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At least 433 records · Page 24

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DOE OSTI2023

Using SWMM for emergency response planning: A case study evaluating biological agent transport under various rainfall scenarios and urban surfaces

To assist in emergency preparedness for a biological agent terrorist attack or accidental pathogen release, potential contaminant levels and migration pathways of spores spread by urban stormwater were evaluated using a Storm Water Management Model (SWMM) of U.S. Coast Guard Base Elizabeth City, North Carolina. The high temporal-spatial resolution SWMM model was built using spore concentrations in stormwater runoff from asphalt, grass, and concrete collected from a point-scale field study. The subsequent modeled contamination scenarios included a notional plume release and point releases mimicking the field study under three rainfall conditions. The rainfall scenarios included a 6-hour natural rainfall event on Dec. 8, 2021 and two design storms (2-year and 100-year events). The observed spore concentrations from asphalt and concrete from the actual field experiment were applied to calibrate the washoff parameters in the SWMM model, using an exponential washoff function. The calibrated washoff coefficient (c 1 ) and exponent (c 2 ) were 0.01 and 1.00 for asphalt, 0.05 and 1.45 for grass, and 2.45 and 1.00 for concrete, respectively. The calibrated SWMM model simulated spore concentrations in runoff at times and magnitudes similar to the field study data. In the point release modeled scenario, the concrete surface generated 55.6% higher average spore concentrations than asphalt. Similarly, in the field experiment, a 175% (p < 0.05) higher average spore concentration in surface runoff was observed from concrete than from asphalt. Here, this study demonstrates how SWMM may be used to evaluate spore washoff from urban surfaces under different precipitation amounts, intensities, and durations, and how visualized spatial migration pathways in stormwater runoff may be used for emergency planning and remediation.

54 ENVIRONMENTAL SCIENCES↗

NASA NTRS1957

Investigation of Heat Transfer from a Stationary and Rotating Conical Forebody

The convective heat transfer from the surface of a conical forebody having a hemispherical nose, an included angle of approximately 30 deg, and. a maximum diameter of 18.9 inches was investigated in a wind tunnel for both stationary and. rotating operation. The range of test conditions included free-stream velocities up to 400 feet per second, rotational speeds up to 1200 rpm, and. angles of attack of 0 deg and 6 deg. Both a uniform surface temperature and a uniform heater input power density were used. The Nusselt-Reynolds number relations provided good correlation of the heat-transfer data for the complete operating range at 0 deg angle of attack with and without spinner rotation, and for 6deg angle of attack with rotation. Rotational speeds up to 1200 rpm had no apparent effect on the heat-transfer characteristics of the spinner. The results obtained at 6 deg angle of attack with rotation were essentially the same as those obtained at 0 deg angle of attack without rotation. The experimental heat-transfer characteristics in the turbulent flow region were consistently in closer agreement with the results predicted for a two-dimensional body than with those predicted. for a cone. For stationary operation at 60 angle of attack, the measured heat-transfer coefficients in the turbulent flow region were from 6 to 13 percent greater on the lower surface (windward. side) than on the upper surface (sheltered side) for corresponding surface locations. The spinner-nose geometry appeared to cause early boundary-layer transition. Transition was initiated at a fairly constant Reynolds number (based on surface distance from nose) of 8.0 x 10(exp 4). Transition was completed at Reynolds numbers less than 5.0 x 10(exp 5) for all conditions investigated.

Ruggeri, Robert S.↗

NASA NTRS1948

Lifting-surface-theory aspect-ratio corrections to the lift and hinge-moment parameters for full-span elevators on horizontal tail surfaces

A limited number of lifting-surface-theory solutions for wings with chordwise loadings resulting from angle of attack, parabolic-ac camber, and flap deflection are now available. These solutions were studied with the purpose of determining methods of extrapolating the results in such a way that they could be used to determine lifting-surface-theory values of the aspect-ratio corrections to the lift and hinge-moment parameters for both angle-of-attack and flap-deflection-type loading that could be used to predict the characteristics of horizontal tail surfaces from section data with sufficient accuracy for engineering purposes. Such a method was devised for horizontal tail surfaces with full-span elevators. In spite of the fact that the theory involved is rather complex, the method is simple to apply and may be applied without any knowledge of lifting-surface theory. A comparison of experimental finite-span and section value and of the estimated values of the lift and hinge-moment parameters for three horizontal tail surfaces was made to provide an experimental verification of the method suggested. (author)

Swanson, Robert S↗

NASA NTRS1987

Aerothermal evaluation of a spherically blunted body with a trapezoidal cross section in the Langley 8-foot high-temperature tunnel

A model to be used in the flow studies and curved Thermal Protection System (TPS) evaluations was tested in the Langley 8 Foot High-Temperature Tunnel at a nominal Mach number of 6.8. The purpose of the study was to define the surface pressure and heating rates at high angles of attack (in support of curved metallic TPS studies) and to determine the conditions for which the model would be suitable as a test bed for aerothermal load studies. The present study was conducted at a nominal total temperature of 2400 and 3300 R, dynamic pressures from 2.3 to 10.9 psia, and free-stream Reynolds numbers from 4000,000 to 1,700,000/ft. The measurements consisted primarily of surface pressure and cold-wall (530 R) heating rates. Qualitative comparisons between predictions and data show that for this configuration, aerothermal tests should be limited to angles of attack between 10 and -10 degrees. Outside this range, the effects of free-stream flow nonuniformity appear in the data, as a result of the long length of the model. However, for TPS testing, this is not a concern and tests can be performed at angles of attack ranging from 20 to -20 degrees. Laminar and naturally turbulent boundary layers are available over limited ranges of conditions.

Albertson, Cindy W.↗

NASA NTRS1984

Inlet flow field investigation. Part 1: Transonic flow field survey

A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.

Yetter, J. A.↗

NASA NTRS1991

Investigation of nose bluntness and angle of attack effects on slender bodies in viscous hypersonic flows

Hypersonic flows over cones and straight biconic configurations are calculated for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. Effect of angle of attack and nose bluntness on these slender cones in air is studied extensively. The numerical procedures are based on the solution of complete Navier-Stokes equations at the nose section and parabolized Navier-Stokes equations further downstream. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The complete flow field is thoroughly analyzed with respect to velocity, temperature, pressure, and entropy profiles. The post shock flow field is studied in detail from the contour plots of Mach number, density, pressure, and temperature. The effect of nose bluntness for slender cones persists as far as 200 nose radii downstream.

Sehgal, A. K.↗

NASA NTRS1992

Combined effect of nose bluntness and angle of attack on slender bodies in viscous hypersonic flows

Tiwari, S. N.↗

NASA NTRS1961

Laminar Heat-Transfer and Pressure Measurements at a Mach Number of 6 on Sharp and Blunt 15 deg Half-Angle Cones at Angles of Attack Up to 90 deg

Two circular conical configurations having 15 deg half-angles were tested in laminar boundary layer at a Mach number of 6 and angles of attack up to 90 deg. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted nose with a bluntness ratio of 0.1428 and a fineness ratio of 1.66. Pressure measurements and schlieren pictures of the flow showed that near-conical flow existed up to an angle of attack of approximately 60 deg. At angles of attack above 70 deg high-pressure areas were present near the base and the bow shock wave was considerably curved. Comparison of the results with simply applied theories showed that on the stagnation line pressures may be predicted by Newtonian theory, and heat transfer by local yawed-cylinder theory based on the yaw angle of the windward generator and the local radius of the cone. Base effects increased the heat transfer in a region extending forward approximately 15 to 30 percent of the windward generator. Circumferential pressure distributions were higher than the corresponding Newtonian distribution and a better prediction was obtained by modifying the theory to match the pressure at 90 deg from the windward generator to that on the surface of the cone at an angle of attack of 0 deg. Circumferential heat-transfer distributions were predicted satisfactorily up to about 60 deg from the stagnation line by using Lees' heat-flux distribution based on the Newtonian pressure. The effects of nose bluntness at large angles of attack were very small in the region beyond two nose radii from the point of tangency.

Conti, Raul J.↗

NASA NTRS1991

An experimental study of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack

An experimental study of the effects of aft blowing on the forebody vortex asymmetry over a 3.0 caliber tangent ogive body at high angles of attack was conducted. The tip of the ogive body was equipped with a single blowing nozzle whose position could be adjusted. The tests were conducted in a subsonic wind tunnel at laminar flow conditions. The effects of model roll, angle of attack, blowing coefficient, and blowing nozzle axial position were independently studied. Surface pressure measurements and flow visualization results were obtained. Aft blowing was observed to alleviate the degree of vortex asymmetry at all angles of attack. The blowing was found to be more effective at the higher angles of attack. However, proportional control of the degree of vortex asymmetry was not observed, because the initial flowfield was highly asymmetric.

Gittner, Nathan M.↗

NASA NTRS1960

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds

An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.

Boltz, Frederick W.↗

NASA NTRS1985

Low-speed experimental study of the vortex flow effects of a fighter forebody having unconventional cross-section

A low speed wind tunnel investigation is conducted for the vortex flow characteristics of a tailless fighter configuration that features full length, fuselage-blended forebody strakes. Static pressure distributions were obtained on the forebody upper surface, and six-component forces and moments were measured on the complete, 60-deg cropped delta wing-fuselage-centerline vertical fin model at angles of attack and sideslip respectively reaching 50 deg and + or - 20 deg. The surface pressure data reveal that development of concentrated vortices from the sharp edged forebody strakes occurred at moderate and high angles of attack. The blended forebody was effective in capturing the windward vortical flow up to high incidence and sideslip angles. The direct suction effect of the windward vortex on the forebody sidewall was the primary contributor to static directional stability at high angles of attack. These improvements are potentially offset, however, by reduced or unstable roll and yaw damping.

Erickson, G. E.↗

NASA NTRS1998

Spatial Characteristics of the Unsteady Differential Pressures on 16 percent F/A-18 Vertical Tails

Buffeting is an aeroelastic phenomenon which plagues high performance aircraft at high angles of attack. For the F/A-18 at high angles of attack, vortices emanating from wing/fuselage leading edge extensions burst, immersing the vertical tails in their turbulent wake. The resulting buffeting of the vertical tails is a concern from fatigue and inspection points of view. Previous flight and wind-tunnel investigations to determine the buffet loads on the tail did not provide a complete description of the spatial characteristics of the unsteady differential pressures. Consequently, the unsteady differential pressures were considered to be fully correlated in the analyses of buffet and buffeting. The use of fully correlated pressures in estimating the generalized aerodynamic forces for the analysis of buffeting yielded responses that exceeded those measured in flight and in the wind tunnel. To learn more about the spatial characteristics of the unsteady differential pressures, an available 16%, sting-mounted, F-18 wind-tunnel model was modified and tested in the Transonic Dynamics Tunnel (TDT) at the NASA Langley Research Center as part of the ACROBAT (Actively Controlled Response Of Buffet-Affected Tails) program. Surface pressures were measured at high angles of attack on flexible and rigid tails. Cross-correlation and cross-spectral analyses of the pressure time histories indicate that the unsteady differential pressures are not fully correlated. In fact, the unsteady differential pressure resemble a wave that travels along the tail. At constant angle of attack, the pressure correlation varies with flight speed.

Moses, Robert W.↗

NASA NTRS1972

Effect of thermally expanding surfaces on aerodynamic roll torques for cones.

An experimental and analytical study has been made of the aerodynamic roll torque on smoothly ablating, spinning cones. The tests, conducted in the Ames 3.5-Foot Hypersonic Wind Tunnel, indicated that the smoothly ablating model rolling at 0 deg angle of attack experiences roll damping, but when rolling at angle of attack an aerodynamic torque in the direction of spin arises if thermal expansion of the body surface occurs. An analysis of the spinning body at angle of attack is included to show how the relative importance of the ablator thermal expansion and the surface recession rate determines whether the motion is damped or undamped.

Mcdevitt, J. B.↗

NASA NTRS1977

Flow-field surveys on the windward side of the NASA 040A space shuttle orbiter at 31 deg angle of attack and Mach 20 in helium

Pitot pressure and flow angle distributions in the windward flow field of the NASA 040A space shuttle orbiter configuration and surface pressures were measured, at a Mach number of 20 and an angle of attack of 31 deg. The free stream Reynolds number, based on model length, was 5.39 x 10 to the 6th power. Results show that cores of high pitot pressure, which are related to the body-shock-wing-shock intersections, occur on the windward plane of symmetry in the vicinity of the wing-body junction and near midspan on the wing. Theoretical estimates of the flow field pitot pressures show that conical flow values for the windward plane of symmetry surface are representative of the average level over the entire lower surface.

Ashby, G. C., Jr.↗

NASA NTRS1953

Preliminary Results of Heat Transfer from a Stationary and Rotating Ellipsoidal Spinner

Convective heat-transfer coefficients in dry air were obtained for an ellipsoidal spinner of 30-inch maximum diameter for both stationary and rotating operation over a range of conditions including airspeeds up to 275 miles per hour, rotational speeds up to 1200 rpm, and angles of attack of zero and 40 The results are presented in terms of Nusselt numbers, Reynolds numbers, and convective heat-transfer coefficients. The studies included both uniform heating densities over the spinner and uniform surface temperatures.. In general, the results showed that rotation will increase the convective heat transfer from a spinner, especially in the turbulent-flow regions. Rotation of the spinner at 1200 rpm and at a free-stream velocity of 275 miles per hour increased the Nusselt number parameter in the turbulent-flow region by 32 percent over that obtained with a stationary spinner; whereas in the nose region, where the flow was laminar, an increase of only 18 percent was observed. Transition from laminar to turbulent flow occurred over a large range of Reynolds numbers primarily because of surface roughness of the spinner. Operation at an angle of attack of 40 had only small effects on the local convective heat transfer for the model studied.

vonGlahn, U.↗

NASA NTRS1980

The linear and non-linear aerodynamics of three-surface aircraft concepts

It is noted that most modern fighter aircraft rely on vortex interaction to provide lift enhancement at maneuvering angles of attack. It is shown that the close-coupled horizontal canard in a three-surface configuration provides a control surface which in addition to its other control functions, can be used to optimize this vortex interaction. Attention is given to a study intended to provide a detailed understanding of the aerodynamics of this type of configuration. The discussion examines the results of this investigation and hypotheses are presented to explain the linear and nonlinear aerodynamic phenomena observed.

Agnew, J. W.↗

NASA NTRS1973

Investigation of configuration effects on entry heating distributions at Mach no. equal 8.0 (OH41)

Aerodynamic heating data were obtained on 0.006 scale models of four Rockwell International SSV double delta wing Orbiters in the Mach 8 variable density tunnel. A model of two previously tested Rockwell International Orbiters which are identified in the Configuration Description of this report were also tested. Orbiter surfaces were thermally mapped from the laminar through turbulent flight regimes during re-entry. Various modifications were made to model lower surfaces to determine the cause of transition in the vicinity of 3.0 million Reynolds number per foot. Re-entry data were acquired for angles of attack from 25 through 35 degrees at nominal Reynolds numbers per foot of 1.0, 2.0, 2.3, 2.5, 3.0, 3.5, 4.5 and 6.0 million utilizing the phase change paint technique. Launch data were acquired on the model upper surfaces for angles of attack of 0 and -5 degrees at nominal Reynolds numbers per foot of 3.0 and 6.0 million. A total of 70 orbiter heating runs and 6 material sample sphere runs were completed.

Gorowitz, H.↗

NASA NTRS1987

Distortion of the vortex core during blade-vortex interaction

The effect of vortex distortion on the flow arising from the interaction of a convecting vortex and an airfoil is analyzed for two-dimensional, incompressible, inviscid flow. Significant distortion of the vortex structure, owing to interaction with the airfoil, can occur, especially during close interactions. The vorticity field is represented by multiple vortices. Vortex instability and diffusion, resulting from discretization in space and time, are discussed. An integral boundary element method is used to represent the airfoil surface. The body panels are re-distributed on the surface as the vortices are convected past to calculate the effect of the vorticity field more accurately. The method is applied to several cases to examine effects of vortex strength, distribution, initial separation position, and airfoil angle of attack. Calculations of vortex trajectory and distortion, and airfoil surface-pressure distribution, lift, drag, and circulation as functions of time are presented.

Lee, G. J.↗