulting higher heat flux increased the metal temperatures. Simulated coating loss was also investigated and shown to increase local metal temperatures. However, the metal temperatures in the leading edge region remained below those of the uncoated vane tested at similar conditions. Metal temperatures in the trailing edge region exceeded those of the uncoated vane.

Effects of a ceramic coating on metal temperatures of an air-cooled turbine vane

NASA Technical Reports Server (NTRS)

Gladden, H. J.; Liebert, C. H.

1980-01-01

The metal temperatures of air cooled turbine vanes both uncoated and coated with the NASA thermal barrier system were studied experimentally. Current and advanced gas turbine engine conditions were simulated at reduced temperatures and pressures. Airfoil metal temperatures were significantly reduced, both locally and on the average, by use of the the coating. However, at low gas Reynolds number, the ceramic coating tripped a laminar boundary layer on the suction surface, and the resulting higher heat flux increased the metal temperatures. Simulated coating loss was also investigated and shown to increase local metal temperatures. However, the metal temperatures in the leading edge region remained below those of the uncoated vane tested at similar conditions. Metal temperatures in the trailing edge region exceeded those of the uncoated vane.

Design optimization using adjoint of Long-time LES for the trailing edge of a transonic turbine vane

NASA Astrophysics Data System (ADS)

Talnikar, Chaitanya; Wang, Qiqi

2017-11-01

Adjoint-based design optimization methods have been applied to low-fidelity simulation methods like Reynolds Averaged Navier-Stokes (RANS) and are useful for designing fluid machinery components. But to reliably capture the complex flow phenomena involved in turbomachinery, high fidelity simulations like large eddy simulation (LES) are required. Unfortunately due to the chaotic dynamics of turbulence, the unsteady adjoint method for LES diverges and produces incorrect gradients. Using a viscosity stabilized unsteady adjoint method developed for LES, the gradient can be obtained with reasonable accuracy. In this paper, design of the trailing edge of a gas turbine inlet guide vane is performed with the objective to reduce stagnation pressure loss and heat transfer over the surface of the vane. Slight changes in the shape of trailing edge can significantly impact these quantities by altering the boundary layer development process and separation points. The trailing edge is parameterized using a linear combination of 5 convex designs. Bayesian optimization is used as a global optimizer with the objective function evaluated from the LES and gradients obtained using the viscosity adjoint method. Results from the optimization, performed on the supercomputer Mira, are presented.

SiC Composite Turbine Vanes

NASA Technical Reports Server (NTRS)

Calomino, Anthony M.; Verilli, Michael J.

2006-01-01

Turbine inlet guide vanes have been fabricated from composites of silicon carbide fibers in silicon carbide matrices. A unique design for a cloth made from SiC fibers makes it possible to realize the geometric features necessary to form these vanes in the same airfoil shapes as those of prior metal vanes. The fiber component of each of these vanes was made from SiC-fiber cloth coated with boron nitride. The matrix was formed by chemical-vapor infiltration with SiC, then slurry-casting of SiC, followed by melt infiltration with silicon. These SiC/SiC vanes were found to be capable of withstanding temperatures 400 F (222 C) greater than those that can be withstood by nickel-base-superalloy turbine airfoils now in common use in gas turbine engines. The higher temperature capability of SiC/SiC parts is expected to make it possible to use them with significantly less cooling than is used for metallic parts, thereby enabling engines to operate more efficiently while emitting smaller amounts of NOx and CO. The SiC/SiC composite vanes were fabricated in two different configurations. Each vane of one of the configurations has two internal cavities formed by a web between the suction and the pressure sides of the vane. Each vane of the other configuration has no web (see Figure 1). It is difficult to fabricate components having small radii, like those of the trailing edges of these vanes, by use of stiff stoichiometric SiC fibers currently preferred for SiC/SiC composites. To satisfy the severe geometric and structural requirements for these vanes, the aforementioned unique cloth design, denoted by the term Y-cloth, was conceived (see Figure 2). In the regions away from the trailing edge, the Y-cloth features a fiber architecture that had been well characterized and successfully demonstrated in combustor liners. To form a sharp trailing edge (having a radius of 0.3 mm), the cloth was split into two planes during the weaving process. The fiber tows forming the trailing-edge

Numerical solution for the temperature distribution in a cooled guide vane blade of a radial gas turbine

NASA Technical Reports Server (NTRS)

Hosny, W. M.; Tabakoff, W.

1977-01-01

A two dimensional finite difference numerical technique is presented to determine the temperature distribution of an internal cooled blade of radial turbine guide vanes. A simple convection cooling is assumed inside the guide vane blade. Such cooling has relatively small cooling effectiveness at the leading edge and at the trailing edge. Heat transfer augmentation in these critical areas may be achieved by using impingement jets and film cooling. A computer program is written in FORTRAN IV for IBM 370/165 computer.

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.

Design Considerations for Ceramic Matrix Composite Vanes for High Pressure Turbine Applications

NASA Technical Reports Server (NTRS)

Boyle, Robert J.; Parikh, Ankur H.; Nagpal, Vinod K.; Halbig, Michael C.

2013-01-01

Issues associated with replacing conventional metallic vanes with Ceramic Matrix Composite (CMC) vanes in the first stage of the High Pressure Turbine (HPT) are explored. CMC materials have higher temperature capability than conventional HPT vanes, and less vane cooling is required. The benefits of less vane coolant are less NOx production and improved vane efficiency. Comparisons between CMC and metal vanes are made at current rotor inlet temperatures and at an vane inlet pressure of 50 atm.. CMC materials have directionally dependent strength characteristics, and vane designs must accommodate these characteristics. The benefits of reduced NOx and improved cycle efficiency obtainable from using CMC vanes. are quantified Results are given for vane shapes made of a two dimensional CMC weave. Stress components due to thermal and pressure loads are shown for all configurations. The effects on stresses of: (1) a rib connecting vane pressure and suction surfaces; (2) variation in wall thickness; and (3) trailing edge region cooling options are discussed. The approach used to obtain vane temperature distributions is discussed. Film cooling and trailing edge ejection were required to avoid excessive vane material temperature gradients. Stresses due to temperature gradients are sometimes compressive in regions where pressure loads result in high tensile stresses.

Gas turbine engine exhaust diffuser including circumferential vane

DOEpatents

Orosa, John A.; Matys, Pawel

2015-05-19

A flow passage defined between an inner and an outer boundary for guiding a fluid flow in an axial direction. A flow control vane is supported at a radial location between the inner and outer boundaries. A fluid discharge opening is provided for discharging a flow of the compressed fluid from a trailing edge of the vane, and a fluid control surface is provided adjacent to the fluid discharge opening and extends in the axial direction at the trailing edge of the vane. The fluid control surface has a curved trailing edge forming a Coanda surface. The fluid discharge opening is selectively provided with a compressed fluid to produce a Coanda effect along the control surface. The Coanda effect has a component in the radial direction effecting a turning of the fluid flow in the flow path radially inward or outward toward one of the inner and outer boundaries.

Secondary flow and heat transfer control in gas turbine inlet nozzle guide vanes

NASA Astrophysics Data System (ADS)

Burd, Steven Wayne

1998-12-01

Endwall heat transfer is a very serious problem in the inlet nozzle guide vane region of gas turbine engines. To resolve heat transfer concerns and provide the desired thermal protection, modern cooling flows for the vane endwalls tend to be excessive leading to lossy and inefficient designs. Coolant introduction is further complicated by the flow patterns along vane endwall surfaces. They are three-dimensional and dominated by strong, complex secondary flows. To achieve performance goals for next-generation engines, more aerodynamically efficient and advanced cooling concepts, including combustor bleed cooling, must be investigated. To this end, the overall performance characteristics of several combustor bleed flow designs are assessed in this experimental study. In particular, their contributions toward secondary flow control and component cooling are documented. Testing is performed in a large-scale, guide vane simulator comprised of three airfoils encased between one contoured and one flat endwall. Core flow is supplied to this simulator at an inlet chord Reynolds number of 350,000 and turbulence intensity of 9.5%. Combustor bleed cooling flow is injected through the contoured endwall via inclined slots. The slots vary in cross-sectional area, have equivalent slot widths, and are positioned with their leeward edges 10% of the axial chord ahead of the airfoil leading edges. Measurements with hot-wire anemometry characterize the inlet and exit flow fields of the cascade. Total and static pressure measurements document aerodynamic performance. Thermocouple measurements detail thermal fields and permit evaluation of surface adiabatic effectiveness. To elucidate the effects of bleed injection, data are compared to an experiment taken without bleed. The influence of bleed mass flow rate and slot geometry on the aerodynamic losses and thermal protection arc given. This study suggests that such combustor bleed flow cooling offers significant thermal protection without

CMC vane assembly apparatus and method

DOEpatents

Schiavo, Anthony L; Gonzalez, Malberto F; Huang, Kuangwei; Radonovich, David C

2012-10-23

A metal vane core or strut (64) is formed integrally with an outer backing plate (40). An inner backing plate (38) is formed separately. A spring (74) with holes (75) is installed in a peripheral spring chamber (76) on the strut. Inner and outer CMC shroud covers (46, 48) are formed, cured, then attached to facing surfaces of the inner and outer backing plates (38, 40). A CMC vane airfoil (22) is formed, cured, and slid over the strut (64). The spring (74) urges continuous contact between the strut (64) and airfoil (66), eliminating vibrations while allowing differential expansion. The inner end (88) of the strut is fastened to the inner backing plate (38). A cooling channel (68) in the strut is connected by holes (69) along the leading edge of the strut to peripheral cooling paths (70, 71) around the strut. Coolant flows through and around the strut, including through the spring holes.

Turbine stator vane segment having internal cooling circuits

DOEpatents

Jones, Raymond Joseph; Burns, James Lee; Bojappa, Parvangada Ganapathy; Jones, Schotsch Margaret

2003-01-01

A turbine stator vane includes outer and inner walls each having outer and inner chambers and a vane extending between the outer and inner walls. The vane includes first, second, third, fourth and fifth cavities for flowing a cooling medium. The cooling medium enters the outer chamber of the outer wall, flows through an impingement plate for impingement cooling of the outer band wall defining in part the hot gas path and through openings in the first, second and fourth cavities for flow radially inwardly, cooling the vane. The spent cooling medium flows into the inner wall and inner chamber for flow through an impingement plate radially outwardly to cool the inner wall. The spent cooling medium flows through the third cavity for egress from the turbine vane segment from the outer wall. The first, second or third cavities contain inserts having impingement openings for impingement cooling of the vane walls. The fifth cavity provides air cooling for the trailing edge.

Stagnation region gas film cooling for turbine blade leading edge applications

NASA Technical Reports Server (NTRS)

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

1976-01-01

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

Influence of vane sweep on rotor-stator interaction noise

NASA Technical Reports Server (NTRS)

Envia, Edmane; Kerschen, Edward J.

1990-01-01

The influence of vane sweep in rotor-stator interaction noise is investigated. In an analytical approach, the interaction of a convected gust representing the rotor viscous wake, with a cascade of cascade of finite span swept airfoils, representing the stator, is analyzed. The analysis is based on the solution of the exact linearized equations of motion. High frequency convected gusts for which noise generation is concentrated near the leading edge of airfoils is considered. In a preliminary study, the problem of an isolated finite span swept airfoil interacting with a convected gust is analyzed. Results indicate that sweep can substantially reduce the farfield noise levels for a single airfoil. Using the single airfoil model, an approximate solution to the problem of noise radiation from a cascade of finite span swept airfoils interacting with a convected gust is derived. A parametric study of noise generated by gust cascade interaction is carried out to assess the effectiveness of vane sweep in reducing rotor-stator interaction noise. The results show that sweep is beneficial in reducing noise levels. Rotor wake twist or circumferential lean substantially influences the effectiveness of vane sweep. The orientation of vane sweep must be chosen to enhance the natural phase lag caused by wake lean, in which case rather small sweep angles substantially reduce the noise levels.

Influence of vane sweep on rotor-stator interaction noise

NASA Astrophysics Data System (ADS)

Envia, Edmane; Kerschen, Edward J.

1990-12-01

The influence of vane sweep in rotor-stator interaction noise is investigated. In an analytical approach, the interaction of a convected gust representing the rotor viscous wake, with a cascade of cascade of finite span swept airfoils, representing the stator, is analyzed. The analysis is based on the solution of the exact linearized equations of motion. High frequency convected gusts for which noise generation is concentrated near the leading edge of airfoils is considered. In a preliminary study, the problem of an isolated finite span swept airfoil interacting with a convected gust is analyzed. Results indicate that sweep can substantially reduce the farfield noise levels for a single airfoil. Using the single airfoil model, an approximate solution to the problem of noise radiation from a cascade of finite span swept airfoils interacting with a convected gust is derived. A parametric study of noise generated by gust cascade interaction is carried out to assess the effectiveness of vane sweep in reducing rotor-stator interaction noise. The results show that sweep is beneficial in reducing noise levels. Rotor wake twist or circumferential lean substantially influences the effectiveness of vane sweep. The orientation of vane sweep must be chosen to enhance the natural phase lag caused by wake lean, in which case rather small sweep angles substantially reduce the noise levels.

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.

Numerical Modeling of Thermal Edge Flow

NASA Astrophysics Data System (ADS)

Ibrayeva, Aizhan

A gas flow can be induced between two interdigitated arrays of thin vanes, when one of the arrays is uniformly heated or cooled. Sharply curved isotherms near the vane edges leads to momentum imbalance among incident particles, which creates Knudsen force to the vane and thermal edge flow in a gas. The flow is observed in a rarefied gas, when the mean free path of the molecules are comparable with the characteristic length scale of the system. In order to understand a physical mechanism of the flow and Knudsen force, the configuration was numerically investigated under different gas rarefication degrees and temperature gradients in the system by direct simulation Monte Carlo (DSMC) method. From simulations, the highest force value is obtained when Knudsen number is around 0.5 and becomes negligible in free molecular and continuum regimes. DSMC results are analyzed from the theoretical point of view and compared to experimental data. Validation of the simulations is done by the RKDG method. An effect of various geometric parameters to the performance of the actuator was investigated and suggestions were made for improved design of the device.

Comparison of Predicted and Measured Turbine Vane Rough Surface Heat Transfer

NASA Technical Reports Server (NTRS)

Boyle, R. J.; Spuckler, C. M.; Lucci, B. L.

2000-01-01

The proposed paper compares predicted turbine vane heat transfer for a rough surface over a wide range of test conditions with experimental data. Predictions were made for the entire vane surface. However, measurements were made only over the suction surface of the vane, and the leading edge region of the pressure surface. Comparisons are shown for a wide range of test conditions. Inlet pressures varied between 3 and 15 psia, and exit Mach numbers ranged between 0.3 and 0.9. Thus, while a single roughened vane was used for the tests, the effective rougness,(k(sup +)), varied by more than a factor of ten. Results were obtained for freestream turbulence levels of 1 and 10%. Heat transfer predictions were obtained using the Navier-Stokes computer code RVCQ3D. Two turbulence models, suitable for rough surface analysis, are incorporated in this code. The Cebeci-Chang roughness model is part of the algebraic turbulence model. The k-omega turbulence model accounts for the effect of roughness in the application of the boundary condition. Roughness causes turbulent flow over the vane surface. Even after accounting for transition, surface roughness significantly increased heat transfer compared to a smooth surface. The k-omega results agreed better with the data than the Cebeci-Chang model. However, the low Reynolds number k-omega model did not accurately account for roughness when the freestream turbulence level was low. The high Reynolds number version of this model was more suitable when the freestream turbulence was low.

Hybrid flow control of a transport truck side-mirror using AC-DBD plasma actuated guide vane

NASA Astrophysics Data System (ADS)

Michelis, Theodoros; Kotsonis, Marios

2014-11-01

A wind-tunnel study is conducted towards hybrid flow control of a full-scale transport truck side-mirror (Re = 4 ×105) . The mirror is mounted on a structure that models the truck cabin. PIV measurements are performed at a range of velocities from 15 to 25 m/s and from leeward to windward angles of -5° to +5° . A slim guide vane of 6cm chord is employed along the span of the hub of the mirror for redirecting high momentum flow towards the wake region. Separation from the leading edge of the guide vane is reduced or eliminated by means of AC-DBD plasma actuator, operating at voltage of 35 kV peak-to-peak and frequency of 200 Hz. Time-averaged velocity fields are obtained at the centre of the mirror for three scenarios: a) reference case lacking any control elements; b) guide vane only and c) combination of the guide vane and the AC-DBD. The comparison of cases demonstrates that at 25 m/s windward conditions (-5°) the guide vane is capable of increasing momentum (+20%) in the wake of the mirror with additional improvement when plasma actuation is applied (+21%). In contrast, at leeward conditions (+5°) , the guide vane reduces momentum (-20%), though with actuation an increase is observed (+5%). Total recovered momentum is 25%.

Aerodynamic performance of high turning core turbine vanes in a two dimensional cascade

NASA Technical Reports Server (NTRS)

Schwab, J. R.

1982-01-01

Experimental and theoretical aerodynamic performance data are presented for four uncooled high turning core turbine vanes with exit angles of 74.9, 75.0, 77.5, and 79.6 degrees in a two dimensional cascade. Data for a more conservative 67.0 degree vane are included for comparison. Correction of the experimental aftermix kinetic energy losses to a common 0.100 centimeter trailing edge thickness yields a linear trend of increased loss from 0.020 to 0.025 as the vane exit angle increases from 67.0 to 79.6 degrees. The theoretical losses show a similar trend. The experimental and theoretical vane surface velocity distributions generally agree within approximately five percent, although the suction surface theoretical velocities are generally higher than the experimental velocities as the vane exit angle increases.

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.

Characterization of Ceramic Matrix Composite Vane Subelements Subjected to Rig Testing in a Gas Turbine Environment

NASA Technical Reports Server (NTRS)

Verrilli, Michael; Calomino, Anthony; Thomas, David J.; Robinson, R. Craig

2004-01-01

Vane subelements were fabricated from a silicon carbide fiber-reinforced silicon carbide matrix (SiC/SiC) composite. A cross-sectional slice of an aircraft engine metal vane was the basis of the vane subelement geometry. To fabricate the small radius of the vane's trailing edge using stiff Sylramic SiC fibers, a unique SiC fiber architecture was developed. A test configuration for the vanes in a high pressure gas turbine environment was designed and fabricated. Testing was conducted using a pressure of 6 atm and combustion flow rate of 0.5 kg/sec, and consisted of fifty hours of steady state operation followed by 102 2-minute thermal cycles. A surface temperature of 1320 C was obtained for the EBC-coated SiC/SiC vane subelement. This paper will briefly discuss the vane fabrication, test configuration, and results of the vane testing. The emphasis of the paper is on characterization of the post-test condition of the vanes.

A Three-Dimensional Coupled Internal/External Simulation of a Film-Cooled Turbine Vane

NASA Technical Reports Server (NTRS)

Heidmann, James D.; Rigby, David L.; Ameri, Ali A.

1999-01-01

A three-dimensional Navier-Stokes simulation has been performed for a realistic film-cooled turbine vane using the LeRC-HT code. The simulation includes the flow regions inside the coolant plena and film cooling holes in addition to the external flow. The vane is the subject of an upcoming NASA Glenn Research Center experiment and has both circular cross-section and shaped film cooling holes. This complex geometry is modeled using a multi-block grid which accurately discretizes the actual vane geometry including shaped holes. The simulation matches operating conditions for the planned experiment and assumes periodicity in the spanwise direction on the scale of one pitch of the film cooling hole pattern. Two computations were performed for different isothermal wall temperatures, allowing independent determination of heat transfer coefficients and film effectiveness values. The results indicate separate localized regions of high heat transfer coefficient values, while the shaped holes provide a reduction in heat flux through both parameters. Hole exit data indicate rather simple skewed profiles for the round holes, but complex profiles for the shaped holes with mass fluxes skewed strongly toward their leading edges.

Leading-edge vortex lifts swifts.

PubMed

Videler, J J; Stamhuis, E J; Povel, G D E

2004-12-10

The current understanding of how birds fly must be revised, because birds use their hand-wings in an unconventional way to generate lift and drag. Physical models of a common swift wing in gliding posture with a 60 degrees sweep of the sharp hand-wing leading edge were tested in a water tunnel. Interactions with the flow were measured quantitatively with digital particle image velocimetry at Reynolds numbers realistic for the gliding flight of a swift between 3750 and 37,500. The results show that gliding swifts can generate stable leading-edge vortices at small (5 degrees to 10 degrees) angles of attack. We suggest that the flow around the arm-wings of most birds can remain conventionally attached, whereas the swept-back hand-wings generate lift with leading-edge vortices.

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.

Sliding vane geometry turbines

DOEpatents

Sun, Harold Huimin; Zhang, Jizhong; Hu, Liangjun; Hanna, Dave R

2014-12-30

Various systems and methods are described for a variable geometry turbine. In one example, a turbine nozzle comprises a central axis and a nozzle vane. The nozzle vane includes a stationary vane and a sliding vane. The sliding vane is positioned to slide in a direction substantially tangent to an inner circumference of the turbine nozzle and in contact with the stationary vane.

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.

Reduction of background noise induced by wind tunnel jet exit vanes

NASA Technical Reports Server (NTRS)

Martin, R. M.; Brooks, T. F.; Hoad, D. R.

1985-01-01

The NASA-Langley 4 x 7 m wind tunnel develops low frequency flow pulsations at certain velocity ranges during open throat mode operation, affecting the aerodynamics of the flow and degrading the resulting model test data. Triangular vanes attached to the trailing edge of flat steel rails, mounted 10 cm from the inside of the jet exit walls, have been used to reduce this effect; attention is presently given to methods used to reduce the inherent noise generation of the vanes while retaining their pulsation reduction features.

Leading edge gypsy moth population dynamics

Treesearch

M. R. Carter; F. W. Ravlin; M. L. McManus

1991-01-01

Leading edge gypsy moth populations have been the focus of several intervention programs (MDIPM, AIPM). Knowledge of gypsy moth population dynamics in leading edge area is crucial for effective management. Populations in these areas tend to reach outbreak levels (noticeable defoliation) within three to four years after egg masses are first detected. Pheromone traps...

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.

Phase 2: HGM air flow tests in support of HEX vane investigation

NASA Technical Reports Server (NTRS)

Cox, G. B., Jr.; Steele, L. L.; Eisenhart, D. W.

1993-01-01

Following the start of SSME certification testing for the Pratt and Whitney Alternate Turbopump Development (ATD) High Pressure Oxidizer Turbopump (HPOTP), cracking of the leading edge of the inner HEX vane was experienced. The HEX vane, at the inlet of the oxidizer bowl in the Hot Gas Manifold (HGM), accepts the HPOTP turbine discharge flow and turns it toward the Gaseous Oxidizer Heat Exchanger (GOX HEX) coil. The cracking consistently initiated over a specific circumferential region of the hex vane, with other circumferential locations appearing with increased run time. Since cracking had not to date been seen with the baseline HPOTP, a fluid-structural interaction involving the ATD HPOTP turbine exit flowfield and the HEX inner vane was suspected. As part of NASA contract NAS8-36801, Pratt and Whitney conducted air flow tests of the ATD HPOTP turbine turnaround duct flowpath in the MSFC Phase 2 HGM air flow model. These tests included HEX vane strain gages and additional fluctuating pressure gages in the turnaround duct and HEX vane flowpath area. Three-dimensional flow probe measurements at two stations downstream of the turbine simulator exit plane were also made. Modifications to the HPOTP turbine simulator investigated the effects on turbine exit flow profile and velocity components, with the objective of reproducing flow conditions calculated for the actual ATD HPOTP hardware. Testing was done at the MSFC SSME Dynamic Fluid Air Flow (Dual-Leg) Facility, at air supply pressures between 50 and 250 psia. Combinations of turbine exit Mach number and pressure level were run to investigate the effect of flow regime. Information presented includes: (1) Descriptions of turbine simulator modifications to produce the desired flow environment; (2) Types and locations for instrumentation added to the flow model for improved diagnostic capability; (3) Evaluation of the effect of changes to the turbine simulator flowpath on the turbine exit flow environment; and (4

Cold-air annular-cascade investigation of aerodynamic performance of cooled turbine vanes. 2: Trailing-edge ejection, film cooling, and transpiration cooling

NASA Technical Reports Server (NTRS)

Goldman, L. J.; Mclallin, K. L.

1975-01-01

The aerodynamic performance of four different cooled vane configurations was experimentally determined in a full-annular cascade at a primary- to coolant-total-temperature ratio of 1.0. The vanes were tested over a range of coolant flow rates and pressure ratios. Overall vane efficiencies were obtained and compared, where possible, with the results obtained in a four-vane, annular-sector cascade. The vane efficiency and exit flow conditions as functions of radial position were also determined and compared with solid (uncooled) vane results.

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.

The Effect of Break Edge Configuration on the Aerodynamics of Anti-Ice Jet Flow

NASA Astrophysics Data System (ADS)

Tatar, V.; Yildizay, H.; Aras, H.

2015-05-01

One of the components of a turboprop gas turbine engine is the Front Bearing Structure (FBS) which leads air into the compressor. FBS directly encounters with ambient air, as a consequence ice accretion may occur on its static vanes. There are several aerodynamic parameters which should be considered in the design of anti-icing system of FBS, such as diameter, position, exit angle of discharge holes, etc. This research focuses on the effects of break edge configuration over anti-ice jet flow. Break edge operation is a process which is applied to the hole in order to avoid sharp edges which cause high stress concentration. Numerical analyses and flow visualization test have been conducted. Four different break edge configurations were used for this investigation; without break edge, 0.35xD, 74xD, 0.87xD. Three mainstream flow conditions at the inlet of the channel are defined; 10m/s, 20 m/s and 40 m/s. Shear stresses are extracted from numerical analyses near the trailing edge of pressure surface where ice may occur under icing conditions. A specific flow visualization method was used for the experimental study. Vane surface near the trailing edge was dyed and thinner was injected into anti-ice jet flow in order to remove dye from the vane surface. Hence, film effect on the surface could be computed for each testing condition. Thickness of the dye removal area of each case was examined. The results show noticeable effects of break edge operation on jet flow, and the air film effectiveness decreases when mainstream inlet velocity decreases.

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.

Development of a design model for airfoil leading edge film cooling

NASA Astrophysics Data System (ADS)

Wadia, A. R.; Nealy, D. A.

1985-03-01

A series of experiments on scaled cylinder models having injection through holes inclined at 20, 30, 45, and 90 degrees are presented. The experiments were conducted in a wind tunnel on several stainless steel test specimens in which flow and heat transfer parameters were measured over simulated airfoil leading edge surfaces. On the basis of the experimental results, an engineering design model is proposed that treats the gas-to-surface heat transfer coefficient with film cooling in a manner suggested by Luckey and L'Ecuyer (1981). It is shown that the main factor influencing the averaged film cooling effectiveness in the showerhead region is the inclination of the injection holes. The effectiveness parameter was not affected by variations in the coolant-to-gas stream pressure ratio, the freestream Mach number, the gas to coolant temperature ratio, or the gas stream Reynolds number. Experience in the wind tunnel tests is reflected in the design of the model in which the coolant side heat transfer coefficient is offset by a simultaneous increase in the gas side film coefficient. The design applications of the analytical model are discussed, with emphasis given to high temperature first stage turbine vanes and rotor blades.

Dynamic Stall Characteristics of Drooped Leading Edge Airfoils

NASA Technical Reports Server (NTRS)

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

2000-01-01

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

Symmetric airfoil geometry effects on leading edge noise.

PubMed

Gill, James; Zhang, X; Joseph, P

2013-10-01

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

Wing Leading Edge Debris Analysis

NASA Technical Reports Server (NTRS)

Shah, Sandeep; Jerman, Gregory

2004-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-04-01

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

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.

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.

Numerical and Experimental Study of a Cooling for Vanes in a Small Turbine Engine

NASA Astrophysics Data System (ADS)

Šimák, Jan; Michálek, Jan

2016-03-01

This paper is concerned with a cooling system for inlet guide vanes of a small turbine engine which are exposed to a high temperature gas leaving a combustion chamber. Because of small dimensions of the vanes, only a simple internal cavity and cooling holes can be realized. The idea was to utilize a film cooling technique. The proposed solution was simulated by means of a numerical method based on a coupling of CFD and heat transfer solvers. The numerical results of various scenarios (different coolant temperature, heat transfer to surroundings) showed a desired decrease of the temperature, especially on the most critical part - the trailing edge. The numerical data are compared to results obtained by experimental measurements performed in a test facility in our institute. A quarter segment model of the inlet guide vanes wheel was equipped with thermocouples in order to verify an effect of cooling. Despite some uncertainty in the results, a verifiable decrease of the vane temperature was observed.

Engineering diagnostics for vortex-induced stay vanes cracks in a Francis turbine

NASA Astrophysics Data System (ADS)

D'Agostini Neto, Alexandre; Gissoni, Humberto, Dr.; Gonçalves, Manuel, Dr.; Cardoso, Rogério; Jung, Alexander, Dr.; Meneghini, Julio, Prof.

2016-11-01

Despite the fact that vortex-induced vibration (VIV) in hydraulic turbines components (especially in stay vanes) is a well-known phenomenon, it still remains challenging for operation and maintenance teams in several power plants around the world. Since the first publication of a similar problem in 1967, literature shows that at least 27 other turbines witnessed strong stay vane vibrations associated with vortex shedding. Recurrent stay vane cracks in a 250 MW Francis turbine in Brazil motivated an engineering study involving prototype measurements, structural and Computational Fluid Dynamics (CFD) analysis in order to determine a proper geometry modification that could eliminate the periodic vortex wake generated at the stay vanes trailing edge. First cracks appeared in 1978 just after the machine was put into operation. A study published in 1982 associated these cracks with dynamic excitations caused by the water flow at high flow conditions. New stay vane profiles were proposed and executed as well as improved welding recommendations. Cracks however, continued to appear requiring welding repairs roughly every two years. Although Voith Hydro was not the original equipment manufacturer for these units, the necessary information was available to study the issue and propose and execute new stay vane profiles. This paper details the approach taken for the study. First, indirect vibration measurements were used to determine vibration frequencies to help to characterize the affected mode shapes. These results were compared to finite element (FE) calculations. Strain gage measurements performed afterwards confirmed the conclusions of this analysis. Next, transient CFD calculations were run to reproduce the measured phenomenon and to serve as a basis for a new stay vane geometry. This modification was then implemented in the actual turbine stay vanes. A new set of indirect vibration measurements indicated the effectiveness of the proposed solution. Final confirmation

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.

Partially turbulated trailing edge cooling passages for gas turbine nozzles

DOEpatents

Thatcher, Jonathan Carl; Burdgick, Steven Sebastian

2001-01-01

A plurality of passages are spaced one from the other along the length of a trailing edge of a nozzle vane in a gas turbine. The passages lie in communication with a cavity in the vane for flowing cooling air from the cavity through the passages through the tip of the trailing edge into the hot gas path. Each passage is partially turbulated and includes ribs in an aft portion thereof to provide enhanced cooling effects adjacent the tip of the trailing edge. The major portions of the passages are smooth bore. By this arrangement, reduced temperature gradients across the trailing edge metal are provided. Additionally, the inlets to each of the passages have a restriction whereby a reduced magnitude of compressor bleed discharge air is utilized for trailing edge cooling purposes.

Span efficiency of wings with leading edge protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2013-11-01

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

Heat-Pipe-Cooled Leading Edges for Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Glass, David E.

2006-01-01

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

Cooling circuit for steam and air-cooled turbine nozzle stage

DOEpatents

Itzel, Gary Michael; Yu, Yufeng

2002-01-01

The turbine vane segment includes inner and outer walls with a vane extending therebetween. The vane includes leading and trailing edge cavities and intermediate cavities. An impingement plate is spaced from the outer wall to impingement-cool the outer wall. Post-impingement cooling air flows through holes in the outer wall to form a thin air-cooling film along the outer wall. Cooling air is supplied an insert sleeve with openings in the leading edge cavity for impingement-cooling the leading edge. Holes through the leading edge afford thin-film cooling about the leading edge. Cooling air is provided the trailing edge cavity and passes through holes in the side walls of the vane for thin-film cooling of the trailing edge. Steam flows through a pair of intermediate cavities for impingement-cooling of the side walls. Post-impingement steam flows to the inner wall for impingement-cooling of the inner wall and returns the post-impingement cooling steam through inserts in other intermediate cavities for impingement-cooling the side walls of the vane.

61. DETAIL OF HEAD, VANE STEM, VANE HINGE, AND WHEEL ...

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

61. DETAIL OF HEAD, VANE STEM, VANE HINGE, AND WHEEL OF AN ELI WINDMILL ON THE GROUND AT THE STOLL RESIDENCE ABOUT 1-1/2 MILES WEST OF NEBRASKA CITY ON STEAM WAGON ROAD. - Kregel Windmill Company Factory, 1416 Central Avenue, Nebraska City, Otoe County, NE

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.

Laminar Flow Control Leading Edge Systems in Simulated Airline Service

NASA Technical Reports Server (NTRS)

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

1988-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

Effects of blade-vane ratio and rotor-stator spacing of fan noise with forward velocity

NASA Astrophysics Data System (ADS)

Woodward, R. P.; Glaser, F. W.

1981-08-01

A research fan stage was acoustically tested in an anechoic wind tunnel with a 41 m/sec tunnel flow. Two stator vane numbers giving cut-on and cut-off conditions were tested at three rotor-stator spacings ranging from about 0.5 to 2.0 rotor chords. These two stators were designed for similar aerodynamic performance. Hot film anemometer turbulence measurements were made at the leading edge of the stator for each spacing. The cut-off criterion strongly controlled the fundamental tone level at all spacings. The trends with spacing of the wake defect upwash component at the stator tip showed good agreement with the corresponding cut-on acoustic tone levels.

Inverse design of centrifugal compressor vaned diffusers in inlet shear flows

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

Zangeneh, M.

1996-04-01

A three-dimensional inverse design method in which the blade (or vane) geometry is designed for specified distributions of circulation and blade thickness is applied to the design of centrifugal compressor vaned diffusers. Two generic diffusers are designed, one with uniform inlet flow (equivalent to a conventional design) and the other with a sheared inlet flow. The inlet shear flow effects are modeled in the design method by using the so-called ``Secondary Flow Approximation`` in which the Bernoulli surfaces are convected by the tangentially mean inviscid flow field. The difference between the vane geometry of the uniform inlet flow and nonuniformmore » inlet flow diffusers is found to be most significant from 50 percent chord to the trailing edge region. The flows through both diffusers are computed by using Denton`s three-dimensional inviscid Euler solver and Dawes` three-dimensional Navier-Stokes solver under sheared in-flow conditions. The predictions indicate improved pressure recovery and internal flow field for the diffuser designed for shear inlet flow conditions.« less

Leading-edge receptivity for blunt-nose bodies

NASA Technical Reports Server (NTRS)

Kerschen, Edward J.

1991-01-01

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

Wing Leading Edge Concepts for Noise Reduction

NASA Technical Reports Server (NTRS)

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

2010-01-01

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

Gas turbine vane platform element

DOEpatents

Campbell, Christian X [Oviedo, FL; Schiavo, Anthony L [Oviedo, FL; Morrison, Jay A [Oviedo, FL

2012-08-28

A gas turbine CMC shroud plate (48A) with a vane-receiving opening (79) that matches a cross-section profile of a turbine vane airfoil (22). The shroud plate (48A) has first and second curved circumferential sides (73A, 74A) that generally follow the curves of respective first and second curved sides (81, 82) of the vane-receiving opening. Walls (75A, 76A, 77A, 78A, 80, 88) extend perpendicularly from the shroud plate forming a cross-bracing structure for the shroud plate. A vane (22) may be attached to the shroud plate by pins (83) or by hoop-tension rings (106) that clamp tabs (103) of the shroud plate against bosses (105) of the vane. A circular array (20) of shroud plates (48A) may be assembled to form a vane shroud ring in which adjacent shroud plates are separated by compressible ceramic seals (93).

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

NASA Technical Reports Server (NTRS)

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

2016-01-01

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

The mechanism performance of improved oil pump with micro-structured vanes

NASA Astrophysics Data System (ADS)

Li, Ping; Xie, Jin; Qi, Dongtao; Li, Houbu

2017-09-01

The wear of oil pump vanes easily leads to the noise and vibration, even results the decrease of volume efficiency and total efficiency. In order to reduce the friction and improve the lubrication between the vane and the pump inner wall, the micro-machining of micro-structure on the oil pump vanes is proposed. First, the micro-V-grooves with the depth ranging from 500μm to 50μm were micro-grinding on the top of the vanes by a diamond grinding wheel. Secondly, the experiments were conducted to test the actual flow rate, the output power and the overall efficiency of the oil pump with and without the micro-groove vanes. Then, the computational fluid dynamics (CFD) method was adopted to simulate the pump internal flow field. Finally, the micro-flow field between the internal wall of the oil pump and the top of micro-grooved vanes was analyzed. The results shows that the pump overall efficiency increased as the decrease of micro-groove depth from 500 μm to 50μm and not be affected by the rotate speed and working frequency of the pump rotator. Especially the micro-groove with depth of 50μm, the actual flow rate, the output power and the overall efficiency reached to the maximum. From CFD simulation, the velocity of the micro-flow between the surfaces of the vane and inner wall was larger than the pump linear velocity when the microstructure depth is larger than 50μm, leading to an internal leakage. When the micro-groove depth is between10-50μm, the velocity of the micro-flow was less than the pump linear velocity and no internal leakage was found, but the oil film thickness is too small to be beneficial to lubrication according to the fluid dynamic characteristics. Thus, for the oil pump equipping with micro-grooved vane with the depth of 50 μm, the internal leakage not only is avoided but the lubrication efficiency is improved and the oil pump efficiency is also enhanced.

The leading-edge vortex of yacht sails

NASA Astrophysics Data System (ADS)

Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

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

Structural Health Monitoring Analysis for the Orbiter Wing Leading Edge

NASA Technical Reports Server (NTRS)

Yap, Keng C.

2010-01-01

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

Turning Vanes inside the Altitude Wind Tunnel

NASA Image and Video Library

1944-02-21

Men stand in front of turning vanes inside the Altitude Wind Tunnel (AWT) at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The AWT was the only wind tunnel capable of testing full-size aircraft engines in simulated altitude conditions. A large wooden drive fan, located on the other side of these vanes, created wind speeds up to 500 miles per hour. The drive shaft connected the fan to the induction motor located in an adjacent building. Turning vanes were located in each corner of the rectangular tunnel to straighten the airflow and direct it around the corners. This set of vanes was located in the 31-foot-diameter southeast corner of the tunnel. These elliptical panels consisted of 36 to 42 vertical vanes that were supported by three horizontal supports. The individual vanes were 2.5 feet long and half-moon shaped. The panel of vanes was affixed to the curved corner rings of the tunnel. Each set of turning vanes had a moveable vane in the middle of the lower level for personnel access. Each set of vanes took weeks to assemble before they were installed during the summer of 1943. This publicity photograph was taken just weeks after the tunnel became operational in February 1944.

The Flow Field on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2008-11-01

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

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

NASA Technical Reports Server (NTRS)

Middleton, W. D.

1982-01-01

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

High-Frequency Testing of Composite Fan Vanes With Erosion-Resistant Coating Conducted

NASA Technical Reports Server (NTRS)

Bowman, Cheryl L.; Sutter, James K.; Naik, Subhash; Otten, Kim D.; Perusek, Gail P.

2003-01-01

The mechanical integrity of hard, erosion-resistant coatings were tested using the Structural Dynamics Laboratory at the NASA Glenn Research Center. Under the guidance of Structural Mechanics and Dynamics Branch personnel, fixturing and test procedures were developed at Glenn to simulate engine vibratory conditions on coated polymer-matrix- composite bypass vanes using a slip table in the Structural Dynamics Laboratory. Results from the high-frequency mechanical bench testing, along with concurrent erosion testing of coupons and vanes, provided sufficient confidence to engine-endurance test similarly coated vane segments. The knowledge gained from this program will be applied to the development of oxidation- and erosion-resistant coatings for polymer matrix composite blades and vanes in future advanced turbine engines. Fan bypass vanes from the AE3007 (Rolls Royce America, Indianapolis, IN) gas turbine engine were coated by Engelhard (Windsor, CT) with compliant bond coatings and hard ceramic coatings. The coatings were developed collaboratively by Glenn and Allison Advanced Development Corporation (AADC)/Rolls Royce America through research sponsored by the High-Temperature Engine Materials Technology Project (HITEMP) and the Higher Operating Temperature Propulsion Components (HOTPC) project. High-cycle fatigue was performed through high-frequency vibratory testing on a shaker table. Vane resonant frequency modes were surveyed from 50 to 3000 Hz at input loads from 1g to 55g on both uncoated production vanes and vanes with the erosion-resistant coating. Vanes were instrumented with both lightweight accelerometers and strain gauges to establish resonance, mode shape, and strain amplitudes. Two high-frequency dwell conditions were chosen to excite two strain levels: one approaching the vane's maximum allowable design strain and another near the expected maximum strain during engine operation. Six specimens were tested per dwell condition. Pretest and posttest

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.

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

NASA Astrophysics Data System (ADS)

Modlin, James Michael

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

Thermal management of tungsten leading edges in DIII-D

DOE PAGES

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

2017-04-29

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

Thermal management of tungsten leading edges in DIII-D

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

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

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

Method of joining a vane cavity insert to a nozzle segment of a gas turbine

DOEpatents

Burdgick, Steven Sebastian

2002-01-01

An insert containing apertures for impingement cooling a nozzle vane of a nozzle segment in a gas turbine is inserted into one end of the vane. The leading end of the insert is positioned slightly past a rib adjacent the opposite end of the vane through which the insert is inserted. The end of the insert is formed or swaged into conformance with the inner margin of the rib. The insert is then brazed or welded to the rib.

Hook nozzle arrangement for supporting airfoil vanes

DOEpatents

Shaffer, James E.; Norton, Paul F.

1996-01-01

A gas turbine engine's nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic.

Hook nozzle arrangement for supporting airfoil vanes

DOEpatents

Shaffer, J.E.; Norton, P.F.

1996-02-20

A gas turbine engine`s nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic. 8 figs.

Vortex interaction with a leading-edge of finite thickness

NASA Technical Reports Server (NTRS)

Sohn, D.; Rockwell, Donald

1987-01-01

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

Theory of step on leading edge of negative corona current pulse

NASA Astrophysics Data System (ADS)

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

2000-03-01

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

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

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

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

1980-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1987-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1959-01-01

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

Fan Noise Source Diagnostic Test: Vane Unsteady Pressure Results

NASA Technical Reports Server (NTRS)

Envia, Edmane

2002-01-01

To investigate the nature of fan outlet guide vane pressure fluctuations and their link to rotor-stator interaction noise, time histories of vane fluctuating pressures were digitally acquired as part of the Fan Noise Source Diagnostic Test. Vane unsteady pressures were measured at seven fan tip speeds for both a radial and a swept vane configuration. Using time-domain averaging and spectral analysis, the blade passing frequency (BPF) harmonic and broadband contents of the vane pressures were individually analyzed. Significant Sound Pressure Level (SPL) reductions were observed for the swept vane relative to the radial vane for the BPF harmonics of vane pressure, but vane broadband reductions due to sweep turned out to be much smaller especially on an average basis. Cross-correlation analysis was used to establish the level of spatial coherence of broadband pressures between different locations on the vane and integral length scales of pressure fluctuations were estimated from these correlations. Two main results of this work are: (1) the average broadband level on the vane (in dB) increases linearly with the fan tip speed for both the radial and swept vanes, and (2) the broadband pressure distribution on the vane is nearly homogeneous and its integral length scale is a monotonically decreasing function of fan tip speed.

Bed erosion control at 60 degree river confluence using vanes

NASA Astrophysics Data System (ADS)

Wuppukondur, Ananth; Chandra, Venu

2017-04-01

, for the same discharge ratios (q), Rmax reduces to 0.64 and 0.72 with 3h spacing between the vanes, respectively. Obstruction to the flow increases with an increase of vane angle which leads to decrease of bed erosion. Also, the bed erosion increases with an increase of spacing between the vanes. Hence, vanes placed at 600 vane angle and 2h spacing exhibit better performance.

Inner vane fringes of barn owl feathers reconsidered: morphometric data and functional aspects

PubMed Central

Bachmann, Thomas; Wagner, Hermann; Tropea, Cameron

2012-01-01

It is a challenge to understand how barn owls (Tyto alba) reduce noise during flight to be able to hunt small mammals by audition. Several specializations of the wing and the wing feathers have been implicated in noise reduction. What has been overlooked so far are the fringes at the inner vanes of remiges. We demonstrated, by using precise imaging techniques combined with morphometric measurements and air-flow studies, that these fringes merge into neighboring feather vanes by gliding into the grooves at the lower wing surface that are formed by parallel-oriented barb shafts. The connection of adjacent feathers results in a smooth lower wing surface and thus reduces sharp and noisy edges. This finding sheds new light on the mechanisms underlying noise reduction of flying owls. PMID:22471670

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

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2017-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2017-06-01

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

Experimental Investigation on Design Enhancement of Axial Fan Using Fixed Guide Vane

NASA Astrophysics Data System (ADS)

Munisamy, K. M.; Govindasamy, R.; Thangaraju, S. K.

2015-09-01

Airflow passes through the rotating blade in an axial flow fan will experience a helical flow pattern. This swirling effect leads the system to experience swirl energy losses or pressure drop yet reducing the total efficiency of the fan system. A robust tool to encounter this air spin past the blade is by introducing guide vane to the system. Owing to its importance, a new approach in designing outlet guide vane design for a commercial usage 1250mm diameter axial fan with a 30° pitch angle impeller has been introduced in this paper. A single line metal of proper curvature guide vane design technique has been adopted for this study. By choosing fan total efficiency as a target variable to be improved, the total and static pressure on the design point were set to be constraints. Therefore, the guide vane design was done based on the improvement target on the static pressure in system. The research shows that, with the improvement in static pressure by 29.63% through guide vane installation, the total fan efficiency is increased by 5.12%, thus reduces the fan power by 5.32%. Good agreement were found, that when the fan total efficiency increases, the power consumption of the fan is reduced. Therefore, this new approach of guide vane design can be applied to improve axial fan performance.

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.

Variable area nozzle including a plurality of convexly vanes with a crowned contour, in a vane to vane sealing arrangement and with nonuniform lengths

NASA Technical Reports Server (NTRS)

Arnold, Steven M. (Inventor); Penney, Nicholas (Inventor)

2008-01-01

A variable area nozzle comprising a concentric support and a plurality of convexly contoured self sealing vanes is disclosed and claimed. The vanes are circumferentially and rotatably mounted to the concentric support forming a nozzle infinitely positionable between a first position corresponding to a minimum area nozzle and a second position corresponding to a maximum area nozzle. A closer, which is preferably a shape memory alloy (SMA), urges the nozzle toward the first position corresponding to a minimum area nozzle. Periodically spaced openers act between adjacent vanes to urge the nozzle to a second position corresponding to a maximum area nozzle.

Leading edge flap system for aircraft control augmentation

NASA Technical Reports Server (NTRS)

Rao, D. M. (Inventor)

1984-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1989-01-01

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

Radial inlet guide vanes for a combustor

DOEpatents

Zuo, Baifang; Simons, Derrick; York, William; Ziminsky, Willy S

2013-02-12

A combustor may include an interior flow path therethrough, a number of fuel nozzles in communication with the interior flow path, and an inlet guide vane system positioned about the interior flow path to create a swirled flow therein. The inlet guide vane system may include a number of windows positioned circumferentially around the fuel nozzles. The inlet guide vane system may also include a number of inlet guide vanes positioned circumferentially around the fuel nozzles and adjacent to the windows to create a swirled flow within the interior flow path.

Methods and Apparatus for Deployable Swirl Vanes

NASA Technical Reports Server (NTRS)

Shah, Parthiv N. (Inventor)

2017-01-01

An aircraft control structure for drag management includes a nozzle structure configured to exhaust a swirling fluid stream. A plurality of swirl vanes are positioned within the nozzle structure, and an actuation subsystem is configured to cause the plurality of swirl vanes to move from a deployed state to a non-deployed state. In the non-deployed state, the plurality of swirl vanes are substantially flush with the inner surface of the nozzle structure. In the deployed state, the plurality of swirl vanes produce the swirling fluid stream.

Vane segment support and alignment device

DOEpatents

McLaurin, Leroy Dixon; Sizemore, John Derek

1999-01-01

A support and alignment assembly for supporting and aligning a vane segment is provided. The support and alignment assembly comprises a torque plate which defines an opening for receiving an eccentric pin and a locking end member for receiving a lock socket member. An eccentric pin adjustably supported by the torque plate opening for supporting and aligning a vane segment is provided. A lock socket member adapted to securely receive the eccentric pin and rotated therewith, and adjustably engage the torque plate locking end is provided. The lock socket member receives the eccentric pin, such that when the eccentric pin is adjusted to align the vane segment, the lock socket member engages the torque plate locking end to secure the vane segment in the desired position.

Design Concepts for Cooled Ceramic Matrix Composite Turbine Vanes

NASA Technical Reports Server (NTRS)

Boyle, Robert

2014-01-01

This project demonstrated that higher temperature capabilities of ceramic matrix composites (CMCs) can be used to reduce emissions and improve fuel consumption in gas turbine engines. The work involved closely coupling aerothermal and structural analyses for the first-stage vane of a high-pressure turbine (HPT). These vanes are actively cooled, typically using film cooling. Ceramic materials have structural and thermal properties different from conventional metals used for the first-stage HPT vane. This project identified vane configurations that satisfy CMC structural strength and life constraints while maintaining vane aerodynamic efficiency and reducing vane cooling to improve engine performance and reduce emissions. The project examined modifications to vane internal configurations to achieve the desired objectives. Thermal and pressure stresses are equally important, and both were analyzed using an ANSYS® structural analysis. Three-dimensional fluid and heat transfer analyses were used to determine vane aerodynamic performance and heat load distributions.

Low transient thermal stress turbine engine components

DOEpatents

Shi, Jun [Glastonbury, CT; Schmidt, Wayde R [Pomfret Center, CT

2011-06-28

A turbine vane includes a platform; and at least one airfoil mounted to the platform and having a trailing edge and a leading edge, wherein the vane is composed of a functionally graded material having a first material and a second material, wherein the trailing edge includes a greater amount of the first material than the second material, and the leading edge includes a greater amount of the second material than the first material.

Broadband Noise Reduction of a Low-Speed Fan Noise Using Trailing Edge Blowing

NASA Technical Reports Server (NTRS)

Sutliff, Daniel L.

2005-01-01

An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through the use of rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a continuous trailing edge slot. Hollow blades with interior guide vanes create flow channels through which externally supplied air flows from the blade root to the trailing edge. A previous paper documented the substantial tonal reductions of this Trailing Edge Rotor Blowing (TERB) fan. This report documents the broadband characteristics of TERB. The Active Noise Control Fan (ANCF), located at the NASA Glenn Research Center, was used as the proof-of-concept test bed. Two-component hotwire data behind the rotor, unsteady surface pressures on the stator vane, and farfield directivity acoustic data were acquired at blowing rates of 1.1, 1.5, and 1.8 percent of the total fan mass flow. The results indicate a substantial reduction in the rotor wake turbulent velocity and in the stator vane unsteady surface pressures. Based on the physics of the noise generation, these indirect measurements indicate the prospect of broadband noise reduction. However, since the broadband noise generated by the ANCF is rotor-dominated, any change in the rotor-stator interaction broadband noise levels is barely distinguishable in the farfield measurements.

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

Analysis of S Characteristics and Pressure Pulsations in a Pump-Turbine With Misaligned Guide Vanes.

PubMed

Sun, Hui; Xiao, Ruofu; Liu, Weichao; Wang, Fujun

2013-05-01

Growing environmental concerns and the need for better power balancing and frequency control have increased attention in renewable energy sources such as the reversible pump-turbine which can provide both power generation and energy storage. Pump-turbine operation along the S-shaped curve can lead to difficulties in loading the rejection process with unusual increases in water pressure, which lead to machine vibrations. Pressure fluctuations are the primary reason for unstable operation of pump-turbines. Misaligned guide vanes (MGVs) are widely used to control the stability in the S region. There have been experimental investigations and computational fluid dynamics (CFD) simulations of scale models with aligned guide vanes and MGVs with spectral analyses of the S curve characteristics and the pressure pulsations in the frequency and time-frequency domains at runaway conditions. The course of the S characteristic is related to the centrifugal force and the large incident angle at low flow conditions with large vortices forming between the guide vanes and the blade inlets and strong flow recirculation inside the vaneless space as the main factors that lead to the S-shaped characteristics. Preopening some of the guide vanes enables the pump-turbine to avoid the influence of the S characteristic. However, the increase of the flow during runaway destroys the flow symmetry in the runner leading to all asymmetry forces on the runner that leads to hydraulic system oscillations. The MGV technique also increases the pressure fluctuations in the draft tube and has a negative impact on stable operation of the unit.

System Study for Axial Vane Engine Technology

NASA Technical Reports Server (NTRS)

Badley, Patrick R.; Smith, Michael R.; Gould, Cedric O.

2008-01-01

The purpose of this engine feasibility study was to determine the benefits that can be achieved by incorporating positive displacement axial vane compression and expansion stages into high bypass turbofan engines. These positive-displacement stages would replace some or all of the conventional compressor and turbine stages in the turbine engine, but not the fan. The study considered combustion occurring internal to an axial vane component (i.e., Diesel engine replacing the standard turbine engine combustor, burner, and turbine); and external continuous flow combustion with an axial vane compressor and an axial vane turbine replacing conventional compressor and turbine systems.

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

NASA Astrophysics Data System (ADS)

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

2017-11-01

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

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

NASA Technical Reports Server (NTRS)

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

1998-01-01

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

A flight test of laminar flow control leading-edge systems

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

Improved Method for Prediction of Attainable Wing Leading-Edge Thrust

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

Intersecting vane machines

DOEpatents

Bailey, H. Sterling; Chomyszak, Stephen M.

2007-01-16

The invention provides a toroidal intersecting vane machine incorporating intersecting rotors to form primary and secondary chambers whose porting configurations minimize friction and maximize efficiency. Specifically, it is an object of the invention to provide a toroidal intersecting vane machine that greatly reduces the frictional losses through meshing surfaces without the need for external gearing by modifying the function of one or the other of the rotors from that of "fluid moving" to that of "valving" thereby reducing the pressure loads and associated inefficiencies at the interface of the meshing surfaces. The inventions described herein relate to these improvements.

Vane segment support and alignment device

DOEpatents

McLaurin, L.D.; Sizemore, J.D.

1999-07-13

A support and alignment assembly for supporting and aligning a vane segment is provided. The support and alignment assembly comprises a torque plate which defines an opening for receiving an eccentric pin and a locking end member for receiving a lock socket member. An eccentric pin adjustably supported by the torque plate opening for supporting and aligning a vane segment is provided. A lock socket member adapted to securely receive the eccentric pin and rotated therewith, and adjustably engage the torque plate locking end is provided. The lock socket member receives the eccentric pin, such that when the eccentric pin is adjusted to align the vane segment, the lock socket member engages the torque plate locking end to secure the vane segment in the desired position. 5 figs.

Vane segment support and alignment device

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

McLaurin, L.D.; Sizemore, J.D.

1999-07-13

A support and alignment assembly for supporting and aligning a vane segment is provided. The support and alignment assembly comprises a torque plate which defines an opening for receiving an eccentric pin and a locking end member for receiving a lock socket member. An eccentric pin adjustably supported by the torque plate opening for supporting and aligning a vane segment is provided. A lock socket member adapted to securely receive the eccentric pin and rotated therewith, and adjustably engage the torque plate locking end is provided. The lock socket member receives the eccentric pin, such that when the eccentric pinmore » is adjusted to align the vane segment, the lock socket member engages the torque plate locking end to secure the vane segment in the desired position. 5 figs.« less

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

NASA Astrophysics Data System (ADS)

Raghu Gowda, Belagumba Venkatachalaiah

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

Sealing intersecting vane machines

DOEpatents

Martin, Jedd N.; Chomyszak, Stephen M.

2005-06-07

The invention provides a toroidal intersecting vane machine incorporating intersecting rotors to form primary and secondary chambers whose porting configurations minimize friction and maximize efficiency. Specifically, it is an object of the invention to provide a toroidal intersecting vane machine that greatly reduces the frictional losses through intersecting surfaces without the need for external gearing by modifying the width of one or both tracks at the point of intermeshing. The inventions described herein relate to these improvements.

Sealing intersecting vane machines

DOEpatents

Martin, Jedd N [Providence, RI; Chomyszak, Stephen M [Attleboro, MA

2007-06-05

The invention provides a toroidal intersecting vane machine incorporating intersecting rotors to form primary and secondary chambers whose porting configurations minimize friction and maximize efficiency. Specifically, it is an object of the invention to provide a toroidal intersecting vane machine that greatly reduces the frictional losses through intersecting surfaces without the need for external gearing by modifying the width of one or both tracks at the point of intermeshing. The inventions described herein relate to these improvements.

Experimental And Numerical Study Of CMC Leading Edges In Hypersonic Flows

NASA Astrophysics Data System (ADS)

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

2011-05-01

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

Aerothermal Performance Constraint Analysis of Sharp Nosecaps and Leading Edges

NASA Technical Reports Server (NTRS)

Rizk, Yehia; Gee, Ken

2004-01-01

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

Use of leaning vanes in a two stage fan

NASA Technical Reports Server (NTRS)

Rao, G. V. R.; Digumarthi, R. V.

1975-01-01

The use of leaning vanes for tone noise reduction was examined in terms of their application in a typical two-stage high pressure ratio fan. In particular for stages designed with outlet guide vanes and zero swirl between stages, leaning the vanes of the first stage stator was studied, since increasing the number of vanes and the gap between stages do not provide the desired advantage. It was shown that noise reduction at higher harmonics of blade passing frequency can be obtained by leaning the vanes.

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.

Design Concepts for Cooled Ceramic Composite Turbine Vane

NASA Technical Reports Server (NTRS)

Boyle, Robert J.; Parikh, Ankur H.; Nagpal, VInod K.

2015-01-01

The objective of this work was to develop design concepts for a cooled ceramic vane to be used in the first stage of the High Pressure Turbine(HPT). To insure that the design concepts were relevant to the gas turbine industry needs, Honeywell International Inc. was subcontracted to provide technical guidance for this work. The work performed under this contract can be divided into three broad categories. The first was an analysis of the cycle benefits arising from the higher temperature capability of Ceramic Matrix Composite(CMC) compared with conventional metallic vane materials. The second category was a series of structural analyses for variations in the internal configuration of first stage vane for the High Pressure Turbine(HPT) of a CF6 class commercial airline engine. The third category was analysis for a radial cooled turbine vanes for use in turboshaft engine applications. The size, shape and internal configuration of the turboshaft engine vanes were selected to investigate a cooling concept appropriate to small CMC vanes.

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.

Nozzle cavity impingement/area reduction insert

DOEpatents

Yu, Yufeng Phillip; Itzel, Gary Michael; Osgood, Sarah Jane

2002-01-01

A turbine vane segment is provided that has inner and outer walls spaced from one another, a vane extending between the inner and outer walls and having leading and trailing edges and pressure and suction sides, the vane including discrete leading edge, intermediate, aft and trailing edge cavities between the leading and trailing edges and extending lengthwise of the vane for flowing a cooling medium; and an insert sleeve within at least one of the cavities and spaced from interior wall surfaces thereof. The insert sleeve has an inlet for flowing the cooling medium into the insert sleeve and has impingement holes defined in first and second walls thereof that respectively face the pressure and suction sides of the vane. The impingement holes of at least one of those first and second walls are defined along substantially only a first, upstream portion thereof, whereby the cooling flow is predominantly impingement cooling along a first region of the insert wall corresponding to the first, upstream portion and the cooling flow is predominantly convective cooling along a second region corresponding to a second, downstream portion of the at least one wall of the insert sleeve.

Membrane tension controls adhesion positioning at the leading edge of cells

PubMed Central

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

2017-01-01

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

Airfoil noise reductions through leading edge serrations

NASA Astrophysics Data System (ADS)

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

2015-02-01

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

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.

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

NASA Technical Reports Server (NTRS)

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

1980-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

REMOTE RECORDING ANNULAR VANE ASSEMBLY

DOEpatents

Wehmann, G.

1963-06-25

A weather vane apparatus is described which is capable of movement in horizontal and vertical planes. Associated with the vane are tangent potentiometers, commutators, and other electrical apparatus for deriving electrical output voltages as a function of the wind direction. The apparatus is particularly adapted for use with an anemometer to provide an electrical output indicating the amount and direction of an up or down draft. (AEC)

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

NASA Astrophysics Data System (ADS)

Hrynuk, John; Bohl, Douglas

2013-11-01

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

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

NASA Astrophysics Data System (ADS)

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

2017-08-01

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

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

PubMed Central

Muir, Rowan Eveline; Arredondo-Galeana, Abel

2017-01-01

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

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

PubMed

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

2017-08-01

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

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

NASA Astrophysics Data System (ADS)

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

2006-11-01

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

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.

Modeling of Electron Transpiration Cooling for Leading Edges of Hypersonic Vehicles

NASA Astrophysics Data System (ADS)

Hanquist, Kyle Matthew

The development of aeronautics has been largely driven by the passion to fly faster. From the flight of the Wright Flyer that flew 48 km/hr to the recent advances in hypersonic flight, most notably NASA's X-43A that flew at over 3 km/s, the velocity of flight has steadily increased. However, as these hypersonic speeds are reached and increased, contradicting aerothermodynamic design requirements present themselves. For example, a hypersonic cruise vehicle requires sharp leading edges to decrease the drag in order to maximize the range. However, the aerodynamic performance gains obtained by having a sharp leading edge come at the cost of very high, localized heating rates. There is currently no ideal way to manage these heating loads for sustained hypersonic flight, especially as flight velocities continue to increase. An approach that has been recently proposed involves using thermo-electric materials on these sharp leading edges to manage the heating loads. When exposed to high convective heating rates, these materials emit a current of electrons that leads to a cooling effect of the surface of the vehicle called electron transpiration cooling (ETC). This dissertation focuses on developing a modeling approach to investigate this phenomenon. The research includes developing and implementing an approach for ETC into a computational fluid dynamics code for simulation of hypersonic flow that accounts for electron emission from the surface. Models for space-charge-limited emission are also developed and implemented in order to accurately determine the level of emission from the surface. This work involves developing analytic models and assessing them using a direct-kinetic plasma sheath solver. Electric field effects are also implemented in the modeling approach, which accounts for forced diffusion and Joule heating. Finally, the modeling approach is coupled to a material response code in order to model the heat transfer into the material surface. Using this modeling

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.

Ceramic matrix composite turbine engine vane

NASA Technical Reports Server (NTRS)

Schaff, Jeffery R. (Inventor); Shi, Jun (Inventor)

2012-01-01

A vane has an airfoil shell and a spar within the shell. The vane has an outboard shroud at an outboard end of the shell and an inboard platform at an inboard end of the shell. The shell includes a region having a coefficient of thermal expansion (CTE) varying with depth.

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

Performance of a vane driven-gear pump

NASA Technical Reports Server (NTRS)

Heald, R H

1921-01-01

Given here are the results of a test conducted in a wind tunnel on the performance of a vane-driven gear pump used to pump gasoline upward into a small tank located within the upper wing from which it flows by gravity to the engine carburetor. Information is given on the efficiency of the pump, the head resistance of the vanes, the performance and characteristics of the unit with and without housing about the vanes, the pump performance when motor driven, and resistance and power characteristics.

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

PubMed Central

Laplante, Caroline

2011-01-01

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

Flow analysis in a vane-type surface tension propellant tank

NASA Astrophysics Data System (ADS)

Yu, A.; Ji, B.; Zhuang, B. T.; Hu, Q.; Luo, X. W.; Y Xu, H.

2013-12-01

Vane-type surface tension tanks are widely used as the propellant management devices in spacecrafts. This paper treats the two-phase flow inside a vane-type surface tension tank. The study indicates that the present numerical methods such as time-dependent Navier-Stokes equations, VOF model can reasonably predict the flow inside a propellant tank. It is clear that the vane geometry has important effects on transmission performance of the liquid. for a vane type propellant tank, the vane having larger width, folding angle, height of folded side and clearance is preferable if possible.

Impact of vane size and separation on radiometric forces for microactuation

NASA Astrophysics Data System (ADS)

Gimelshein, Natalia; Gimelshein, Sergey; Ketsdever, Andrew; Selden, Nathaniel

2011-04-01

A kinetic approach is used to study the feasibility of increasing the efficiency of microactuators that use radiometric force through etching holes in a single radiometer vane. It has been shown that a radiometer that consists of small vanes is capable of producing at least an order of magnitude larger force than a single-vane radiometer that takes up the same area. The optimum gap between the vanes is found to be slightly smaller than the vane size, with the optimum Knudsen number of about 0.05 based on the vane height.

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

PubMed

Rao, Chen; Liu, Hao

2018-06-08

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

Numerical investigation of a centrifugal compressor with circumferential grooves in vane diffuser

NASA Astrophysics Data System (ADS)

Chen, X. F.; Qin, G. L.; Ai, Z. J.

2015-08-01

Enhancing stall and surge margin has a great importance for the development of turbo compressors. The application of casing treatment is an effective measure to expand the stall margin and stable operation range. Numerical investigations were conducted to predict the performance of a low flow rate centrifugal compressor with circumferential groove casing treatment in vane diffuser. Numerical cases with different radial location, radial width and axial depth of a circumferential single groove and different numbers of circumferential grooves were carried out to compare the results. The CFD analyses results show that the centrifugal compressor with circumferential grooves in diffuser can extend stable range by about 9% while the efficiency over the whole operating range decreases by 0.2 to 1.7%. The evaluation based on stall margin improvement showed the optimal position for the groove to be located was indicated to exist near the leading edge of the diffuser, and a combination of position, width, depth and numbers of circumferential grooves that will maximize both surge margin range and efficiency.

Heat pipes for wing leading edges of hypersonic vehicles

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2016-09-01

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

Validation and Analysis of Numerical Results for a Two-Pass Trapezoidal Channel With Different Cooling Configurations of Trailing Edge.

PubMed

Siddique, Waseem; El-Gabry, Lamyaa; Shevchuk, Igor V; Fransson, Torsten H

2013-01-01

High inlet temperatures in a gas turbine lead to an increase in the thermal efficiency of the gas turbine. This results in the requirement of cooling of gas turbine blades/vanes. Internal cooling of the gas turbine blade/vanes with the help of two-pass channels is one of the effective methods to reduce the metal temperatures. In particular, the trailing edge of a turbine vane is a critical area, where effective cooling is required. The trailing edge can be modeled as a trapezoidal channel. This paper describes the numerical validation of the heat transfer and pressure drop in a trapezoidal channel with and without orthogonal ribs at the bottom surface. A new concept of ribbed trailing edge has been introduced in this paper which presents a numerical study of several trailing edge cooling configurations based on the placement of ribs at different walls. The baseline geometries are two-pass trapezoidal channels with and without orthogonal ribs at the bottom surface of the channel. Ribs induce secondary flow which results in enhancement of heat transfer; therefore, for enhancement of heat transfer at the trailing edge, ribs are placed at the trailing edge surface in three different configurations: first without ribs at the bottom surface, then ribs at the trailing edge surface in-line with the ribs at the bottom surface, and finally staggered ribs. Heat transfer and pressure drop is calculated at Reynolds number equal to 9400 for all configurations. Different turbulent models are used for the validation of the numerical results. For the smooth channel low-Re k-ɛ model, realizable k-ɛ model, the RNG k-ω model, low-Re k-ω model, and SST k-ω models are compared, whereas for ribbed channel, low-Re k-ɛ model and SST k-ω models are compared. The results show that the low-Re k-ɛ model, which predicts the heat transfer in outlet pass of the smooth channels with difference of +7%, underpredicts the heat transfer by -17% in case of ribbed channel compared to

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.

Turbocharger with sliding piston, and having vanes and leakage dams

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

Roberts, Quentin; Alnega, Ahmed

2011-12-06

A turbocharger having a sliding piston for regulating exhaust gas flow into the turbine wheel includes a set of first vanes mounted on a fixed first wall of the turbine nozzle and projecting axially toward an opposite second wall of the nozzle, and/or a set of second vanes mounted on the end of the piston and projecting in an opposite axial direction toward the first wall of the nozzle. For the/each set of vanes, there are leakage dams formed on the wall that is adjacent the vane tips when the piston is closed. The leakage dams are closely adjacent themore » vane tips and discourage exhaust gas from leaking in a generally radial direction past the vane tips as the piston just begins to open from its fully closed position.« less

On the leading edge vortex of thin wings

NASA Astrophysics Data System (ADS)

Arredondo, Abel; Viola, Ignazio Maria

2016-11-01

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

Cooled turbine vane with endcaps

DOEpatents

Cunha, Frank J.; Schiavo, Jr., Anthony L.; Nordlund, Raymond Scott; Malow, Thomas; McKinley, Barry L.

2002-01-01

A turbine vane assembly which includes an outer endcap having a plurality of generally straight passages and passage segments therethrough, an inner endcap having a plurality of passages and passage segments therethrough, and a vane assembly having an outer shroud, an airfoil body, and an inner shroud. The outer shroud, airfoil body and inner shroud each have a plurality of generally straight passages and passage segments therethrough as well. The outer endcap is coupled to the outer shroud so that outer endcap passages and said outer shroud passages form a fluid circuit. The inner endcap is coupled to the inner shroud so that the inner end cap passages and the inner shroud passages from a fluid circuit. Passages in the vane casting are in fluid communication with both the outer shroud passages and the inner shroud passages. Passages in the outer endcap may be coupled to a cooling system that supplies a coolant and takes away the heated exhaust.

Ablative overlays for Space Shuttle leading edge ascent heat protection

NASA Technical Reports Server (NTRS)

Strauss, E. L.

1975-01-01

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

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

Probabilistic Assessment of a CMC Turbine Vane

NASA Technical Reports Server (NTRS)

Murthy, Pappu L. N.; Brewer, Dave; Mital, Subodh K.

2004-01-01

In order to demonstrate the advanced CMC technology under development within the Ultra Efficient Engine Technology (UEET) program, it has been planned to fabricate, test and analyze an all CMC turbine vane made of a SiC/SiC composite material. The objective was to utilize a 5-II Satin Weave SiC/CVI SiC/ and MI SiC matrix material that was developed in-house under the Enabling Propulsion Materials (EPM) program, to design and fabricate a stator vane that can endure successfully 1000 hours of engine service conditions operation. The design requirements for the vane are to be able to withstand a maximum of 2400 F within the substrate and the hot surface temperature of 2700 F with the aid of an in-house developed Environmental/Thermal Barrier Coating (EBC/TBC) system. The vane will be tested in a High Pressure Burner Rig at NASA Glenn Research Center facility. This rig is capable of simulating the engine service environment. The present paper focuses on a probabilistic assessment of the vane. The material stress/strain relationship shows a bilinear behavior with a distinct knee corresponding to what is often termed as first matrix cracking strength. This is a critical life limiting consideration for these materials. The vane is therefore designed such that the maximum stresses are within this limit so that the structure is never subjected to loads beyond the first matrix cracking strength. Any violation of this design requirement is considered as failure. Probabilistic analysis is performed in order to determine the probability of failure based on this assumption. In the analysis, material properties, strength, and pressures are considered random variables. The variations in properties and strength are based on the actual experimental data generated in house. The mean values for the pressures on the upper surface and the lower surface are known but their distributions are unknown. In the present analysis the pressures are considered normally distributed with a nominal

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.

Flow and heat transfer in an L-shaped cooling passage with ribs and pin fins for the trailing edge of a gas-turbine vane and blade

NASA Astrophysics Data System (ADS)

Pardeshi, Irsha

Efficient and effective cooling of the trailing edges of gas-turbine vanes and blades is challenging because there is very little space to work with. In this study, CFD simulations based on steady RANS closed by the shear-stress transport turbulence model were performed to study the flow and heat transfer in an L-shaped duct for the trailing edge under two operating conditions. One operating condition, referred to as the laboratory condition, where experimental measurements were made, has a Reynolds number at the duct inlet of ReD = 15,000, coolant inlet temperature of Tinlet = 300 K, wall temperature of Twall = 335 K, a back pressure of Pb = 1 atm. When rotating, the angular speed was O = 1,000 rpm. The other condition, referred to as the engine-relevant condition, has Re D = 150,000 at the duct inlet, Tinlet = 673 K, Twall = 1,173 K, and Pb = 25 atm. When rotating, O was 3,600 rpm. The objective is to understand the nature of the flow and heat transfer in an L-shaped cooling passage for the trailing edge that has a combination of ribs and pin fins under rotating and non-rotating conditions with focus on how pin fins and ribs distribute the flow throughout the passage and to understand what features of the flow and heat transfer can or cannot be extrapolated from the laboratory to the engine-relevant operating conditions. When there is no rotation, results obtained show that for both operating conditions, the pin fins minimized the size of the separation bubble when the flow exits the inlet duct into the expanded portion of the L-shaped duct. The size of the separation bubble at the tip of the L-shaped duct created by the adverse pressure gradient is quite large for the laboratory condition and relatively small for the engine condition. Each rib was found to create two sets of recirculating flows, one just upstream of the rib because of the adverse pressure gradient induced by the rib and one just downstream of the rib because of flow separation from a sharp edge

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

NASA Technical Reports Server (NTRS)

Darden, C. M.

1985-01-01

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

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

NASA Technical Reports Server (NTRS)

Silverstein, C. C.

1971-01-01

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

Solar energy system with wind vane

DOEpatents

Grip, Robert E

2015-11-03

A solar energy system including a pedestal defining a longitudinal axis, a frame that is supported by the pedestal and that is rotateable relative to the pedestal about the longitudinal axis, the frame including at least one solar device, and a wind vane operatively connected to the frame to urge the frame relative to the pedestal about the longitudinal axis in response to wind acting on the wind vane.

Noise generated by convected gusts interacting with swept airfoil cascades

NASA Astrophysics Data System (ADS)

Envia, E.; Kerschen, E. J.

1986-07-01

An analysis is developed for the noise generated by the interaction of a rotor viscous wake with a cascade of swept stator vanes. The stator vanes span a channel formed by infinite parallel walls and containing a subsonic mean flow. High frequency interactions, for which the noise generation is concentrated at the vane leading edge, are considered. The analysis utilizes a superposition of the solution to the isolated stator vane problem, presented in an earlier paper, to develop an approximate solution to the cascade problem. The rotor wake model includes the features of wake circumferential lean and a linear spanwise variation of the magnitude of the wake deficit velocity. Calculations are presented which show that, for rotor wakes with moderate circumferential lean, stator sweep produces substantial reductions in noise level. The vane sweep must be oriented to enhance the phase lags along the vane leading edge produced by wake lean. The noise levels are found to be fairly insensitive to spanwise variations in the wake deficit.

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

PubMed

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

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

Improving the Hydrodynamic Performance of Diffuser Vanes via Shape Optimization

NASA Technical Reports Server (NTRS)

Goel, Tushar; Dorney, Daniel J.; Haftka, Raphael T.; Shyy, Wei

2007-01-01

The performance of a diffuser in a pump stage depends on its configuration and placement within the stage. The influence of vane shape on the hydrodynamic performance of a diffuser has been studied. The goal of this effort has been to improve the performance of a pump stage by optimizing the shape of the diffuser vanes. The shape of the vanes was defined using Bezier curves and circular arcs. Surrogate model based tools were used to identify regions of the vane that have a strong influence on its performance. Optimization of the vane shape, in the absence of manufacturing, and stress constraints, led to a nearly nine percent reduction in the total pressure losses compared to the baseline design by reducing the extent of the base separation.

Optimization analysis of a new vane MRF damper

NASA Astrophysics Data System (ADS)

Zhang, J. Q.; Feng, Z. Z.; Jing, Q.

2009-02-01

The primary purpose of this study was to provide the optimization analysis certain characteristics and benefits of a vane MRF damper. Based on the structure of conventional vane hydraulic damper for heavy vehicle, a narrow arc gap between clapboard and rotary vane axle, which one rotates relative to the other, was designed for MRF valve and the mathematical model of damping was deduced. Subsequently, the finite element analysis of electromagnetic circuit was done by ANSYS to perform the optimization process. Some ways were presented to augment the damping adjustable multiple under the condition of keeping initial damping forces and to increase fluid dwell time through the magnetic field. The results show that the method is useful in the design of MR dampers and the damping adjustable range of vane MRF damper can meet the requirement of heavy vehicle semi-active suspension system.

High Temperature Ceramic Guide Vane Temperature and Pressure Distribution Calculation for Flow with Cooling Jets

NASA Technical Reports Server (NTRS)

Srivastava, Rakesh

2004-01-01

A ceramic guide vane has been designed and tested for operation under high temperature. Previous efforts have suggested that some cooling flow may be required to alleviate the high temperatures observed near the trailing edge region. The present report describes briefly a three-dimensional viscous analysis carried out to calculate the temperature and pressure distribution on the blade surface and in the flow path with a jet of cooling air exiting from the suction surface near the trailing edge region. The data for analysis was obtained from Dr. Craig Robinson. The surface temperature and pressure distribution along with a flowfield distribution is shown in the results. The surface distribution is also given in a tabular form at the end of the document.

Leading-Edge Flow Sensing for Aerodynamic Parameter Estimation

NASA Astrophysics Data System (ADS)

Saini, Aditya

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

Industrial stator vane with sequential impingement cooling inserts

DOEpatents

Jones, Russell B; Fedock, John A; Goebel, Gloria E; Krueger, Judson J; Rawlings, Christopher K; Memmen, Robert L

2013-08-06

A turbine stator vane for an industrial engine, the vane having two impingement cooling inserts that produce a series of impingement cooling from the pressure side to the suction side of the vane walls. Each insert includes a spar with a row of alternating impingement cooling channels and return air channels extending in a radial direction. Impingement cooling plates cover the two sides of the insert and having rows of impingement cooling holes aligned with the impingement cooling channels and return air openings aligned with the return air channel.

Ceramic matrix composite turbine engine vane

NASA Technical Reports Server (NTRS)

Prill, Lisa A. (Inventor); Schaff, Jeffery R. (Inventor); Shi, Jun (Inventor)

2012-01-01

A vane has an airfoil shell and a spar within the shell. The vane has an outboard shroud at an outboard end of the shell and an inboard platform at an inboard end of the shell. The shell includes a region having a depth-wise coefficient of thermal expansion and a second coefficient of thermal expansion transverse thereto, the depth-wise coefficient of thermal expansion being greater than the second coefficient of thermal expansion.

Vane array with one or more non-integral platforms

DOEpatents

Lohaus, Andrew S.; Campbell, Christian Xavier; Miller, Jr, Samuel R.; Marra, John J.

2016-07-12

A vane array adapted to be coupled to a vane carrier within a gas turbine engine is provided comprising: a plurality of elongated airfoils comprising at least a first airfoil and a second airfoil located adjacent to one another; a U-ring; first connector structure for coupling a radially inner end section of each of the first and second airfoils to the U-ring; second connector structure for coupling a radially outer end section of each of the first and second airfoils to the vane carrier; a platform extending between the first and second airfoils; and platform connector structure for coupling the platform to one of the U-ring and the vane carrier.

Streakline flow visualization study of a horseshoe vortex in a large-scale, two-dimensional turbine stator cascade

NASA Technical Reports Server (NTRS)

Gaugler, R. E.; Russell, L. M.

1980-01-01

Neutrally buoyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large-scale, two-dimensional, turbine stator cascade. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16-mm movie film. Individual frames from the film have been selected, and overlayed to show the details of the horseshoe vortex around the leading edge. The transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

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

NASA Technical Reports Server (NTRS)

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

1999-01-01

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

The Flow Field on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2009-11-01

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

The effectiveness of vane-aileron excitation in the experimental determination of flutter speed by parameter identification

NASA Technical Reports Server (NTRS)

Nissim, Eli

1990-01-01

The effectiveness of aerodynamic excitation is evaluated analytically in conjunction with the experimental determination of flutter dynamic pressure by parameter identification. Existing control surfaces were used, with an additional vane located at the wingtip. The equations leading to the identification of the equations of motion were reformulated to accommodate excitation forces of aerodynamic origin. The aerodynamic coefficients of the excitation forces do not need to be known since they are determined by the identification procedure. The 12 degree-of-freedom numerical example treated in this work revealed the best wingtip vane locations, and demonstrated the effectiveness of the aileron-vane excitation system. Results from simulated data gathered at much lower dynamic pressures (approximately half the value of flutter dynamic pressure) predicted flutter dynamic pressures with 2-percent errors.

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

NASA Technical Reports Server (NTRS)

1972-01-01

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

Potential environmental effects of the leading edge hydrokinetic energy technology.

DOT National Transportation Integrated Search

2017-05-01

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

Detailed film cooling effectiveness and three component velocity field measurements on a first stage turbine vane subject to high freestream turbulence

NASA Astrophysics Data System (ADS)

Polanka, Marcus Damian

1999-11-01

This experimental program studied the effects of high freestream turbulence on film cooling for a turbine vane. This investigation focussed on the showerhead and pressure surface of an airfoil. An emphasis of this study was to acquire highly detailed film cooling effectiveness and velocity measurements in the showerhead region. Acquisition of both pieces of information resulted in detailed knowledge of the physics involved in the interaction of the coolant jets and the freestream flow in this region of an airfoil. By generating a 18% turbulence level at the leading edge of the airfoil, the impact of elevated freestream turbulence was also studied. Of further interest was the affect of a highly turbulent flow resulting from both the freestream flow as well as that generated from the showerhead jets themselves, further downstream. The impact of this turbulent approach flow will have significant consequence on downstream film cooling designs. In order to achieve the desired goals, modification to the existing closed loop wind tunnel facility was required. The new tunnel consisted of a test section containing a center, instrumented airfoil with inner and outer walls positioned to match the flow parameters around the center airfoil. The center airfoil was built at a nine times scale ratio. In utilizing this large scale vane and still matching the engine conditions, a better understanding of leading edge film cooling was gained. This was a result of the high spatial resolution of the flow field gained from the large scale of the airfoil. This benefited both the Laser Doppler Velocimeter (LDV) system for velocity measurements and the infrared camera used for thermal field measurements. High effectiveness levels were measured throughout the showerhead region. This was attributed to a build up of coolant along the span of the airfoil. The introduction of a high freestream turbulence level increased the uniformity at the expense of lower overall effectiveness levels

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.

Research on liquid sloshing performance in vane type tank under microgravity

NASA Astrophysics Data System (ADS)

Hu, Q.; Li, Y.; Liu, J. T.; Liang, J. Q.

2016-05-01

Propellant management device (PMD) in vane type tank mainly comprises of vane type structure parts, whose performance of restraining liquid sloshing should satisfy spacecraft requirements of high stabilization and fast orbital maneuver. Aiming at liquid sloshing performance in vane type tank under microgravity environment, gas-liquid flow model based on the volume of fluid (VOF) method was put forward, and via numerical simulation liquid sloshing performances of vane type PMD with anti-sloshing baffles and without anti-sloshing baffles in microgravity were analyzed and compared. Simulation results reveal that liquid sloshing performance of vane type PMD with anti-sloshing baffles is markedly superior vane type PMD without anti-sloshing baffles and the baffles make liquid surface become stable fast. Then by comparing between results of microgravity experiments and results of numerical simulations, they are very similar. According to present research, vane type PMD with antisloshing baffles has better effects on restraining liquid sloshing and is able to restrain observably propellant sloshing in tanks in order to satisfy spacecraft requirements of high stabilization and fast orbital maneuver.

Liquid–liquid mixing studies in annular centrifugal contactors comparing stationary mixing vane options

DOE PAGES

Wardle, Kent E.

2015-09-11

Comparative studies of multiphase operation of an annular centrifugal contactor show the impact of housing stationary mixing vane configuration. A number of experimental results for several different mixing vane options are reported for operation of a 12.5 cm engineering-scale contactor unit. Fewer straight vanes give greater mixing-zone hold-up compared to curved vanes. Quantitative comparison of droplet size distribution also showed a significant decrease in mean diameter for four straight vanes versus eight curved vanes. This set of measurements gives a compelling case for careful consideration of mixing vane geometry when evaluating hydraulic operation and extraction process efficiency of annular centrifugalmore » contactors.« less

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

NASA Astrophysics Data System (ADS)

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

2017-10-01

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

Chronophin coordinates cell leading edge dynamics by controlling active cofilin levels

PubMed Central

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

2015-01-01

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

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

PubMed Central

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

2016-01-01

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

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

NASA Astrophysics Data System (ADS)

Al Ahmari, Saeed Abdullah Saeed

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

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

NASA Technical Reports Server (NTRS)

Glass, David E.

1998-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2017-11-01

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

Liquid–Liquid Mixing Studies in Annular Centrifugal Contactors Comparing Stationary Mixing Vane Options

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

Wardle, Kent E.

2015-11-10

Comparative studies of multiphase operation of annular centrifugal contactors showing the impact of housing stationary mixing vane configuration. A number of experimental results for several different mixing vane options are reported with selected measurements in a lab-scale 5 cm contactor and 12.5 cm engineering-scale unit. Fewer straight vanes give greater mixingzone hold-up compared to curved vanes. Quantitative comparison of droplet size distribution also showed a significant decrease in mean diameter for four straight vanes versus eight curved vanes. This set of measurements gives a compelling case for careful consideration of mixing vane geometry when evaluating hydraulic operation and extraction processmore » efficiency of annular centrifugal contactors.« less

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.

Demonstration and Field Evaluation of Streambank Stabilization with Submerged Vanes

USGS Publications Warehouse

Whitman, H.; Hoopes, J.; Poggi, D.; Fitzpatrick, F.; Walz, K.; ,

2001-01-01

The effectiveness of submerged vanes for reducing bank erosion and improving aquatic habitat is being evaluated at a site on North Fish Creek, a Lake Superior tributary. Increased runoff from agricultural areas with clayey soils has increased flood magnitudes and the erosion potential/transport capacity of the stream. Most of the creek's sediment load originates from the erosion of 17 large bluffs. This creek contains important recreational fisheries that are potentially limited by the loss of aquatic habitat from deposition of sediment on spawning beds. Submerged vanes are a cost effective and environmentally less intrusive alternative to traditional structural stabilization measures. Submerged vanes protrude from a channel bed, are oriented at an angle to the local velocity, and are distributed along a portion of channel. They induce a transverse force and torque on the flow along with longitudinal vortexes that alter the cross sectional shape and alignment of the channel. Submerged vanes were installed at a bluff/bend site in summer and fall 2000. The number, size, and layout of the vanes were based upon the channel morphology under estimated bankfull conditions. The effectiveness of the vanes will be evaluated by comparing surveys of the bluff face, streamflow, and channel conditions for several years after installation of the submerged vanes with surveys before and immediately after their installation.

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.

Optimization of a Low Heat Load Turbine Nozzle Guide Vane

DTIC Science & Technology

2006-03-01

HEAT LOAD TURBINE NOZZLE GUIDE VANE THESIS Presented to the Faculty Department of Aeronautical and Astronautical Engineering ...a function of turbine inlet temperature. .................... 2 Figure 2 Traditional turbofan engine and stator vane location (from Ref [1...the non-rotating stator vanes within a cross-section of a classical two-spool turbofan engine which has an inlet, 4 compressor, combustor, turbine

Turbocharger with variable nozzle having vane sealing surfaces

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

Arnold, Philippe; Petitjean, Dominique; Ruquart, Anthony

2011-11-15

A variable nozzle for a turbocharger includes a plurality of vanes rotatably mounted on a nozzle ring and disposed in a nozzle flow path defined between the nozzle ring and an opposite nozzle wall. Either or both of the faces of the nozzle ring and nozzle wall include(s) at least one step that defines sealing surfaces positioned to be substantially abutted by airfoil surfaces of the vanes in the closed position of the vanes and to be spaced from the airfoil surfaces in positions other than the closed position. This substantial abutment between the airfoil surfaces and the sealing surfacesmore » serves to substantially prevent exhaust gas from leaking past the ends of the airfoil portions. At the same time, clearances between the nozzle ring face and the end faces of the airfoil portions can be sufficiently large to prevent binding of the vanes under all operating conditions.« less

Experimental investigation of in-cylinder air flow to optimize number of helical guide vanes to enhance DI diesel engine performance using mamey sapote biodiesel

NASA Astrophysics Data System (ADS)

Kumar, A. Raj; Janardhana Raju, G.; Hemachandra Reddy, K.

2018-03-01

The current research work investigates the influence of helical guide vanes in to the intake runner of a D.I diesel engine operating by the high viscous Mamey Sapote biodiesel to enhance in-cylinder suction air flow features. Helical guide vanes of different number of vanes are produced from 3D printing and placed in the intake manifold to examine the air flow characteristics. Four different helical guide vane devices namely 3, 4, 5 and 6 vanes of the same dimensions are tested in a D.I diesel engine operating with Mamey Sapote biodiesel blend. As per the experimental results of engine performance and emission characteristics, it is found that 5 vanes helical guide vane swirl device exhibited in addition number of increased improvements such as the brake power and bake thermal efficiency by 2.4% and 8.63% respectively and the HC, NOx, Carbon monoxide and, Smoke densities are reduced by 15.62%, 4.23%, 14.27% and 9.6% at peak load operating conditions as collate with normal engine at the same load. Hence this investigation concluded that Helical Guide Vane Devices successfully enhanced the in-cylinder air flow to improve better addition of Mamey Sapote biodiesel with air leading in better performance of the engine than without vanes.

Design Evolution: From Rolling Piston to Revolving Vane to Cross-Vane Expander-compressor unit

NASA Astrophysics Data System (ADS)

T, K.; S, Yap K.

2015-08-01

In this paper the design evolution from rolling piston compressor (RP) to revolving vane compressor (RV) and finally to cross-vane expander-compressor unit (CVEC) is presented and discussed. The details in the design philosophy which initiated this evolution will be presented and discussed. It is estimated that more than 90% of the room air-conditioners uses RP in its compressor. This is because of its advantages: it has the few parts, it is simple geometrically and it is reliable. However, it is with no weaknesses. RP's weaknesses lie in three parts: too many rubbing surfaces with high relative rubbing velocity with each other which give rise to high frictional losses; components (eccentric and roller) are not rotating at their centres and resulted in unnecessary inherent vibration; vane tip is constantly rubbing against the roller making it a weakest part in design. To overcome RP's weaknesses, RV is introduced. As compared to RP, RV has fewer rubbing surfaces, and the relative velocities among these surfaces are reduced Components in RV rotate at their own centres and there is no inherent vibration; the rubbing at the vane tip has been eliminated completely. However, like RP, RV also has a large rotor which occupied "useful" space and making the working chamber relatively small. To overcome this latter problem, CVEC is introduced. In this newly invented CVCE, not only parts are all concentric and rotate at their own centres, the unit also recovers expansion energy and hence significantly reduces energy required by the refrigeration systems. And, more importantly, all these are carried out not at the expense of the additional cost. In this paper, details on these three compressors are explained, compared and their respective uniqueness are shown and discussed.

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.

Simulation of Flow Through Breach in Leading Edge at Mach 24

NASA Technical Reports Server (NTRS)

Gnoffo, Peter A.; Alter, Stephen J.

2004-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

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

NASA Technical Reports Server (NTRS)

Squires, Becky

1993-01-01

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

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

PubMed

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

2017-07-04

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

Broadband Venetian-Blind Polarizer With Dual Vanes

NASA Technical Reports Server (NTRS)

Conroy, Bruce L.; Hoppe, Daniel J.

1995-01-01

Improved venetian-blind polarizer features optimized tandem, two-layer vane configuration reducing undesired reflections and deformation of radiation pattern below those of prior single-layer vane configuration. Consists of number of thin, parallel metal strips placed in path of propagating radio-frequency beam. Offers simple way to convert polarization from linear to circular or from circular to linear. Particularly useful for beam-wave-guide applications.

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

NASA Technical Reports Server (NTRS)

Luckring, James M.

2004-01-01

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

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

NASA Technical Reports Server (NTRS)

Fisher, David F.; Fischer, Michael C.

1987-01-01

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

Study on Orbital Liquid Transport and Interface Behavior in Vane Tank

NASA Astrophysics Data System (ADS)

Kang, Qi; Rui, Wei

2016-07-01

Liquid propellant tank is used to supply gas free liquid for spacecraft as an important part of propulsion system. The liquid behavior dominated by surface tension in microgravity is obviously different with that on the ground, which put forward a new challenge to the liquid transport and relocation. The experiments which are investigated at drop tower in National Microgravity Lab have concentrated on liquid relocation following thruster firing. Considered that the liquid located at the bottom in the direction of the acceleration vector, a sphere scale vane tank is used to study the liquid-gas interface behaviors with different acceleration vector and different filling independently and we obtain a series of stable equilibrium interface and relocation time. We find that there is an obvious sedimentation in the direction of acceleration vector when fill rate greater than 2% fill. Suggestions have been put forward that outer vanes transferring liquid to the outlet should be fixed and small holes should be dogged at the vane close to the center post to improve the liquid flow between different vanes when B0 is greater than 2.5. The research about liquid transport alone ribbon vanes is simulated though software Flow3D. The simulation process is verified by comparing the liquid lip and vapor-liquid interface obtained from drop tower experiment and simulation result when fill rate is 15%. Then the influence of fill rate, numbers of vanes and the gap between vane and wall is studied through the same simulate process. Vanes' configurations are also changed to study the effect on the lip and liquid volume below some section. Some suggestions are put forward for the design of vanes.

Low-Speed Fan Noise Reduction With Trailing Edge Blowing

NASA Technical Reports Server (NTRS)

Sutliff, Daniel L.; Tweedt, Daniel L.; Fite, E. Brian; Envia, Edmane

2002-01-01

An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a trailing edge slot. Composite hollow rotor blades with internal flow passages were designed based on analytical codes modeling the internal flow. The hollow blade with interior guide vanes creates flow channels through which externally supplied air flows from the root of the blade to the trailing edge. The impact of the rotor wake-stator interaction on the acoustics was also predicted analytically. The Active Noise Control Fan, located at the NASA Glenn Research Center, was used as the proof- of-concept test bed. In-duct mode and farfield directivity acoustic data were acquired at blowing rates (defined as mass supplied to trailing edge blowing system divided by fan mass flow) ranging from 0.5 to 2.0 percent. The first three blade passing frequency harmonics at fan rotational speeds of 1700 to 1900 rpm were analyzed. The acoustic tone power levels (PWL) in the inlet and exhaust were reduced 11.5 and -0.1, 7.2 and 11.4, 11.8 and 19.4 PWL dB, respectively. The farfield tone power levels at the first three harmonics were reduced 5.4, 10.6, and 12.4 dB PWL. At selected conditions, two-component hotwire and stator vane unsteady surface pressures were acquired. These measurements illustrate the physics behind the noise reduction.

Active Control of Fan Noise by Vane Actuators

NASA Technical Reports Server (NTRS)

Curtis, Alan R. D.

1999-01-01

An active noise control system for ducted fan noise was built that uses actuators located in stator vanes. The actuators were piezoelectric benders manufactured using the THUNDER technology and were custom designed for the application. The active noise control system was installed in the NASA ANCF rig. Four actuator array with a total of 168 actuators in 28 stator vanes were used. Simultaneous reductions of acoustic power in both the inlet and exhaust duct were demonstrated for a fan disturbance that contained two radial mode orders in both inlet and exhaust. Total power levels in the target modes were reduced by up to 9 dB in the inlet and total tone levels by over 6 dB while exhaust power levels were reduced by up to 3 dB. Far field sound pressure level reductions of up to 17 dB were observed. A simpler control system, matched to the location of the disturbance with two radial actuator arrays, was demonstrated to control total acoustic power in four disturbance modes simultaneously in inlet and exhaust. The vane actuator met the requirements given for the ANCF, although in practice the performance of the system was limited by the constraints of the power amplifiers and the presence of control spillover. The vane actuators were robust. None of the 168 vane actuators failed during the tests.

Characterization of Rotating Detonation Engine Exhaust Through Nozzle Guide Vanes

DTIC Science & Technology

2013-03-21

THROUGH NOZZLE GUIDE VANES THESIS Presented to the Faculty Department of Aeronautics and Astronautics Graduate School of Engineering and Management Air...the first Nozzle Guide Vane (NGV) section from a T63 gas turbine engine to a 6 inch diameter RDE was designed and built for this study. Pressure...CHARACTERIZATION OF ROTATING DETONATION ENGINE EXHAUST THROUGH NOZZLE GUIDE VANES THESIS Nick D. DeBarmore, Second Lieutenant, USAF AFIT/GAE/ENY/13

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.

Heat transfer enhancement with mixing vane spacers using the field synergy principle

NASA Astrophysics Data System (ADS)

Yang, Lixin; Zhou, Mengjun; Tian, Zihao

2017-01-01

The single-phase heat transfer characteristics in a PWR fuel assembly are important. Many investigations attempt to obtain the heat transfer characteristics by studying the flow features in a 5 × 5 rod bundle with a spacer grid. The field synergy principle is used to discuss the mechanism of heat transfer enhancement using mixing vanes according to computational fluid dynamics results, including a spacer grid without mixing vanes, one with a split mixing vane, and one with a separate mixing vane. The results show that the field synergy principle is feasible to explain the mechanism of heat transfer enhancement in a fuel assembly. The enhancement in subchannels is more effective than on the rod's surface. If the pressure loss is ignored, the performance of the split mixing vane is superior to the separate mixing vane based on the enhanced heat transfer. Increasing the blending angle of the split mixing vane improves heat transfer enhancement, the maximum of which is 7.1%. Increasing the blending angle of the separate mixing vane did not significantly enhance heat transfer in the rod bundle, and even prevented heat transfer at a blending angle of 50°. This finding testifies to the feasibility of predicting heat transfer in a rod bundle with a spacer grid by field synergy, and upon comparison with analyzed flow features only, the field synergy method may provide more accurate guidance for optimizing the use of mixing vanes.

Hot spot detection system for vanes or blades of a combustion turbine

DOEpatents

Twerdochlib, M.

1999-02-02

This invention includes a detection system that can determine if a turbine component, such as a turbine vane or blade, has exceeded a critical temperature, such as a melting point, along any point along the entire surface of the vane or blade. This system can be employed in a conventional combustion turbine having a compressor, a combustor and a turbine section. Included within this system is a chemical coating disposed along the entire interior surface of a vane or blade and a closed loop cooling system that circulates a coolant through the interior of the vane or blade. If the temperature of the vane or blade exceeds a critical temperature, the chemical coating will be expelled from the vane or blade into the coolant. Since while traversing the closed loop cooling system the coolant passes through a detector, the presence of the chemical coating in the coolant will be sensed by the system. If the chemical coating is detected, this indicates that the vane or blade has exceeded a critical temperature. 5 figs.

Hot spot detection system for vanes or blades of a combustion turbine

DOEpatents

Twerdochlib, Michael

1999-01-01

This invention includes a detection system that can determine if a turbine component, such as a turbine vane or blade, has exceeded a critical temperature, such as a melting point, along any point along the entire surface of the vane or blade. This system can be employed in a conventional combustion turbine having a compressor, a combustor and a turbine section. Included within this system is a chemical coating disposed along the entire interior surface of a vane or blade and a closed loop cooling system that circulates a coolant through the interior of the vane or blade. If the temperature of the vane or blade exceeds a critical temperature, the chemical coating will be expelled from the vane or blade into the coolant. Since while traversing the closed loop cooling system the coolant passes through a detector, the presence of the chemical coating in the coolant will be sensed by the system. If the chemical coating is detected, this indicates that the vane or blade has exceeded a critical temperature.

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

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

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

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

NASA Technical Reports Server (NTRS)

Ghaffari, F.; Chaturvedi, S. K.

1984-01-01

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

Reduction of Unsteady Forcing in a Vaned, Contra-Rotating Transonic Turbine Configuration

NASA Technical Reports Server (NTRS)

Clark, John

2010-01-01

HPT blade unsteadiness in the presence of a downstream vane consistent with contra-rotation is characterized by strong interaction at the first harmonic of downstream vane passing. E An existing stage-and-one-half transonic turbine rig design was used as a baseline to investigate means of reducing such a blade-vane interaction. E Methods assessed included: Aerodynamic shaping of HPT blades 3D stacking of the downstream vane Steady pressure-side blowing E Of the methods assessed, a combination of vane bowing and steady pressure-side blowing produced the most favorable result. E Transonic turbine experiments are planned to assess predictive accuracy for the baseline turbine and any design improvements.

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

USGS Publications Warehouse

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

2015-01-01

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

Flexible Plug Repair for Shuttle Wing Leading Edge

NASA Technical Reports Server (NTRS)

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

Application of composite materials to turbofan engine fan exit guide vanes

NASA Technical Reports Server (NTRS)

Smith, G. T.

1980-01-01

A program was conducted by NASA with the JT9D engine manufacturer to develop a lightweight, cost effective, composite material fan exit guide vane design having satisfactory structural durability for commerical engine use. Based on the results of a previous company supported program, eight graphite/epoxy and graphite-glass/epoxy guide vane designs were evaluated and four were selected for fabrication and testing. Two commercial fabricators each fabricated 13 vanes. Fatigue tests were used to qualify the selected design configurations under nominally dry, 38 C (100 F) and fully wet and 60 C (140 F) environmental conditions. Cost estimates for a production rate of 1000 vanes per month ranged from 1.7 to 2.6 times the cost of an all aluminum vane. This cost is 50 to 80 percent less than the initial program target cost ratio which was 3 times the cost of an aluminum vane. Application to the JT9D commercial engine is projected to provide a weight savings of 236 N (53 lb) per engine.

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

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

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

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2002-01-01

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

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

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2002-01-01

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

Analysis of airfoil leading edge separation bubbles

NASA Technical Reports Server (NTRS)

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

1982-01-01

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

Numerical Simulations of Vortex Generator Vanes and Jets on a Flat Plate

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Yao, Chung-Sheng; Lin, John C.

2002-01-01

Numerical simulations of a single low-profile vortex generator vane, which is only a small fraction of the boundary-layer thickness, and a vortex generating jet have been performed for flows over a flat plate. The numerical simulations were computed by solving the steady-state solution to the Reynolds-averaged Navier-Stokes equations. The vortex generating vane results were evaluated by comparing the strength and trajectory of the streamwise vortex to experimental particle image velocimetry measurements. From the numerical simulations of the vane case, it was observed that the Shear-Stress Transport (SST) turbulence model resulted in a better prediction of the streamwise peak vorticity and trajectory when compared to the Spalart-Allmaras (SA) turbulence model. It is shown in this investigation that the estimation of the turbulent eddy viscosity near the vortex core, for both the vane and jet simulations, was higher for the SA model when compared to the SST model. Even though the numerical simulations of the vortex generating vane were able to predict the trajectory of the stream-wise vortex, the initial magnitude and decay of the peak streamwise vorticity were significantly under predicted. A comparison of the positive circulation associated with the streamwise vortex showed that while the numerical simulations produced a more diffused vortex, the vortex strength compared very well to the experimental observations. A grid resolution study for the vortex generating vane was also performed showing that the diffusion of the vortex was not a result of insufficient grid resolution. Comparisons were also made between a fully modeled trapezoidal vane with finite thickness to a simply modeled rectangular thin vane. The comparisons showed that the simply modeled rectangular vane produced a streamwise vortex which had a strength and trajectory very similar to the fully modeled trapezoidal vane.

Unsteady Flow Simulations in Support of the SSME HEX Turning Vane Cracking Investigation with the ATD HPOTP

NASA Technical Reports Server (NTRS)

Dougherty, N. S.; Burnette, D. W.; Holt, J. B.; Nesman, T.

1993-01-01

Unsteady flow computations are being performed with the P&W (ATD) and the Rocketdyne baseline configurations of the SSME LO2 turbine turnaround duct (TAD) and heat exchanger (HEX). The work is in support of the HEX inner turning vane cracking investigation. Fatigue cracking has occurred during hot firings with the P&W configuration on the HEX inner vane, and it appears the fix will involve changes to the TAD splitter vane position and to the TAD inner wall curvature to reduce the dynamic loading on the inner vane. Unsteady flow computations on the P&W baseline and fix and on the Rocketdyne baseline reference follow steady-flow screening computations done by MSFC/ED32 on several trial configurations arriving at the fix. The P&W TAD inlet velocity profile has a strong radial velocity component that directs the flow toward the inner wall and raises the local velocity a factor of two and the dynamic pressure a factor, of four. The fix is intended to redistribute the flow more evenly across the HEX inner and outer vanes like the Rocketdyne baseline reference. Vane buffeting at frequencies around 4,000 Hz is the leading suspected cause of the problem. Our simulations (work in progress) are being done with the USA 2D axisymmetric code approximating the flow as axisymmetric u+v 2D (axial, u, and radial, v, components only). The HEX coils are included in the model to make sure the fix does not adversely affect the HEX environment. Turbulent kinetic energy, k, levels where k = 1/2 v' rms2 are locally as high as 10,000 ft2/sec2 for the P&W baseline at the engine interface (between the TAD and HEX) at the HEX inner vane location. However, k is less than 8,000 on the HEX outer vane and only about 4,500 on the HEX inner vane for the Rocketdyne baseline. Unsteady turbulence intensity, v'rms/v, and pressure, p', are being computed in the present computations to compare with steady-flow Reynolds-averaged computations where p'rms = const (pk) for overall rms random turbulence from 0

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

NASA Astrophysics Data System (ADS)

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

2017-10-01

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

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

NASA Astrophysics Data System (ADS)

Mathews, J. R.; Peake, N.

2018-05-01

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

Low cost fabrication development for oxide dispersion strengthened alloy vanes

NASA Technical Reports Server (NTRS)

Perkins, R. J.; Bailey, P. G.

1978-01-01

Viable processes were developed for secondary working of oxide dispersion strengthened (ODS) alloys to near-net shapes (NNS) for aircraft turbine vanes. These processes were shown capable of producing required microstructure and properties for vane applications. Material cost savings of 40 to 50% are projected for the NNS process over the current procedures which involve machining from rectangular bar. Additional machining cost savings are projected. Of three secondary working processes evaluated, directional forging and plate bending were determined to be viable NNS processes for ODS vanes. Directional forging was deemed most applicable to high pressure turbine (HPT) vanes with their large thickness variations while plate bending was determined to be most cost effective for low pressure turbine (LPT) vanes because of their limited thickness variations. Since the F101 LPT vane was selected for study in this program, development of plate bending was carried through to establishment of a preliminary process. Preparation of ODS alloy plate for bending was found to be a straight forward process using currently available bar stock, providing that the capability for reheating between roll passes is available. Advanced ODS-NiCrAl and ODS-FeCrAl alloys were utilized on this program. Workability of all alloys was adequate for directional forging and plate bending, but only the ODS-FeCrAl had adequate workability for shaped preform extrustion.

Streakline flow visualization study of a horseshoe vortex in a large-scale, two-dimensional turbine stator cascade

NASA Technical Reports Server (NTRS)

Gaugler, R. E.; Russell, L. M.

1979-01-01

Neutrally bouyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large scale, two dimensional, turbine stator cascade. Inlet Reynolds number, based on true chord, ranged between 100,000 to 300,000. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16 mm movie film. Individual frames from the film were selected, and overlayed to show the details of the horseshoe vortex around the leading edge. The transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

Flow behavior in inlet guide vanes of radial turbines

NASA Technical Reports Server (NTRS)

Sokhey, J.; Tabakoff, W.; Hosny, W. M.

1975-01-01

Scroll flow is discussed. Streamline pattern and velocity distribution in the guide vanes are calculated. The blade surface temperature distribution is also determined. The effects of the blade shapes and the nozzle channel width on the velocity profiles at inlet to the guide vanes are investigated.

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

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

NASA Technical Reports Server (NTRS)

Camarda, Charles J.; Glass, David E.

1992-01-01

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

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

NASA Astrophysics Data System (ADS)

Camarda, Charles J.; Glass, David E.

1992-10-01

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

A lightweight, biological structure with tailored stiffness: The feather vane.

PubMed

Sullivan, Tarah N; Pissarenko, Andreï; Herrera, Steven A; Kisailus, David; Lubarda, Vlado A; Meyers, Marc A

2016-09-01

The flying feathers of birds are keratinous appendages designed for maximum performance with a minimum weight penalty. Thus, their design contains ingenious combinations of components that optimize lift, stiffness, aerodynamics, and damage resistance. This design involves two main parts: a central shaft that prescribes stiffness and lateral vanes which allows for the capture of air. Within the feather vane, barbs branch from the shaft and barbules branch from barbs, forming a flat surface which ensures lift. Microhooks at the end of barbules hold barbs tightly together, providing the close-knit, unified structure of the feather vane and enabling a repair of the structure through the reattachment of un-hooked junctions. Both the shaft and barbs are lightweight biological structures constructed of keratin using the common motif of a solid shell and cellular interior. The cellular core increases the resistance to buckling with little added weight. Here we analyze the detailed structure of the feather barb and, for the first time, explain its flexural stiffness in terms of the mechanics of asymmetric foam-filled beams subjected to bending. The results are correlated and validated with finite element modeling. We compare the flexure of single barbs as well as arrays of barbs and find that the interlocking adherence of barbs to one another enables a more robust structure due to minimized barb rotation during deflection. Thus, the flexure behavior of the feather vane can be tailored by the adhesive hooking between barbs, creating a system that mitigates damage. A simplified three-dimensional physical model for this interlocking mechanism is constructed by additive manufacturing. The exceptional architecture of the feather vane will motivate the design of bioinspired structures with tailored and unique properties ranging from adhesives to aerospace materials. Despite its importance to bird flight, literature characterizing the feather vane is extremely limited. The feather

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.

Ceramic Matrix Composite Vane Subelement Burst Testing

NASA Technical Reports Server (NTRS)

Brewer, David N.; Verrilli, Michael; Calomino, Anthony

2006-01-01

Burst tests were performed on Ceramic Matrix Composite (CMC) vane specimens, manufactured by two vendors, under the Ultra Efficient Engine Technology (UEET) project. Burst specimens were machined from the ends of 76mm long vane sub-elements blanks and from High Pressure Burner Rig (HPBR) tested specimens. The results of burst tests will be used to compare virgin specimens with specimens that have had an Environmental Barrier Coating (EBC) applied, both HPBR tested and untested, as well as a comparison between vendors.

Lesson from Tungsten Leading Edge Heat Load Analysis in KSTAR Divertor

NASA Astrophysics Data System (ADS)

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

2016-10-01

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

Controlled vortical flow on delta wings through unsteady leading edge blowing

NASA Technical Reports Server (NTRS)

Lee, K. T.; Roberts, Leonard

1990-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

Numerical investigation of performance of vane-type propellant management device by VOF methods

NASA Astrophysics Data System (ADS)

Liu, J. T.; Zhou, C.; Wu, Y. L.; Zhuang, B. T.; Li, Y.

2015-01-01

The orbital propellant management performance of the vane-type tank is so important for the propellant system and it determines the lifetime of the satellite. The propellant in the tank can be extruded by helium gas. To study the two phase distribution in the vane-type surface tension tank and the capability of the vane-type propellant management device (PMD), a large volume vane-type surface tension tank is analysed using 3-D unsteady numerical simulations. VOF methods are used to analyse the location of the interface of the two phase. Performances of the propellant acquisition vanes and propellant refillable reservoir in the tank are investigated. The flow conductivity of the propellant acquisition vanes and the liquid storage capacity of propellant refillable reservoir can be affected by the value of the gravity and the volume of the propellant in the tank. To avoid the large resistance causing by surface tension in an outflow of a small hole, the design of the vanes in a propellant refillable reservoir should have suitable space.

Optimal Micro-Vane Flow Control for Compact Air Vehicle Inlets

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Miller, Daniel N.; Addington, Gregory A.; Agrell, Johan

2004-01-01

The purpose of this study on micro-vane secondary flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to optimally design micro-vane secondary flow control arrays, and to establish that the aeromechanical effects of engine face distortion can also be included in the design and optimization process. These statistical design concepts were used to investigate the design characteristics of "low unit strength" micro-effector arrays. "Low unit strength" micro-effectors are micro-vanes set at very low angles-of-incidence with very long chord lengths. They were designed to influence the near wall inlet flow over an extended streamwise distance, and their advantage lies in low total pressure loss and high effectiveness in managing engine face distortion. Therefore, this report examines optimal micro-vane secondary flow control array designs for compact inlets through a Response Surface Methodology.

Numerical modelling of multi-vane expander operating conditions in ORC system

NASA Astrophysics Data System (ADS)

Rak, Józef; Błasiak, Przemysław; Kolasiński, Piotr

2017-11-01

Multi-vane expanders are positive displacement volumetric machines which are nowadays considered for application in micro-power domestic ORC systems as promising alternative to micro turbines and other volumetric expanders. The multi-vane expander features very simple design, low gas flow capacity, low expansion ratios, an advantageous ratio of the power output to the external dimensions and are insensitive to the negative influence of the gas-liquid mixture expansion. Moreover, the multi-vane expander can be easily hermetically sealed, which is one of the key issues in the ORC system design. A literature review indicates that issues concerning the application of multi-vane expanders in such systems, especially related to operating of multi-vane expander with different low-boiling working fluids, are innovative, not fully scientifically described and have the potential for practical implementation. In this paper the results of numerical investigations on multi-vane expander operating conditions are presented. The analyses were performed on three-dimensional numerical model of the expander in ANSYS CFX software. The numerical model of the expander was validated using the data obtained from the experiment carried out on a lab test-stand. Then a series of computational analysis were performed using expanders' numerical model in order to determine its operating conditions under various flow conditions of different working fluids.

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

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

1980-01-01

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

Pulsed Film Cooling on a Turbine Blade Leading Edge

DTIC Science & Technology

2009-09-01

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

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

Forced responses on a radial turbine with nozzle guide vanes

NASA Astrophysics Data System (ADS)

Liu, Yixiong; Yang, Ce; Ma, Chaochen; Lao, DaZhong

2014-04-01

Radial turbines with nozzle guide vanes are widely used in various size turbochargers. However, due to the interferences with guide vanes, the blades of impellers are exposed to intense unsteady aerodynamic excitations, which cause blade vibrations and lead to high cycle failures (HCF). Moreover, the harmonic resonance in some frequency regions are unavoidable due to the wide operation conditions. Aiming to achieve a detail insight into vibration characteristics of radial flow turbine, a numerical method based on fluid structure interaction (FSI) is presented. Firstly, the unsteady aerodynamic loads are determined by computational fluid dynamics (CFD). And the fluctuating pressures are transformed from time domain to frequency domain by fast Fourier-transform (FFT). Then, the entire rotor model is adopted to analyze frequencies and mode shapes considering mistuning in finite element (FE) method. Meanwhile, harmonic analyses, applying the pressure fluctuation from CFD, are conducted to investigate the impeller vibration behavior and blade forced response in frequency domain. The prediction of the vibration dynamic stress shows acceptable agreement to the blade actual damage in consistent tendency.

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

Simulation and Optimization of an Airfoil with Leading Edge Slat

NASA Astrophysics Data System (ADS)

Schramm, Matthias; Stoevesandt, Bernhard; Peinke, Joachim

2016-09-01

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

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

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

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

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

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

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

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

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

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

Impact of Separation Distance on Multi-Vane Radiometer Configurations

NASA Astrophysics Data System (ADS)

Cornella, B. M.; Ketsdever, A. D.; Gimelshein, N. E.; Gimelshein, S. F.

2011-05-01

The radiometric force produced by a linear array of three radiometer vanes has been assessed numerically using an argon carrier gas and experimentally using air. The separation distance between the three vanes of the array was varied between 0 and 120 percent based on the height of an individual radiometer vane of 40 mm. Qualitative agreement between the numerical and experimental results is shown as a function of operating Knudsen number, vane separation distance, and surrounding chamber geometry. Both sets of results indicate an asymptotic trend in maximum force as the separation distance increases as well as a shift in the maximum force Knudsen number. Small chamber effects for both numerical and experimental results indicate an increase of the total force ranging from a factor of 2.5 to 4. Quantitatively, however, the numerical simulations yield forces approximately an order of magnitude higher than observed in the experiments due to differences in carrier gas and accommodation coefficient as well as the two dimensional nature of the numerical simulations versus the three dimensional experiment.

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.

Flow measurements in two cambered vane diffusers with different passage widths

NASA Astrophysics Data System (ADS)

Stein, W.; Rautenberg, M.

1985-03-01

To investigate the influence of the vaneless space between impeller exit and the diffuser vanes, detailed flow measurements in two diffusers with the same vane geometry but different passage width are compared. The three-dimensional character of the flow changes between impeller exit and the entry to the two dimensional vanes depending on the shape of the shroud. After initial measurements with a constant area vaneless space, the width of the vaned diffuser was later on reduced by 10 percent. The compressor maps show increases in overall pressure rise and efficiency with the width reduction. To get further details of the flow field, measurements of the static pressure distribution at hub and shroud have been performed at several operation points for both diffusers. At the same points, the flow angle and total pressure distribution between hub and shroud upstream and downstream of the vanes have been measured with probes. The maximum efficiency of the narrow diffuser is nearly 2 percent higher than for the wide diffuser. The measurements give further details to explain this improvement.

Numerical analysis and experiment research on fluid orbital performance of vane type propellant management device

NASA Astrophysics Data System (ADS)

Hu, Q.; Li, Y.; Pan, H. L.; Liu, J. T.; Zhuang, B. T.

2015-01-01

Vane type propellant management device (PMD) is one of the key components of the vane-type surface tension tank (STT), and its fluid orbital performance directly determines the STT's success or failure. In present paper, numerical analysis and microgravity experiment study on fluid orbital performance of a vane type PMD were carried out. By using two-phase flow model of volume of fluid (VOF), fluid flow characteristics in the tank with the vane type PMD were numerically calculated, and the rules of fluid transfer and distribution were gotten. A abbreviate model test system of the vane type PMD is established and microgravity drop tower tests were performed, then fluid management and transmission rules of the vane type PMD were obtained under microgravity environment. The analysis and tests results show that the vane type PMD has good and initiative fluid orbital management ability and meets the demands of fluid orbital extrusion in the vane type STT. The results offer valuable guidance for the design and optimization of the new generation of vane type PMD, and also provide a new approach for fluid management and control in space environment.

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.

Fluid-structure coupling in the guide vanes cascade of a pump-turbine scale model

NASA Astrophysics Data System (ADS)

Roth, S.; Hasmatuchi, V.; Botero, F.; Farhat, M.; Avellan, F.

2010-08-01

The present study concerns fluid-structure coupling phenomena occurring in a guide vane cascade of a pump-turbine scale model placed in the EPFL PF3 test rig. An advanced instrument set is used to monitor both vibrating structures and the surrounding flow. The paper highlights the interaction between vibrating guide vanes and the flow behavior. The pressure fluctuations in the stay vanes region are found to be strongly influenced by the amplitude of the vibrating guide vanes. Moreover, the flow induces different hydrodynamic damping on the vibrating guide vanes depending on the operating point of the pump-turbine.

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

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

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

NASA Technical Reports Server (NTRS)

1979-01-01

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

Feather seal slot for vanes

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

Del Mastro, B. P.; Eckart, F.

1985-10-22

The slots for accommodating feather seals in the turbine vanes of a gas turbine engine has the end thereof sealed off by use of weld wire inserted into the slot and simultaneously welded and cut to the required length.

JT90 thermal barrier coated vanes

NASA Technical Reports Server (NTRS)

Sheffler, K. D.; Graziani, R. A.; Sinko, G. C.

1982-01-01

The technology of plasma sprayed thermal barrier coatings applied to turbine vane platforms in modern high temperature commercial engines was advanced to the point of demonstrated feasibility for application to commercial aircraft engines. The three thermal barrier coatings refined under this program are zirconia stabilized with twenty-one percent magnesia (21% MSZ), six percent yttria (6% YSZ), and twenty percent yttria (20% YSZ). Improvement in thermal cyclic endurance by a factor of 40 times was demonstrated in rig tests. A cooling system evolved during the program which featured air impingement cooling for the vane platforms rather than film cooling. The impingement cooling system, in combination with the thermal barrier coatings, reduced platform cooling air requirements by 44% relative to the current film cooling system. Improved durability and reduced cooling air requirements were demonstrated in rig and engine endurance tests. Two engine tests were conducted, one of 1000 cycles and the other of 1500 cycles. All three coatings applied to vanes fabricated with the final cooling system configuration completed the final 1500 cycle engine endurance test. Results of this test clearly demonstrated the durability of the 6% YSZ coating which was in very good condition after the test. The 21% MSZ and 20% YSZ coatings had numerous occurrences of significant spalling in the test.

Active Flow Separation Control of a Stator Vane Using Surface Injection in a Multistage Compressor Experiment

NASA Technical Reports Server (NTRS)

Culley, Dennis E.; Bright, Michelle M.; Prahst, Patricia S.; Strazisar, Anthony J.

2003-01-01

Micro-flow control actuation embedded in a stator vane was used to successfully control separation and improve near stall performance in a multistage compressor rig at NASA Glenn. Using specially designed stator vanes configured with internal actuation to deliver pulsating air through slots along the suction surface, a research study was performed to identify performance benefits using this microflow control approach. Pressure profiles and unsteady pressure measurements along the blade surface and at the shroud provided a dynamic look at the compressor during microflow air injection. These pressure measurements lead to a tracking algorithm to identify the onset of separation. The testing included steady air injection at various slot locations along the vane. The research also examined the benefit of pulsed injection and actively controlled air injection along the stator vane. Two types of actuation schemes were studied, including an embedded actuator for on-blade control. Successful application of an online detection and flow control scheme will be discussed. Testing showed dramatic performance benefit for flow reattachment and subsequent improvement in diffusion through the use of pulsed controlled injection. The paper will discuss the experimental setup, the blade configurations, and preliminary CFD results which guided the slot location along the blade. The paper will also show the pressure profiles and unsteady pressure measurements used to track flow control enhancement, and will conclude with the tracking algorithm for adjusting the control.

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

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2017-12-01

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

Solar updraft power generator with radial and curved vanes

NASA Astrophysics Data System (ADS)

Hafizh, Hadyan; Hamsan, Raziff; Zamri, Aidil Azlan Ahmad; Keprawi, Mohamad Fairuz Mohamad; Shirato, Hiromichi

2018-02-01

Solar radiation is the largest source of energy available on earth and the solar updraft power generator (SUPG) is a renewable energy facility capable of harnessing its abundant power. Unlike the conventional wind turbines that harness natural wind in the atmosphere and often encounter with the intermittent issue or even complete cut-off from airflow, the SUPG creates artificial wind as a result of solar-induced convective flows. However, the SUPG has an inherent low total efficiency due to the conversion of thermal energy into pressure energy. Acknowledging the low efficiency and considering its potential as a renewable energy facility, the current work aims to increase the total efficiency by installing a series of guide walls inside the collector. Two types of guide walls were used i.e. radial and curved vanes. The result with curved vanes showed that the updraft velocity is higher compare to those without vanes. About 18% and 64% improvement of updraft velocity and mechanical power were attained respectively. Furthermore, it was observed that the role of radial vanes configuration was more to produce a smooth updraft velocity profile rather than increasing the total efficiency.

Sequential cooling insert for turbine stator vane

DOEpatents

Jones, Russell B.; Krueger, Judson J.; Plank, William L.

2014-04-01

A sequential impingement cooling insert for a turbine stator vane that forms a double impingement for the pressure and suction sides of the vane or a triple impingement. The insert is formed from a sheet metal formed in a zigzag shape that forms a series of alternating impingement cooling channels with return air channels, where pressure side and suction side impingement cooling plates are secured over the zigzag shaped main piece. Another embodiment includes the insert formed from one or two blocks of material in which the impingement channels and return air channels are machined into each block.

Sequential cooling insert for turbine stator vane

DOEpatents

Jones, Russel B; Krueger, Judson J; Plank, William L

2014-11-04

A sequential impingement cooling insert for a turbine stator vane that forms a double impingement for the pressure and suction sides of the vane or a triple impingement. The insert is formed from a sheet metal formed in a zigzag shape that forms a series of alternating impingement cooling channels with return air channels, where pressure side and suction side impingement cooling plates are secured over the zigzag shaped main piece. Another embodiment includes the insert formed from one or two blocks of material in which the impingement channels and return air channels are machined into each block.

Water table tests of proposed heat transfer tunnels for small turbine vanes

NASA Technical Reports Server (NTRS)

Meitner, P. L.

1974-01-01

Water-table flow tests were conducted for proposed heat-transfer tunnels which were designed to provide uniform flow into their respective test sections of a single core engine turbine vane and a full annular ring of helicopter turbine vanes. Water-table tests were also performed for the single-vane test section of the core engine tunnel. The flow in the heat-transfer tunnels was shown to be acceptable.

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

NASA Technical Reports Server (NTRS)

Luckring, James M.; Boelens, Okko J.

2015-01-01

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

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

NASA Technical Reports Server (NTRS)

Kendall, James M.

1991-01-01

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

Testing of Composite Fan Vanes With Erosion-Resistant Coating Accelerated

NASA Technical Reports Server (NTRS)

Bowman, Cheryl L.; Sutter, James K.; Otten, Kim D.; Samorezov, Sergey; Perusek, Gail P.

2004-01-01

The high-cycle fatigue of composite stator vanes provided an accelerated life-state prior to insertion in a test stand engine. The accelerated testing was performed in the Structural Dynamics Laboratory at the NASA Glenn Research Center under the guidance of Structural Mechanics and Dynamics Branch personnel. Previous research on fixturing and test procedures developed at Glenn determined that engine vibratory conditions could be simulated for polymer matrix composite vanes by using the excitation of a combined slip table and electrodynamic shaker in Glenn's Structural Dynamics Laboratory. Bench-top testing gave researchers the confidence to test the coated vanes in a full-scale engine test.

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2004-01-01

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

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

NASA Astrophysics Data System (ADS)

Pan, Shaowu; Gao, Zhenxun; Lee, Chunhian

2014-12-01

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

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

PubMed

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

2017-12-01

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

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.

A Method for Computing Leading-Edge Loads

NASA Technical Reports Server (NTRS)

Rhode, Richard V; Pearson, Henry A

1933-01-01

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

An Aeroacoustic Study of a Leading Edge Slat Configuration

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

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

NASA Technical Reports Server (NTRS)

Hahne, Daniel E.

1999-01-01

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

Theoretical morphology and development of flight feather vane asymmetry with experimental tests in parrots.

PubMed

Feo, Teresa J; Prum, Richard O

2014-06-01

Asymmetry in flight feather vane width is a major functional innovation associated with the evolution of flight in the ancestors of birds. However, the developmental and morphological basis of feather shape is not simple, and the developmental processes involved in vane width asymmetry are poorly understood. We present a theoretical model of feather morphology and development that describes the possible ways to modify feather development and produce vane asymmetry. Our model finds that the theoretical morphospace of feather shape is redundant, and that many different combinations of parameters could be responsible for vane asymmetry in a given feather. Next, we empirically measured morphological and developmental model parameters in asymmetric and symmetric feathers from two species of parrots to identify which combinations of parameters create vane asymmetry in real feathers. We found that both longer barbs, and larger barb angles in the relatively wider trailing vane drove asymmetry in tail feathers. Developmentally, longer barbs were the result of an offset of the radial position of the new barb locus, whereas larger barb angles were produced by differential expansion of barbs as the feather unfurls from the tubular feather germ. In contrast, the helical angle of barb ridge development did not contribute to vane asymmetry and could be indicative of a constraint. This research provides the first comprehensive description of both the morphological and developmental modifications responsible for vane asymmetry within real feathers, and identifies key steps that must have occurred during the evolution of vane asymmetry. © 2014 Wiley Periodicals, Inc.

Programmable controller system for wind tunnel diversion vanes

NASA Technical Reports Server (NTRS)

King, R. F.

1982-01-01

A programmable controller (PC) system automatic sequence control, which acts as a supervisory controller for the servos, selects the proper drives, and automatically sequences the vanes, was developed for use in a subsonic wind tunnel. Tunnel modifications include a new second test section (80 ft x 100 ft with a maximum air speed capability of 110 knots) and an increase in maximum velocity flow from 200 knots to 300 knots. A completely automatic sequence control is necessary in order to allow intricate motion of the 14 triangularly arranged vanes which can be as large as 70 ft high x 35 ft wide and which require precise acceleration and deceleration control. Rate servos on each drive aid in this control, and servo cost was minimized by using four silicon controlled rectifier controllers to control the 20 dc drives. The PC has a programming capacity which facilitated the implementation of extensive logic design. A series of diagrams sequencing the vanes and a block diagram of the system are included.

Rotor-generated unsteady aerodynamic interactions in a 1½ stage compressor

NASA Astrophysics Data System (ADS)

Papalia, John J.

Because High Cycle Fatigue (HCF) remains the predominant surprise failure mode in gas turbine engines, HCF avoidance design systems are utilized to identify possible failures early in the engine development process. A key requirement of these analyses is accurate determination of the aerodynamic forcing function and corresponding airfoil unsteady response. The current study expands the limited experimental database of blade row interactions necessary for calibration of predictive HCF analyses, with transonic axial-flow compressors of particular interest due to the presence of rotor leading edge shocks. The majority of HCF failures in aircraft engines occur at off-design operating conditions. Therefore, experiments focused on rotor-IGV interactions at off-design are conducted in the Purdue Transonic Research Compressor. The rotor-generated IGV unsteady aerodynamics are quantified when the IGV reset angle causes the vane trailing edge to be nearly aligned with the rotor leading edge shocks. A significant vane response to the impulsive static pressure perturbation associated with a shock is evident in the point measurements at 90% span, with details of this complex interaction revealed in the corresponding time-variant vane-to-vane flow field data. Industry wide implementation of Controlled Diffusion Airfoils (CDA) in modern compressors motivated an investigation of upstream propagating CDA rotor-generated forcing functions. Whole field velocity measurements in the reconfigured Purdue Transonic Research Compressor along the design speedline reveal steady loading had a considerable effect on the rotor shock structure. A detached rotor leading edge shock exists at low loading, with an attached leading edge and mid-chord suction surface normal shock present at nominal loading. These CDA forcing functions are 3--4 times smaller than those generated by the baseline NACA 65 rotor at their respective operating points. However, the IGV unsteady aerodynamic response to the CDA

Interaction of impeller and guide vane in a series-designed axial-flow pump

NASA Astrophysics Data System (ADS)

Kim, S.; Choi, Y. S.; Lee, K. Y.; Kim, J. H.

2012-11-01

In this paper, the interaction of the impeller and guide vane in a series-designed axial-flow pump was examined through the implementation of a commercial CFD code. The impeller series design refers to the general design procedure of the base impeller shape which must satisfy the various flow rate and head requirements by changing the impeller setting angle and number of blades of the base impeller. An arc type meridional shape was used to keep the meridional shape of the hub and shroud with various impeller setting angles. The blade angle and the thickness distribution of the impeller were designed as an NACA airfoil type. In the design of the guide vane, it was necessary to consider the outlet flow condition of the impeller with the given setting angle. The meridional shape of the guide vane were designed taking into consideration the setting angle of the impeller, and the blade angle distribution of the guide vane was determined with a traditional design method using vane plane development. In order to achieve the optimum impeller design and guide vane, three-dimensional computational fluid dynamics and the DOE method were applied. The interaction between the impeller and guide vane with different combination set of impeller setting angles and number of impeller blades was addressed by analyzing the flow field of the computational results.

Analysis of Multi-Vane Radiometers in High-Altitude Propulsion

DTIC Science & Technology

2010-06-17

m] I. Introduction EAR space is generally defined as the altitude region above where jet aircraft can produce enough lift to maintain level...temperature gradient [5]. To maximize the gradient, the hot surface and cold surface are separated by thermal insulator, in this case aerogel . The hot...of the multi-vanes are vehicle specific base on altitude. Each individual vane consists of silicone aerogel sandwiched between two thin aluminum

Observations on Leading-Edge Vortex Development

NASA Astrophysics Data System (ADS)

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

2016-11-01

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

Predicting broadband noise from a stator vane of a gas turbine engine

NASA Technical Reports Server (NTRS)

Hanson, Donald B. (Inventor)

2002-01-01

A computer-implemented model of fan section of a gas turbine engine accounts for the turbulence in the gas flow emanating from the rotor assembly and impinging upon an inlet to the stator vane cascade. The model allows for user-input variations in the sweep and/or lean angles for the stator vanes. The model determines the resulting acoustic response of the fan section as a function of the turbulence and the lean and/or sweep angles of the vanes. The model may be embodied in software that is rapidly executed in a computer. This way, an optimum arrangement in terms of fan noise reduction is quickly determined for the stator vane lean and sweep physical positioning in the fan section of a gas turbine engine.

Performance improvement of a centrifugal compressor stage by using different vaned diffusers

NASA Astrophysics Data System (ADS)

Zhang, Y. C.; Kong, X. Z.; Li, F.; Sun, W.; Chen, Q. G.

2013-12-01

The vaneless diffuser (VLD) is usually adopted in the traditional design of the multi-stage centrifugal compressor because of the stage's match problem. The drawback of the stage with vaneless diffusers is low efficiency. In order to increase the efficiency and at the same time, induce no significant decline in the operating range of the stage, three different types of vaned diffusers are designed and numerically investigated: the traditional vaned diffuser (TVD), the low-solidity cascade diffuser (LSD) and the partial-height vane diffuser (PVD). These three types of vaned diffusers have different influences on the performance of the centrifugal compressor. In the present investigation, the first part investigates the performance of a centrifugal compressor stage with three different vaned diffusers. The second part studies the influences of the height and the position of partial height vanes on the stage performance, and discusses the matching problem between the PVD and the downstream return channel. The stage investigated in this paper includes the impeller, the diffuser, the bend and the return channel. In the process of numerical investigation, the flow is assumed to be steady, and this process includes calculation and simulation. The calculation of 3-D turbulent flow in the stage uses the commercial CFD code NUMECA together with the Spalart-Allmaras turbulence model. The simulation of the computational region includes the impeller passages, the diffuser passages and return channel passages. The structure and surrounding region are assumed to have a perfect cyclic symmetry, so the single channel model and periodic boundary condition are applied at the middle of the passage, that is to reduce the calculation region to only one region. The investigation showed that the low-solidity cascade diffuser would be a better choice as a middle course for the first stage of the multistage centrifugal compressor. Besides, the influences of the height and the position of

47. DETAIL OF ORIGINAL VANE ASSEMBLY AND TWO WHEEL SECTIONS ...

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

47. DETAIL OF ORIGINAL VANE ASSEMBLY AND TWO WHEEL SECTIONS FROM ELI WINDMILLS, THE VANE SHEET BEARING STENCILED PAINTED INSCRIPTION, 'KREGEL WINDMILL CO. ELI NEBRASKA CITY, NEB.' VISIBLE IN THE IMAGE ARE BOTH SIDES OF THE WHEEL SECTIONS, SHOWING THE METHOD OF BLADE MOUNTING FOR ELI WINDMILLS. - Kregel Windmill Company Factory, 1416 Central Avenue, Nebraska City, Otoe County, NE

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

NASA Astrophysics Data System (ADS)

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

2015-05-01

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

Controlling Attitude of a Solar-Sail Spacecraft Using Vanes

NASA Technical Reports Server (NTRS)

Mettler, Edward; Acikmese, Ahmet; Ploen, Scott

2006-01-01

A paper discusses a concept for controlling the attitude and thrust vector of a three-axis stabilized Solar Sail spacecraft using only four single degree-of-freedom articulated spar-tip vanes. The vanes, at the corners of the sail, would be turned to commanded angles about the diagonals of the square sail. Commands would be generated by an adaptive controller that would track a given trajectory while rejecting effects of such disturbance torques as those attributable to offsets between the center of pressure on the sail and the center of mass. The controller would include a standard proportional + derivative part, a feedforward part, and a dynamic component that would act like a generalized integrator. The controller would globally track reference signals, and in the presence of such control-actuator constraints as saturation and delay, the controller would utilize strategies to cancel or reduce their effects. The control scheme would be embodied in a robust, nonlinear algorithm that would allocate torques among the vanes, always finding a stable solution arbitrarily close to the global optimum solution of the control effort allocation problem. The solution would include an acceptably small angle, slow limit-cycle oscillation of the vanes, while providing overall thrust vector pointing stability and performance.

An empirical comparison of primary baffle and vanes for optical astronomical telescope

NASA Astrophysics Data System (ADS)

Li, Taoran; Chen, Yingwei

2017-09-01

In optical astronomical telescopes, the primary baffle is a tube-like structure centering in the hole of the primary mirror and the vanes usually locate inside the baffle, improving the suppression of stray light. They are the most common methods of stray light control. To characterize the performance of primary baffle and vanes, an empirical comparison based on astronomical observations has been made with Xinglong 50cm telescope. Considering the convenience of switching, an independent vanes structure is designed, which can also improve the process of the primary mirror cooling and the air circulation. The comparison of two cases: (1) primary baffle plus vanes and (2) vanes alone involves in-dome and on-sky observations. Both the single star and the various off-axis angles of the stray light source observations are presented. The photometrical images are recorded by CCD to analyze the magnitude and the photometric error. The stray light uniformity of the image background derives from the reduction image which utilizes the MATLAB software to remove the stars. The in-dome experiments results reveal the effectiveness of primary baffle and the independent vanes structure. Meanwhile, the on-sky photometric data indicate there are little differences between them. The stray light uniformity has no difference when the angle between the star and the moon is greater than 20 degrees.

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.

Twisted Vanes Would Enhance Fuel/Air Mixing In Turbines

NASA Technical Reports Server (NTRS)

Nguyen, H. Lee; Micklow, Gerald J.; Dogra, Anju S.

1994-01-01

Computations of flow show performance of high-shear airblast fuel injector in gas-turbine engine enhanced by use of appropriately proportioned twisted (instead of flat) dome swirl vanes. Resultant more nearly uniform fuel/air mixture burns more efficiently, emitting smaller amounts of nitrogen oxides. Twisted-vane high-shear airblast injectors also incorporated into paint sprayers, providing advantages of low pressure drop characteristic of airblast injectors in general and finer atomization of advanced twisted-blade design.

Experimental studies on the aerodynamic performance and dynamic response of flow direction sensing vanes

NASA Technical Reports Server (NTRS)

Barna, P. S.; Crossman, G. R.

1976-01-01

Systematic investigations were performed on a variety of aerodynamic surfaces to obtain their potential for possible application to wind vanes. Among the surfaces tested were: (1) single vanes consisting of flat plates of various planforms having aspect ratios between 0.5 and 5; (2) bi-vanes with aspect ratio 2.5; (3) various cone and box vanes; and (4) various cruciform configurations. The models were subjected to windtunnel tests. In addition to lift and drag force measurements, damping and frequency test were performed under a variety of flow conditions.

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

NASA Astrophysics Data System (ADS)

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

2011-11-01

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

Centrifugal Compressor Surge Margin Improved With Diffuser Hub Surface Air Injection

NASA Technical Reports Server (NTRS)

Skoch, Gary J.

2002-01-01

Aerodynamic stability is an important parameter in the design of compressors for aircraft gas turbine engines. Compression system instabilities can cause compressor surge, which may lead to the loss of an aircraft. As a result, engine designers include a margin of safety between the operating line of the engine and the stability limit line of the compressor. The margin of safety is typically referred to as "surge margin." Achieving the highest possible level of surge margin while meeting design point performance objectives is the goal of the compressor designer. However, performance goals often must be compromised in order to achieve adequate levels of surge margin. Techniques to improve surge margin will permit more aggressive compressor designs. Centrifugal compressor surge margin improvement was demonstrated at the NASA Glenn Research Center by injecting air into the vaned diffuser of a 4:1-pressure-ratio centrifugal compressor. Tests were performed using injector nozzles located on the diffuser hub surface of a vane-island diffuser in the vaneless region between the impeller trailing edge and the diffuser-vane leading edge. The nozzle flow path and discharge shape were designed to produce an air stream that remained tangent to the hub surface as it traveled into the diffuser passage. Injector nozzles were located near the leading edge of 23 of the 24 diffuser vanes. One passage did not contain an injector so that instrumentation located in that passage would be preserved. Several orientations of the injected stream relative to the diffuser vane leading edge were tested over a range of injected flow rates. Only steady flow (nonpulsed) air injection was tested. At 100 percent of the design speed, a 15-percent improvement in the baseline surge margin was achieved with a nozzle orientation that produced a jet that was bisected by the diffuser vane leading edge. Other orientations also improved the baseline surge margin. Tests were conducted at speeds below the

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

PubMed

Uechi, Hiroyuki; Kuranaga, Erina

2017-08-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

PubMed Central

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

2004-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1999-01-01

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

Navier-Stokes analysis of airfoils with leading edge ice accretions

NASA Technical Reports Server (NTRS)

Potapczuk, Mark G.

1993-01-01

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

Methods and systems to enhance flame holding in a gas turbine engine

DOEpatents

Zuo, Baifang [Simpsonville, SC; Lacy, Benjamin Paul [Greer, SC; Stevenson, Christian Xavier [Inman, SC

2012-01-31

A fuel nozzle including a swirler assembly that includes a shroud, a hub, and a plurality of vanes extending between the shroud and the hub. Each vane includes a pressure sidewall and an opposite suction sidewall coupled to the pressure sidewall at a leading edge and at a trailing edge. At least one suction side fuel injection orifice is formed adjacent to the leading edge and extends from a first fuel supply passage to the suction sidewall. A fuel injection angle is oriented with respect to the suction sidewall. The suction side fuel injection orifice is configured to discharge fuel outward from the suction sidewall. At least one pressure side fuel injection orifice extends from a second fuel supply passage to the pressure sidewall and is substantially parallel to the trailing edge. The pressure side fuel injection orifice is configured to discharge fuel tangentially from the trailing edge.

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.

Axial vane-type swirler performance characteristics. M.S. Thesis

NASA Technical Reports Server (NTRS)

Sander, G. F.

1983-01-01

The performance of an axial vane-type swirler was investigated to aid in computer modeling of gas turbine combustor flowfields and in evaluation of turbulence models for swirling confined jet flow. The swirler studied is annular with a hub-to-swirler diameter ratio of 0.25 and ten adjustable vanes of pitch-to-chord ratio 0.68. Measurements of time-mean axial, radial, and tangential velocities were made at the swirler exit plane using a five-hole pitot probe technique with computer data reduction. Nondimensionalized velocities from both radial and azimuthal traverses are tabulated and plotted for a range of swirl vane angles phi from 0 to 70 degrees. A study was done of idealized exit-plane velocity profiles relating the swirl numbers S and S' to the ratio of maximum swirl and axial velocities for each idealized case, and comparing the idealized swirl numbers with ones calculated from measured profiles.

Effect of diffuser vane shape on the performance of a centrifugal compressor stage

NASA Astrophysics Data System (ADS)

Reddy, T. Ch Siva; Ramana Murty, G. V.; Prasad, M. V. S. S. S. M.

2014-04-01

The present paper reports the results of experimental investigations on the effect of diffuser vane shape on the performance of a centrifugal compressor stage. These studies were conducted on the chosen stage having a backward curved impeller of 500 mm tip diameter and 24.5 mm width and its design flow coefficient is ϕd=0.0535. Three different low solidity diffuser vane shapes namely uncambered aerofoil, constant thickness flat plate and circular arc cambered constant thickness plate were chosen as the variants for diffuser vane shape and all the three shapes have the same thickness to chord ratio (t/c=0.1). Flow coefficient, polytropic efficiency, total head coefficient, power coefficient and static pressure recovery coefficient were chosen as the parameters for evaluating the effect of diffuser vane shape on the stage performance. The results show that there is reasonable improvement in stage efficiency and total head coefficient with the use of the chosen diffuser vane shapes as compared to conventional vaneless diffuser. It is also noticed that the aero foil shaped LSD has shown better performance when compared to flat plate and circular arc profiles. The aerofoil vane shape of the diffuser blade is seen to be tolerant over a considerable range of incidence.

Probabilistic Analysis of a SiC/SiC Ceramic Matrix Composite Turbine Vane

NASA Technical Reports Server (NTRS)

Murthy, Pappu L. N.; Nemeth, Noel N.; Brewer, David N.; Mital, Subodh

2004-01-01

To demonstrate the advanced composite materials technology under development within the Ultra-Efficient Engine Technology (UEET) Program, it was planned to fabricate, test, and analyze a turbine vane made entirely of silicon carbide-fiber-reinforced silicon carbide matrix composite (SiC/SiC CMC) material. The objective was to utilize a five-harness satin weave melt-infiltrated (MI) SiC/SiC composite material developed under this program to design and fabricate a stator vane that can endure 1000 hours of engine service conditions. The vane was designed such that the expected maximum stresses were kept within the proportional limit strength of the material. Any violation of this design requirement was considered as the failure. This report presents results of a probabilistic analysis and reliability assessment of the vane. Probability of failure to meet the design requirements was computed. In the analysis, material properties, strength, and pressure loading were considered as random variables. The pressure loads were considered normally distributed with a nominal variation. A temperature profile on the vane was obtained by performing a computational fluid dynamics (CFD) analysis and was assumed to be deterministic. The results suggest that for the current vane design, the chance of not meeting design requirements is about 1.6 percent.

Wind tunnel performance results of swirl recovery vanes as tested with an advanced high speed propeller

NASA Technical Reports Server (NTRS)

Gazzaniga, John A.; Rose, Gayle E.

1992-01-01

Tests of swirl recovery vanes designed for use in conjunction with advanced high speed propellers were carried out at the NASA Lewis Research Center. The eight bladed 62.23 cm vanes were tested with a 62.23 cm SR = 7A high speed propeller in the NASA Lewis 2.44 x 1.83 m Supersonic Wind Tunnel for a Mach number range of 0.60 to 0.80. At the design operating condition for cruise of Mach 0.80 at an advance ratio of 3.26, the vane contribution to the total efficiency approached 2 percent. At lower off-design Mach numbers, the vane efficiency is even higher, approaching 4.5 percent for the Mach 0.60 condition. Use of the swirl recovery vanes essentially shifts the peak of the high speed propeller efficiency to a higher operating speed. This allows a greater degree of freedom in the selection of rpm over a wider operating range. Another unique result of the swirl recovery vane configuration is their essentially constant torque split between the propeller and the swirl vanes over a wide range of operating conditions for the design vane angle.

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

PubMed

Raza, Qanber; Jacobs, J Roger

2016-11-15

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

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

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

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

NASA Technical Reports Server (NTRS)

Ritzert, Frank; Nesbitt, James

2007-01-01

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

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

NASA Technical Reports Server (NTRS)

Carmichael, B. H.

1979-01-01

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

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

NASA Technical Reports Server (NTRS)

Etchberger, F. R.

1983-01-01

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

A Factor Affecting Transonic Leading-edge Flow Separation

NASA Technical Reports Server (NTRS)

Wood, George P; Gooderum, Paul B

1956-01-01

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

The Point of It All: Exploring Variations in Wind Vane Design

ERIC Educational Resources Information Center

Koballa, Thomas, Jr.

2008-01-01

A wind vane is a tool for making observations of wind direction and initiating inquiries about the weather. Its construction and use continue to be mainstays of the science education of elementary students. By providing students with the opportunity to discern critical features associated with the wind vane's operation, you can ensure that…

Mechanical support of a ceramic gas turbine vane ring

DOEpatents

Shi, Jun; Green, Kevin E.; Mosher, Daniel A.; Holowczak, John E.; Reinhardt, Gregory E.

2010-07-27

An assembly for mounting a ceramic turbine vane ring onto a turbine support casing comprises a first metal clamping ring and a second metal clamping ring. The first metal clamping ring is configured to engage with a first side of a tab member of the ceramic turbine vane ring. The second metal clamping ring is configured to engage with a second side of the tab member such that the tab member is disposed between the first and second metal clamping rings.

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.

Flow disturbance due to presence of the vane anemometer

NASA Astrophysics Data System (ADS)

Bujalski, M.; Gawor, M.; Sobczyk, J.

2014-08-01

This paper presents the results of the preliminary experimental investigations of the disturbance of velocity field resulting from placing a vane anemometer in the analyzed air flow. Experiments were conducted in a wind tunnel with a closed loop. For the measurement process, Particle Image Velocimetry (PIV) method was used to visualize the flow structure and evaluate the instantaneous, two-dimensional velocity vector fields. Regions of inflow on the vane anemometer as well as flow behind it were examined. Ensemble averaged velocity distribution and root-mean-square (RMS) velocity fluctuations were determined. The results below are presented in the form of contour-velocity maps and profile plots. In order to investigate velocity fluctuations in the wake of vane anemometer with high temporal resolution hot-wire anemometry (HWA) technique was used. Frequency analysis by means of Fast Fourier Transform was carried out. The obtained results give evidence to a significant spatially and temporally complex flow disturbance in the vicinity of analyzed instrument.

An MDOE Assessment of Nozzle Vanes for High Bypass Ratio Jet Noise Reduction

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Norum, Thomas; Bridges, James

2006-01-01

The effect of airfoil-shaped vanes placed in the fan stream of a BPR 8 coannular nozzle model system on the resulting jet noise was investigated. The experiments used a Modern Design of Experiments approach to investigate the impact of a range of vane parameters on the noise reduction achieved at representative takeoff conditions. The experimental results showed that the installation of the vanes decreased low frequency noise radiation in the downstream peak-noise direction and increased high frequency noise in the upstream direction. Results also showed that improper selection of the vane configuration resulted in increased low frequency noise radiation in the upstream direction. Large angles of attack are shown to reduce noise near the peak jet noise angle and increase noise in the upstream direction. The MDOE analysis yields an optimum design that minimizes perceived noise levels. Limited data taken with a BPR 5 nozzle system showed that the vanes result in better effective perceived noise reduction for lower bypass ratio nozzles than for the BPR 8 model.

Effect of vane twist on the performance of dome swirlers for gas turbine airblast atomizers

NASA Technical Reports Server (NTRS)

Micklow, Gerald J.; Dogra, Anju S.; Nguyen, H. Lee

1990-01-01

For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.

Effect of vane twist on the performance of dome swirlers for gas turbine airblast atomizers

NASA Astrophysics Data System (ADS)

Micklow, Gerald J.; Dogra, Anju S.; Nguyen, H. Lee

1990-07-01

For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.

Effect of vane twist on the performance of dome swirlers for gas turbine airblast atomizers

NASA Astrophysics Data System (ADS)

Micklow, Gerald J.; Dogra, Anju S.; Nguyen, H. Lee

1990-06-01

For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.

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

NASA Technical Reports Server (NTRS)

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

2006-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2011-11-01

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

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.

A Comparison of Hyporheic Transport at a Cross-Vane Structure and Natural Riffle.

PubMed

Smidt, Samuel J; Cullin, Joseph A; Ward, Adam S; Robinson, Jesse; Zimmer, Margaret A; Lautz, Laura K; Endreny, Theodore A

2015-01-01

While restoring hyporheic flowpaths has been cited as a benefit to stream restoration structures, little documentation exists confirming that constructed restoration structures induce comparable hyporheic exchange to natural stream features. This study compares a stream restoration structure (cross-vane) to a natural feature (riffle) concurrently in the same stream reach using time-lapsed electrical resistivity (ER) tomography. Using this hydrogeophysical approach, we were able to quantify hyporheic extent and transport beneath the cross-vane structure and the riffle. We interpret from the geophysical data that the cross-vane and the natural riffle induced spatially and temporally unique hyporheic extent and transport, and the cross-vane created both spatially larger and temporally longer hyporheic flowpaths than the natural riffle. Tracer from the 4.67-h injection was detected along flowpaths for 4.6 h at the cross-vane and 4.2 h at the riffle. The spatial extent of the hyporheic zone at the cross-vane was 12% larger than that at the riffle. We compare ER results of this study to vertical fluxes calculated from temperature profiles and conclude significant differences in the interpretation of hyporheic transport from these different field techniques. Results of this study demonstrate a high degree of heterogeneity in transport metrics at both the cross-vane and the riffle and differences between the hyporheic flowpath networks at the two different features. Our results suggest that restoration structures may be capable of creating sufficient exchange flux and timescales of transport to achieve the same ecological functions as natural features, but engineering of the physical and biogeochemical environment may be necessary to realize these benefits. © 2014, National Ground Water Association.

Vane Separation Control in a Linear Cascade with Area Expansion using AC DBD Plasma Actuators

NASA Astrophysics Data System (ADS)

Kleven, Christopher; Corke, Thomas

2013-11-01

Experiments are presented on the use of AC dielectric barrier discharge (DBD) plasma actuators to prevent flow separation on vanes in a linear cascade with area expansion. The inlet Mach number to the cascade ranged from 0.3 to 0.5, and the vane chord Reynolds numbers ranged from 0 . 9 ×106 to 1 . 5 ×106 . Three cascade designs with different amounts of area expansion, providing different degrees of adverse pressure gradients, were examined. Surface flow visualization revealed a 3-D separation bubble with strong recirculation that formed on the suction side of the vanes. The pattern agreed well with CFD simulations. Plasma actuators were placed on the suction sides of the vanes, just upstream of the flow separation location. Quantitative measurements were performed in the wakes of the vanes using a 5-hole Pitot probe. The measurements were used to determine the effect of the plasma actuator separation control on the pressure loss coefficient, and flow turning angle through the cascades. Overall, the plasma actuators separation control increased the velocity magnitude and dynamic pressure in the passage between the vanes, resulted in a more spanwise-uniform flow turning angle in the vane passage, and significantly lowered the loss coefficient compared to the baseline.

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.

Cyclic stress analysis of an air-cooled turbine vane

NASA Technical Reports Server (NTRS)

Kaufman, A.; Gauntner, D. J.; Gauntner, J. W.

1975-01-01

The effects of gas pressure level, coolant temperature, and coolant flow rate on the stress-strain history and life of an air-cooled vane were analyzed using measured and calculated transient metal temperatures and a turbine blade stress analysis program. Predicted failure locations were compared to results from cyclic tests in a static cascade and engine. The results indicate that a high gas pressure was detrimental, a high coolant flow rate somewhat beneficial, and a low coolant temperature the most beneficial to vane life.

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.

Study of turbine and guide vanes integration to enhance the performance of cross flow vertical axis wind turbine

NASA Astrophysics Data System (ADS)

Wibowo, Andreas; Tjahjana, Dominicus Danardono Dwi Prija; Santoso, Budi; Situmorang, Marcelinus Risky Clinton

2018-02-01

The main purpose of this study is to investigate the best configuration between guide vanes and cross flow vertical axis wind turbine with variation of several parameters including guide vanes tilt angle and the number of turbine and guide vane blades. The experimental test were conducted under various wind speed and directions for testing cross flow wind turbine, consisted of 8, 12 and 16 blades. Two types of guide vane were developed in this study, employing 20° and 60° tilt angle. Both of the two types of guide vane had three variations of blade numbers which had same blade numbers variations as the turbines. The result showed that the configurations between 60° guide vane with 16 blade numbers and turbine with 16 blade numbers had the best configurations. The result also showed that for certain configuration, guide vane was able to increase the power generated by the turbine significantly by 271.39% compared to the baseline configuration without using of guide vane.

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

ERIC Educational Resources Information Center

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

2008-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

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

NASA Astrophysics Data System (ADS)

Acosta, Gregorio I.

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

Detailed flow surveys of turning vanes designed for a 0.1-scale model of NASA Lewis Research Center's proposed altitude wind tunnel

NASA Technical Reports Server (NTRS)

Moore, Royce D.; Shyne, Rickey J.; Boldman, Donald R.; Gelder, Thomas F.

1987-01-01

Detailed flow surveys downstream of the corner turning vanes and downstream of the fan inlet guide vanes have been obtained in a 0.1-scale model of the NASA Lewis Research Center's proposed Altitude Wind Tunnel. Two turning vane designs were evaluated in both corners 1 and 2 (the corners between the test section and the drive fan). Vane A was a controlled-diffusion airfoil and vane B was a circular-arc airfoil. At given flows the turning vane wakes were surveyed to determine the vane pressure losses. For both corners the vane A turning vane configuration gave lower losses than the vane B configuration in the regions where the flow regime should be representative of two-dimensional flow. For both vane sets the vane loss coefficient increased rapidly near the walls.

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

PubMed

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

2012-04-01

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

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

NASA Astrophysics Data System (ADS)

Shen, Lu; Wen, Chih-yung

2017-06-01

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

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

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.

1988-01-01

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

Unsteady behavior and control of vortices in centrifugal compressor

NASA Astrophysics Data System (ADS)

Ohta, Yutaka; Fujisawa, Nobumichi

2014-10-01

Two examples of the use of vortex control to reduce noise and enhance the stable operating range of a centrifugal compressor are presented in this paper. In the case of high-flow operation of a centrifugal compressor with a vaned diffuser, a discrete frequency noise induced by interaction between the impeller-discharge flow and the diffuser vane, which appears most notably in the power spectra of the radiated noise, can be reduced using a tapered diffuser vane (TDV) without affecting the performance of the compressor. Twin longitudinal vortices produced by leakage flow passing through the tapered portion of the diffuser vane induce secondary flow in the direction of the blade surface and prevent flow separation from the leading edge of the diffuser. The use of a TDV can effectively reduce both the discrete frequency noise generated by the interaction between the impeller-discharge flow and the diffuser surface and the broadband turbulent noise component. In the case of low-flow operation, a leading-edge vortex (LEV) that forms on the shroud side of the suction surface near the leading edge of the diffuser increases significantly in size and blocks flow in the diffuser passage. The formation of an LEV may adversely affect the performance of the compressor and may cause the diffuser to stall. Using a one-side tapered diffuser vane to suppress the evolution of an LEV, the stable operating range of the compressor can be increased by more than 12 percent, and the pressure-rise characteristics of the compressor can be improved. The results of a supplementary examination of the structure and unsteady behavior of LEVs, conducted by means of detailed numerical simulations, are also presented.

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

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2017-11-01

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

Aerodynamic Performance of Scale-Model Turbofan Outlet Guide Vanes Designed for Low Noise

NASA Technical Reports Server (NTRS)

Hughes, Christopher E.

2001-01-01

The design of effective new technologies to reduce aircraft propulsion noise is dependent on an understanding of the noise sources and noise generation mechanisms in the modern turbofan engine. In order to more fully understand the physics of noise in a turbofan engine, a comprehensive aeroacoustic wind tunnel test programs was conducted called the 'Source Diagnostic Test.' The text was cooperative effort between NASA and General Electric Aircraft Engines, as part of the NASA Advanced Subsonic Technology Noise Reduction Program. A 1/5-scale model simulator representing the bypass stage of a current technology high bypass ratio turbofan engine was used in the test. The test article consisted of the bypass fan and outlet guide vanes in a flight-type nacelle. The fan used was a medium pressure ratio design with 22 individual, wide chord blades. Three outlet guide vane design configurations were investigated, representing a 54-vane radial Baseline configuration, a 26-vane radial, wide chord Low Count configuration and a 26-vane, wide chord Low Noise configuration with 30 deg of aft sweep. The test was conducted in the NASA Glenn Research Center 9 by 15-Foot Low Speed Wind Tunnel at velocities simulating the takeoff and approach phases of the aircraft flight envelope. The Source Diagnostic Test had several acoustic and aerodynamic technical objectives: (1) establish the performance of a scale model fan selected to represent the current technology turbofan product; (2) assess the performance of the fan stage with each of the three distinct outlet guide vane designs; (3) determine the effect of the outlet guide vane configuration on the fan baseline performance; and (4) conduct detailed flowfield diagnostic surveys, both acoustic and aerodynamic, to characterize and understand the noise generation mechanisms in a turbofan engine. This paper addresses the fan and stage aerodynamic performance results from the Source Diagnostic Test.

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

NASA Technical Reports Server (NTRS)

Creager, Marcus O.

1959-01-01

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

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

PubMed

Ziemba, Brian P; Falke, Joseph J

2018-01-01

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

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

PubMed Central

Ziemba, Brian P.

2018-01-01

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

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

NASA Astrophysics Data System (ADS)

Linehan, Thomas; Mohseni, Kamran

2017-11-01

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

Textbook Multigrid Efficiency for Leading Edge Stagnation

NASA Technical Reports Server (NTRS)

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

2004-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-04-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-05-01

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

Effect of coolant flow ejection on aerodynamic performance of low-aspect-ratio vanes. 1: Performance with coolant ejection holes plugged

NASA Technical Reports Server (NTRS)

Haas, J. E.; Kofskey, M. G.

1976-01-01

The aerodynamic performance of a low aspect ratio turbine vane designed with coolant flow ejection holes on the vane surfaces was experimentally determined in a full-annular cascade with the coolant ejection holes plugged. The purpose was to establish a baseline for comparison with tests where flow is ejected from the vane surfaces. The vanes were tested over a mean-section ideal critical velocity ratio range of 0.64 to 0.98. This ideal critical velocity ratio corresponds to the vane inlet total to vane aftermixed static pressure ratio at the mean section. The variations in vane efficiency and aftermixed flow conditions with circumferential and radial position were obtained.

Performance of a low-pressure-ratio centrifugal compressor with four diffuser designs

NASA Technical Reports Server (NTRS)

Klassen, H. A.

1973-01-01

A low-pressure-ratio centrifugal compressor was tested with four different diffuser configurations. One diffuser had airfoil vanes. Two were pipe diffusers. One pipe diffuser had 7.5 deg cone diffusing passages. The other had trumpet-shaped passages designed for linear static-pressure rise from throat to exit. The fourth configuration had flat vanes with elliptical leading edges similar to those of pipe diffusers. The side walls were contoured to produce a linear pressure rise. Peak compressor efficiencies were 0.82 with the airfoil vane and conical pipe diffusers, 0.80 with the trumpet, and 0.74 with the flat-vane design. Surge margin and useful range were greater for the airfoil-vane diffuser than for the other three.

Similarity tests of turbine vanes, effects of ceramic thermal barrier coatings

NASA Technical Reports Server (NTRS)

Gladden, H. J.

1980-01-01

The role of material thermal conductivity was analyzed for its effect on the thermal performance of air-cooled gas turbine components coated with a ceramic thermal barrier material when tested at reduced temperatures and pressures. It is shown that the thermal performance can be evaluated reliably at reduced gas and coolant conditions; however, thermal conductivity corrections are required for the data at reduced conditions. Corrections for a ceramic thermal barrier coated vane are significantly different than for an uncoated vane. Comparison of uncorrected test data, therefore, would show erroneously that the thermal barrier coating was ineffective. When thermal conductivity corrections are applied to the test data these data are then shown to be representative of engine data and also show that the thermal barrier coating increases the vane cooling effectiveness by 12.5 percent.

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.

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

NASA Technical Reports Server (NTRS)

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

2016-01-01

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

Turbofan gas turbine engine with variable fan outlet guide vanes

NASA Technical Reports Server (NTRS)

Wood, Peter John (Inventor); LaChapelle, Donald George (Inventor); Grant, Carl (Inventor); Zenon, Ruby Lasandra (Inventor); Mielke, Mark Joseph (Inventor)

2010-01-01

A turbofan gas turbine engine includes a forward fan section with a row of fan rotor blades, a core engine, and a fan bypass duct downstream of the forward fan section and radially outwardly of the core engine. The forward fan section has only a single stage of variable fan guide vanes which are variable fan outlet guide vanes downstream of the forward fan rotor blades. An exemplary embodiment of the engine includes an afterburner downstream of the fan bypass duct between the core engine and an exhaust nozzle. The variable fan outlet guide vanes are operable to pivot from a nominal OGV position at take-off to an open OGV position at a high flight Mach Number which may be in a range of between about 2.5-4+. Struts extend radially across a radially inwardly curved portion of a flowpath of the engine between the forward fan section and the core engine.

Heat Transfer Measurements for a Film Cooled Turbine Vane Cascade

NASA Technical Reports Server (NTRS)

Poinsatte, Philip E.; Heidmann, James D.; Thurman, Douglas R.

2008-01-01

Experimental heat transfer and pressure measurements were obtained on a large scale film cooled turbine vane cascade. The objective was to investigate heat transfer on a commercial high pressure first stage turbine vane at near engine Mach and Reynolds number conditions. Additionally blowing ratios and coolant density were also matched. Numerical computations were made with the Glenn-HT code of the same geometry and compared with the experimental results. A transient thermochromic liquid crystal technique was used to obtain steady state heat transfer data on the mid-span geometry of an instrumented vane with 12 rows of circular and shaped film cooling holes. A mixture of SF6 and Argon gases was used for film coolant to match the coolant-to-gas density ratio of a real engine. The exit Mach number and Reynolds number were 0.725 and 2.7 million respectively. Trends from the experimental heat transfer data matched well with the computational prediction, particularly for the film cooled case.

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…

Assessment of Soft Vane and Metal Foam Engine Noise Reduction Concepts

NASA Technical Reports Server (NTRS)

Jones, Michael G.; Parrott, Tony L.; Sutliff, Daniel L.; Hughes, Chris

2009-01-01

Two innovative fan-noise reduction concepts developed by NASA are presented - soft vanes and over-the-rotor metal foam liners. Design methodologies are described for each concept. Soft vanes are outlet guide vanes with internal, resonant chambers that communicate with the exterior aeroacoustic environment via a porous surface. They provide acoustic absorption via viscous losses generated by interaction of unsteady flows with the internal solid structure. Over-the-rotor metal foam liners installed at or near the fan rotor axial plane provide rotor noise absorption. Both concepts also provide pressure-release surfaces that potentially inhibit noise generation. Several configurations for both concepts are evaluated with a normal incidence tube, and the results are used to guide designs for implementation in two NASA fan rigs. For soft vanes, approximately 1 to 2 dB of broadband inlet and aft-radiated fan noise reduction is achieved. For over-the-rotor metal foam liners, up to 3 dB of fan noise reduction is measured in the low-speed fan rig, but minimal reduction is measured in the high-speed fan rig. These metal foam liner results are compared with a static engine test, in which inlet sound power level reductions up to 5 dB were measured. Brief plans for further development are also provided.

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

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2008-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2013-11-01

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

Static thrust-vectoring performance of nonaxisymmetric convergent-divergent nozzles with post-exit yaw vanes. M.S. Thesis - George Washington Univ., Aug. 1988

NASA Technical Reports Server (NTRS)

Foley, Robert J.; Pendergraft, Odis C., Jr.

1991-01-01

A static (wind-off) test was conducted in the Static Test Facility of the 16-ft transonic tunnel to determine the performance and turning effectiveness of post-exit yaw vanes installed on two-dimensional convergent-divergent nozzles. One nozzle design that was previously tested was used as a baseline, simulating dry power and afterburning power nozzles at both 0 and 20 degree pitch vectoring conditions. Vanes were installed on these four nozzle configurations to study the effects of vane deflection angle, longitudinal and lateral location, size, and camber. All vanes were hinged at the nozzle sidewall exit, and in addition, some were also hinged at the vane quarter chord (double-hinged). The vane concepts tested generally produced yaw thrust vectoring angles much less than the geometric vane angles, for (up to 8 percent) resultant thrust losses. When the nozzles were pitch vectored, yawing effectiveness decreased as the vanes were moved downstream. Thrust penalties and yawing effectiveness both decreased rapidly as the vanes were moved outboard (laterally). Vane length and height changes increased yawing effectiveness and thrust ratio losses, while using vane camber, and double-hinged vanes increased resultant yaw angles by 50 to 100 percent.

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

NASA Technical Reports Server (NTRS)

Maki, Ralph L.

1959-01-01

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

Cfd Simulation of Capillary Rise of Liquid in Cylindrical Container with Lateral Vanes

NASA Astrophysics Data System (ADS)

Liu, Xiaolin; Huang, Yiyong; Li, Guangyu

2016-06-01

Orbit refueling is one of the most significant technologies, which has vital strategic meaning. It can enhance the flexibility and prolong the lifetime of the spacecrafts. Space propellant management is one of the key technologies in orbit refueling. Based on the background of space propellant management, CFD simulations of capillary rise of liquid in Cylindrical container with lateral vanes in space condition were carried out in this paper. The influence of the size and the number of the vanes to the capillary flow were analyzed too. The results can be useful to the design and optimization of the propellant management device in the vane type surface tension tank.

A study of high-lift airfoils at high Reynolds numbers in the Langley low-turbulence pressure tunnel

NASA Technical Reports Server (NTRS)

Morgan, Harry L., Jr.; Ferris, James C.; Mcghee, Robert J.

1987-01-01

An experimental study was conducted in the Langley Low Turbulence Pressure Tunnel to determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic performance of two supercritical type airfoils, one equipped with a conventional flap system and the other with an advanced high lift flap system. The conventional flap system consisted of a leading edge slat and a double slotted, trailing edge flap with a small chord vane and a large chord aft flap. The advanced flap system consisted of a leading edge slat and a double slotted, trailing edge flap with a large chord vane and a small chord aft flap. Both models were tested with all elements nested to form the cruise airfoil and with the leading edge slat and with a single or double slotted, trailing edge flap deflected to form the high lift airfoils. The experimental tests were conducted through a Reynolds number range from 2.8 to 20.9 x 1,000,000 and a Mach number range from 0.10 to 0.35. Lift and pitching moment data were obtained. Summaries of the test results obtained are presented and comparisons are made between the observed aerodynamic performance trends for both models. The results showing the effect of leading edge frost and glaze ice formation is given.

Analysis on influence of guide vanes closure laws of pump-turbine on load rejection transient process

NASA Astrophysics Data System (ADS)

Yao, Z.; Bi, H. L.; Huang, Q. S.; Li, Z. J.; Wang, Z. W.

2013-12-01

In load rejection transient process, the sudden shut down of guide vanes may cause units speed rise and a sharp increase in water hammer pressure of diversion system, which endangers the safety operation of the power plant. Adopting reasonable guide vane closure law is a kind of economic and effective measurement to reduce the water hammer pressure and limit rotational speed increases. In this paper, combined with Guangzhou Pumped Storage Power Station plant A, the load rejection condition under different guide vanes closure laws is calculated and the key factor of guide vanes closure laws on the impact of the load rejection transition process is analyzed. The different inflection points, which are the closure modes, on the impact of unit speed change, water level fluctuation of surge tank, and the pressure fluctuation of volute inlet and draft tube inlet are further discussed. By compared with the calculation results, a reasonable guide vanes inflection point position can be determined according to security requirements and a reasonable guide vanes closure law can be attained to effectively coordinate the unit speed rise and the rapid pressure change in the load rejection transient process.

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

NASA Astrophysics Data System (ADS)

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

2017-07-01

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

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

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1999-01-01

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

Analysis of heat-transfer tests of an impingement-convection- and film-cooled vane in a cascade

NASA Technical Reports Server (NTRS)

Gladden, H. J.; Gauntner, D. J.; Livingood, J. N. B.

1971-01-01

Experimental flow and heat transfer data obtained for an air-cooled turbine vane tested in a static cascade at gas temperatures and pressures to 1644 K (2500 F) and 31 N/cm2 (45 psia), respectively, are presented. Average and local vane temperatures were correlated in several ways. Calculated and measured coolant flows and vane temperatures are compared. Potential allowable increases in gas temperature are also discussed.

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.

The image evaluation of iterative motion correction reconstruction algorithm PROPELLER T2-weighted imaging compared with MultiVane T2-weighted imaging

NASA Astrophysics Data System (ADS)

Lee, Suk-Jun; Yu, Seung-Man

2017-08-01

The purpose of this study was to evaluate the usefulness and clinical applications of MultiVaneXD which was applying iterative motion correction reconstruction algorithm T2-weighted images compared with MultiVane images taken with a 3T MRI. A total of 20 patients with suspected pathologies of the liver and pancreatic-biliary system based on clinical and laboratory findings underwent upper abdominal MRI, acquired using the MultiVane and MultiVaneXD techniques. Two reviewers analyzed the MultiVane and MultiVaneXD T2-weighted images qualitatively and quantitatively. Each reviewer evaluated vessel conspicuity by observing motion artifacts and the sharpness of the portal vein, hepatic vein, and upper organs. The signal-to-noise ratio (SNR) and contrast-to-noise ratio (CNR) were calculated by one reviewer for quantitative analysis. The interclass correlation coefficient was evaluated to measure inter-observer reliability. There were significant differences between MultiVane and MultiVaneXD in motion artifact evaluation. Furthermore, MultiVane was given a better score than MultiVaneXD in abdominal organ sharpness and vessel conspicuity, but the difference was insignificant. The reliability coefficient values were over 0.8 in every evaluation. MultiVaneXD (2.12) showed a higher value than did MultiVane (1.98), but the difference was insignificant ( p = 0.135). MultiVaneXD is a motion correction method that is more advanced than MultiVane, and it produced an increased SNR, resulting in a greater ability to detect focal abdominal lesions.

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

NASA Astrophysics Data System (ADS)

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

2018-05-01

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

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

PubMed Central

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

2015-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2010-01-01

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

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

NASA Technical Reports Server (NTRS)

Oh, S.; Tavella, D.

1986-01-01

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

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

NASA Technical Reports Server (NTRS)

Henderson, W. P.

1976-01-01

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

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…

Numerical investigation of rotating stall in centrifugal compressor with vaned and vaneless diffuser

NASA Astrophysics Data System (ADS)

Halawa, Taher; Alqaradawi, Mohamed; Gadala, Mohamed S.; Shahin, Ibrahim; Badr, Osama

2015-06-01

This study presents a numerical simulation of the stall and surge in a centrifugal compressor and presents a descriptionof the stall development in two different cases. The first case is for a compressor with vaneless diffuser and the second is for a compressor with vaned diffuser of the vane island shape. The main aim of this study is to compare the flow characteristics and behavior for the two compressors near the surge operating condition and provide further understanding of the diffuser role when back flow occurs at surge. Results showed that for a locationnear the diffuser entrance, the amplitude of the static pressure fluctuations for the vaneless diffuser case is higher than that for the vaned diffuser case near surge condition. These pressure fluctuations in the case of the vaneless diffuser appear with a gradual decrease of the mean pressure value as a part of the surge cycle. While for the case of the vaned diffuser, the pressure drop during surge occurs faster than the case of the vaneless diffuser. Also, results indicated that during surge in the case of vaneless diffuser, there is a region with low velocity and back flow that appears as a layer connecting all impeller passages near shroud surface and this layer develops in size with time. On the other hand, for the case of vaned diffuser during surge, the low velocity regions appear in random locations in some passages and these regions expand with time towards the shroud surface. Results showed that during stall, the impeller passages are exposed to identical impact from stall cells in the case of vaneless diffuser while the stall effect varies from passage to another in the case of the vaned diffuser.

Performance of an alpha-vane and pitot tube in simulated heavy rain environment

NASA Technical Reports Server (NTRS)

Luers, J. K.; Fiscus, I. B.

1985-01-01

Experimental tests were conducted in the UDRI Environmental Wind/Rain Tunnel to establish the performance of an alpha-vane, that measures angle of attack, in a simulated heavy rain environment. The tests consisted of emersing the alpha-vane in an airstream with a concurrent water spray penetrating vertically through the airstream. The direction of the spray was varied to make an angle of 5.8 to 18 deg with the airstream direction in order to simulate the conditions that occur when an aircraft lands in a heavy rain environment. Rainrates simulated varied from 1000 to 1200 mm/hr which are the most severe ever expected to be encountered by an aircraft over even a 30 second period. Tunnel airspeeds ranged from 85 to 125 miles per hour. The results showed that even the most severe rainrates produced a misalignment in the alpha-vane of only 1 deg away from the airstream direction. Thus for normal rain conditions experienced by landing aircraft no significant deterioration in alpha-vane performance is expected.

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.

Gas turbine row #1 steam cooled vane

DOEpatents

Cunha, Frank J.

2000-01-01

A design for a vane segment having a closed-loop steam cooling system is provided. The vane segment comprises an outer shroud, an inner shroud and an airfoil, each component having a target surface on the inside surface of its walls. A plurality of rectangular waffle structures are provided on the target surface to enhance heat transfer between each component and cooling steam. Channel systems are provided in the shrouds to improve the flow of steam through the shrouds. Insert legs located in cavities in the airfoil are also provided. Each insert leg comprises outer channels located on a perimeter of the leg, each outer channel having an outer wall and impingement holes on the outer wall for producing impingement jets of cooling steam to contact the airfoil's target surface. Each insert leg further comprises a plurality of substantially rectangular-shaped ribs located on the outer wall and a plurality of openings located between outer channels of the leg to minimize cross flow degradation.

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

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2015-11-01

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

Flow control in axial fan inlet guide vanes by synthetic jets

NASA Astrophysics Data System (ADS)

Cyrus, V.; Trávníček, Z.; Wurst, P.; Kordík, J.

2013-04-01

Tested high pressure axial flow fan with hub/tip ratio of 0.70 and external diameter of 600 mm consisted of inlet guide vanes (IGV), rotor and stator blade rows. Fan peripheral velocity was 47 m/s. Air volume flow rate was changed by turning of rear part of the inlet guide vanes. At turning of 20 deg the flow was separated on the IGV profiles. The synthetic jets were introduced through radial holes in machine casing in the location before flow separation origin. Synthetic jet actuator was designed with the use of a speaker by UT AVCR. Its membrane had diameter of 63 mm. Excitation frequency was chosen in the range of 500 Hz - 700 Hz. Synthetic jets favourably influenced separated flow on the vane profiles in the distance of (5 - 12) mm from the casing surface. The reduction of flow separation area caused in the region near the casing the decrease of the profile loss coefficient approximately by 20%.

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

NASA Technical Reports Server (NTRS)

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

1979-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2014-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2015-01-01

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

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

PubMed Central

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

2015-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

Estimating Blade Section Airloads from Blade Leading-Edge Pressure Measurements

NASA Technical Reports Server (NTRS)

vanAken, Johannes M.

2003-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

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

NASA Astrophysics Data System (ADS)

Mishra, Arpit; Ghosh, Parthasarathi

2017-02-01

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

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

PubMed Central

Klarlund, Jes K.

2012-01-01

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

Barb geometry of asymmetrical feathers reveals a transitional morphology in the evolution of avian flight

PubMed Central

Feo, Teresa J.; Field, Daniel J.; Prum, Richard O.

2015-01-01

The geometry of feather barbs (barb length and barb angle) determines feather vane asymmetry and vane rigidity, which are both critical to a feather's aerodynamic performance. Here, we describe the relationship between barb geometry and aerodynamic function across the evolutionary history of asymmetrical flight feathers, from Mesozoic taxa outside of modern avian diversity (Microraptor, Archaeopteryx, Sapeornis, Confuciusornis and the enantiornithine Eopengornis) to an extensive sample of modern birds. Contrary to previous assumptions, we find that barb angle is not related to vane-width asymmetry; instead barb angle varies with vane function, whereas barb length variation determines vane asymmetry. We demonstrate that barb geometry significantly differs among functionally distinct portions of flight feather vanes, and that cutting-edge leading vanes occupy a distinct region of morphospace characterized by small barb angles. This cutting-edge vane morphology is ubiquitous across a phylogenetically and functionally diverse sample of modern birds and Mesozoic stem birds, revealing a fundamental aerodynamic adaptation that has persisted from the Late Jurassic. However, in Mesozoic taxa stemward of Ornithurae and Enantiornithes, trailing vane barb geometry is distinctly different from that of modern birds. In both modern birds and enantiornithines, trailing vanes have larger barb angles than in comparatively stemward taxa like Archaeopteryx, which exhibit small trailing vane barb angles. This discovery reveals a previously unrecognized evolutionary transition in flight feather morphology, which has important implications for the flight capacity of early feathered theropods such as Archaeopteryx and Microraptor. Our findings suggest that the fully modern avian flight feather, and possibly a modern capacity for powered flight, evolved crownward of Confuciusornis, long after the origin of asymmetrical flight feathers, and much later than previously recognized. PMID

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

NASA Technical Reports Server (NTRS)

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

2004-01-01

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

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

NASA Technical Reports Server (NTRS)

1958-01-01

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

Experimental evaluation of two turning vane designs for high-speed corner of 0.1-scale model of NASA Lewis Research Center's proposed altitude wind tunnel

NASA Technical Reports Server (NTRS)

Moore, R. D.; Boldman, D. R.; Shyne, R. J.

1986-01-01

Two turning vane designs were experimentally evaluated for corner 1 (downstream of the test section) of a 0.1-scale model of the NASA Lewis Research Center's proposed Altitude Wind Tunnel (AWT). Vane A was a controlled-diffusion airfoil shape; vane B was a circular-arc airfoil shape. The vane designs were tested over corner inlet Mach numbers from 0.16 to 0.465. Several modifications in vane setting angle and vane spacing were also evaluated for vane A. The overall performance obtained from total pressure rakes indicated that vane B had a slightly lower loss coefficient than vane A. At Mach 0.35 (the design Mach number without the engine exhaust removal scoop), the loss coefficients were 0.150 and 0.178 for vanes B and A, respectively. Resetting the vane A angle by -5 deg. (vane A10) to turn the flow toward the outside corner reduced the loss coefficient to 0.119. The best configuration (vane A10) was also tested with a simulated engine exhaust removal scoop. The loss coefficient for that configuration was 0.164 at Mach 0.41 (the approximate design Mach number with the scoop).

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.

Gas turbine nozzle vane insert and methods of installation

DOEpatents

Miller, William John; Predmore, Daniel Ross; Placko, James Michael

2002-01-01

A pair of hollow elongated insert bodies are disposed in one or more of the nozzle vane cavities of a nozzle stage of a gas turbine. Each insert body has an outer wall portion with apertures for impingement-cooling of nozzle wall portions in registration with the outer wall portion. The insert bodies are installed into the cavity separately and spreaders flex the bodies toward and to engage standoffs against wall portions of the nozzle whereby the designed impingement gap between the outer wall portions of the insert bodies and the nozzle wall portions is achieved. The spreaders are secured to the inner wall portions of the insert bodies and the bodies are secured to one another and to the nozzle vane by welding or brazing.

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

DTIC Science & Technology

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

A theoretical study of fluid forces on a centrifugal impeller rotating and whirling in a vaned diffuser

NASA Technical Reports Server (NTRS)

Tsujimoto, Yoshinobu; Acosta, Allan J.; Yoshida, Yoshiki

1989-01-01

The fluid forces on a centrifugal impeller rotating and whirling in a vaned diffuser are analyzed on the assumption that the number of impeller and diffuser vanes is so large that the flows are perfectly guided by the vanes. The flow is taken to be two dimensional, inviscid, and incompressible, but the effects of impeller and diffuser losses are taken into account. It is shown that the interaction with the vaned diffuser may cause destabilizing fluid forces. From these discussions, it is found that the whirling forces are closely related to the steady head-capacity characteristics of the impeller. This physical understanding of the whirling forces can be applied also to the cases with volute casings. At partial capacities, it is shown that the impeller forces change greatly when the flow rate and whirl velocity are near to the impeller or vaned diffuser attributed rotating stall onset capacity, and the stall propagation velocity, respectively. In such cases the impeller forces may become destabilizing for impeller whirl.

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

NASA Technical Reports Server (NTRS)

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

2015-01-01

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

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

NASA Technical Reports Server (NTRS)

1975-01-01

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

An Electronic Weather Vane for Field Science

ERIC Educational Resources Information Center

Burman, J.; Talbert, R.; Carlton, K.

2014-01-01

This paper details the construction of a weather vane for the measurement of wind direction in field situations. The purpose of its construction was to analyse how wind direction affected the attractiveness of an insect pheromone in a dynamic outdoor environment, where wind could be a significant contributor to odour movement. The apparatus…

Turbine vane segment and impingement insert configuration for fail-safe impingement insert retention

DOEpatents

Burdgick, Steven Sebastian; Kellock, Iain Robertson

2003-05-13

An impingement insert sleeve is provided that is adapted to be disposed in a coolant cavity defined through a stator vane. The insert has a generally open inlet end and first and second pairs of diametrically opposed side walls, and at least one fail-safe tab defined at a longitudinal end of the insert for limiting radial displacement of the insert with respect to the stator vane.

Effect of a part-span variable inlet guide vane on the performance of a high-bypass turbofan engine

NASA Technical Reports Server (NTRS)

Bobula, G. A.; Soeder, R. H.; Burkardt, L. A.

1981-01-01

The ability of a part span variable inlet guide vane (VIGV) to modulate the thrust of a high bypass turbofan engine was evaluated at altitude/Mach number conditions of 4572 m/0.6 and 9144 m/0.93. Fan tip, gas generator and supercharger performance were also determined, both on operating lines and during fan duct throttling. The evaluation was repeated with the bypass splitter extended forward to near the fan blade trailing edge. Gross thrust attentuation of over 50 percent was achieved with 50 degree VIGV closure at 100 percent corrected fan speed. Gas generator supercharger performance fell off with VIGV closure, but this loss was reduced when a splitter extension was added. The effect of VIVG closure on gas generator performance was minimal.

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

NASA Technical Reports Server (NTRS)

Goin, Kennith L

1951-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

Energy efficient engine high-pressure turbine single crystal vane and blade fabrication technology report

NASA Technical Reports Server (NTRS)

Giamei, A. F.; Salkeld, R. W.; Hayes, C. W.

1981-01-01

The objective of the High-Pressure Turbine Fabrication Program was to demonstrate the application and feasibility of Pratt & Whitney Aircraft-developed two-piece, single crystal casting and bonding technology on the turbine blade and vane configurations required for the high-pressure turbine in the Energy Efficient Engine. During the first phase of the program, casting feasibility was demonstrated. Several blade and vane halves were made for the bonding trials, plus solid blades and vanes were successfully cast for materials evaluation tests. Specimens exhibited the required microstructure and chemical composition. Bonding feasibility was demonstrated in the second phase of the effort. Bonding yields of 75 percent for the vane and 30 percent for the blade were achieved, and methods for improving these yield percentages were identified. A bond process was established for PWA 1480 single crystal material which incorporated a transient liquid phase interlayer. Bond properties were substantiated and sensitivities determined. Tooling die materials were identified, and an advanced differential thermal expansion tooling concept was incorporated into the bond process.

Erosion estimation of guide vane end clearance in hydraulic turbines with sediment water flow

NASA Astrophysics Data System (ADS)

Han, Wei; Kang, Jingbo; Wang, Jie; Peng, Guoyi; Li, Lianyuan; Su, Min

2018-04-01

The end surface of guide vane or head cover is one of the most serious parts of sediment erosion for high-head hydraulic turbines. In order to investigate the relationship between erosion depth of wall surface and the characteristic parameter of erosion, an estimative method including a simplified flow model and a modificatory erosion calculative function is proposed in this paper. The flow between the end surfaces of guide vane and head cover is simplified as a clearance flow around a circular cylinder with a backward facing step. Erosion characteristic parameter of csws3 is calculated with the mixture model for multiphase flow and the renormalization group (RNG) k-𝜀 turbulence model under the actual working conditions, based on which, erosion depths of guide vane and head cover end surfaces are estimated with a modification of erosion coefficient K. The estimation results agree well with the actual situation. It is shown that the estimative method is reasonable for erosion prediction of guide vane and can provide a significant reference to determine the optimal maintenance cycle for hydraulic turbine in the future.

Ceramic turbine stator vane and shroud support

DOEpatents

Glenn, Robert G.

1981-01-01

A support system for supporting the stationary ceramic vanes and ceramic outer shrouds which define the motive fluid gas path in a gas turbine engine is shown. Each individual segment of the ceramic component whether a vane or shroud segment has an integral radially outwardly projecting stem portion. The stem is enclosed in a split collet member of a high-temperature alloy material having a cavity configured to interlock with the stem portion. The generally cylindrical external surface of the collet engages a mating internal cylindrical surface of an aperture through a supporting arcuate ring segment with mating camming surfaces on the two facing cylindrical surfaces such that radially outward movement of the collet relative to the ring causes the internal cavity of the collet to be reduced in diameter to tightly engage the ceramic stem disposed therein. A portion of the collet extends outwardly through the ring segment opposite the ceramic piece and is threaded for receiving a nut and a compression washer for retaining the collet in the ring segment under a continuous biasing force urging the collet radially outwardly.

Numerical investigation on performance and sediment erosion of Francis runner with different guide vane profiles

NASA Astrophysics Data System (ADS)

Lama, R.; Dahal, D. R.; Gautam, S.; Acharya, N.; Neopane, H.; Thapa, B. S.

2018-06-01

Francis turbine are ideal turbines for Himalayan and Andes region where both low and high-altitude mountains are located. Turbines operating in such regions face operational and maintenance problems due to the sediment erosion. In order to reduce the erosion effects on these components the design of components for higher sediment handling is essence. This paper presents performance analysis of Francis runner and prediction of sediment erosion on the runner blades for different operating conditions with different guide vane profiles. The simulations were carried out for 11 guide vane opening angles using Tabakoff erosion model. At full load and best efficiency point the erosion was localized at pressure side of runner blades outlet due to higher relative velocity. On the other hand, at part load condition, erosion was observed at suction side of the blades. Application of asymmetric guide vane profile NACA 4412 showed higher efficiency for all operating conditions with minimum erosion on runner blades in compare to symmetric guide vane profile NACA 0012.

Background noise measurements from jet exit vanes designed to reduced flow pulsations in an open-jet wind tunnel

NASA Technical Reports Server (NTRS)

Hoad, D. R.; Martin, R. M.

1985-01-01

Many open jet wind tunnels experience pulsations of the flow which are typically characterized by periodic low frequency velocity and pressure variations. One method of reducing these fluctuations is to install vanes around the perimeter of the jet exit to protrude into the flow. Although these vanes were shown to be effective in reducing the fluctuation content, they can also increase the test section background noise level. The results of an experimental acoustic program in the Langley 4- by 7-Meter Tunnel is presented which evaluates the effect on tunnel background noise of such modifications to the jet exit nozzle. Noise levels for the baseline tunnel configuration are compared with those for three jet exit nozzle modifications, including an enhanced noise reduction configuration that minimizes the effect of the vanes on the background noise. Although the noise levels for this modified vane configuration were comparable to baseline tunnel background noise levels in this facility, installation of these modified vanes in an acoustic tunnel may be of concern because the noise levels for the vanes could be well above background noise levels in a quiet facility.

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…

Loss model for off-design performance analysis of radial turbines with pivoting-vane, variable-area stators

NASA Technical Reports Server (NTRS)