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At least 613 records · Page 34

Experimental surface flow patterns and flow-field phenomena of a delta-wing space-shuttle orbiter

Composite photographs of the surface flow and shadowgraphs of the shock wave pattern are presented that depict the hypersonic flow field of a typical delta wing space shuttle orbiter. Results from a wind tunnel test in air are given in side, oblique, and projected plan views for angles of attack from 0 to 60 deg. The tests were conducted at a Mach number of 7.4 and for Reynolds numbers based on body length of 6,000,000 and 9,000,000. The interrelationship is shown for the intersecting bow and wing leading edge waves with the surface flow for angles of attack for which: (1) the leading edge wave is attached, and (2) the leading edge wave is detached.

Cleary, J. W.↗

Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

Kim, Hyun D.↗

Mechanistic Similarities and Differences for Hydrogenation of Aromatic Heterocycles and Aliphatic Carbonyls on Sulfided Ru Nanoparticles

This study establishes the contrasting reactivity trends for the hydrogenation of aromatic hydrocarbons (AHCs = C n H m X, X = N, S, O, and C) and aliphatic carbonyls [R a C(O)R b ; R a = alkyl group, R b = CH 3 or H] on sulfided Ru clusters arising from the difference in the elementary proton and hydride attack catalytic sequence. Both reactions require sequential additions of a proton from either Ru n+ –(SH 2 ) or S 2 –(H δ+ ) species and a hydride from Ru n+ –(H δ– ) species to the unsaturated C=X bonds. For the five-membered-ring aromatic heterocycles (AHCs = pyrrole, thiophene, and furan), an initial proton addition limits the catalytic turnovers; thus, their hydrogenation reactivity increases with increasing gas-phase proton affinities of the AHCs. Pyridine as the more basic six-membered N-AHC is more susceptible to protonation; therefore, it is more reactive, and its initial proton addition is quasi-equilibrated, followed by the kinetically relevant hydride addition. Conversely, aliphatic carbonyls prefer to undergo hydrogen additions in a reverse sequence, where a Ru n+ –(H δ– ) hydride initially attacks the electron-deficient carbonyl C atom as the kinetically relevant step before a subsequent rapid S 2 –(H δ+ ) proton addition on the electron-rich O atom, as confirmed by isotopic exchange studies with butanal-D 2 and 1-butanol-D 2 mixtures and density functional theory calculations on S-deficient RuS 2 (100) surfaces. For these reasons, their hydrogenation reactivity increases with increasing gas-phase hydride affinities of the carbonyls. On metallic Ru surfaces without sulfur, hydrogen adatoms (H*) are the only reactive hydrogen species; the reactivity of their attack on aromatic heterocycles increases with increasing reactant proton affinities much more sensitively than that on sulfided Ru surfaces. This work illustrates the distinct catalytic roles of the diverse hydrogen species in hydrogenation—the interplay between proton and hydride additions has marked catalytic consequences in shaping the free energy landscape of the reactions, which in turn leads to the observed kinetic dependences, kinetic parameters, reactivity trends, and scaling relations between the measured barriers and the appropriate kinetic descriptors, that is, proton affinity of AHCs and hydride affinity of carbonyls in hydrogenation catalysis. Here, these mechanistic similarities and differences provide explanations of the observed reactivity trends and thus have profound implications for industrial hydrogenation and hydrotreating catalysis.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗

Current Methods for Modeling and Simulating Icing Effects on Aircraft Performance, Stability and Control

Icing alters the shape and surface characteristics of aircraft components, which results in altered aerodynamic forces and moments caused by air flow over those iced components. The typical effects of icing are increased drag, reduced stall angle of attack, and reduced maximum lift. In addition to the performance changes, icing can also affect control surface effectiveness, hinge moments, and damping. These effects result in altered aircraft stability and control and flying qualities. Over the past 80 years, methods have been developed to understand how icing affects performance, stability and control. Emphasis has been on wind tunnel testing of two-dimensional subscale airfoils with various ice shapes to understand their effect on the flow field and ultimately the aerodynamics. This research has led to wind tunnel testing of subscale complete aircraft models to identify the integrated effects of icing on the aircraft system in terms of performance, stability, and control. Data sets of this nature enable pilot in the loop simulations to be performed for pilot training, or engineering evaluation of system failure impacts or control system design.

Ralvasky, Thomas P.↗

Current Methods Modeling and Simulating Icing Effects on Aircraft Performance, Stability, Control

Icing alters the shape and surface characteristics of aircraft components, which results in altered aerodynamic forces and moments caused by air flow over those iced components. The typical effects of icing are increased drag, reduced stall angle of attack, and reduced maximum lift. In addition to the performance changes, icing can also affect control surface effectiveness, hinge moments, and damping. These effects result in altered aircraft stability and control and flying qualities. Over the past 80 years, methods have been developed to understand how icing affects performance, stability, and control. Emphasis has been on wind-tunnel testing of two-dimensional subscale airfoils with various ice shapes to understand their effect on the flowfield and ultimately the aerodynamics. This research has led to wind-tunnel testing of subscale complete aircraft models to identify the integrated effects of icing on the aircraft system in terms of performance, stability, and control. Data sets of this nature enable pilot-in-the-loop simulations to be performed for pilot training or engineering evaluation of system failure impacts or control system design.

Ratvasky, Thomas P.↗

An engineering aerodynamic heating method for hypersonic flow

A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.

Riley, Christopher J.↗

An Engineering Aerodynamic Heating Method for Hypersonic Flow

A capability to calculate surface heating rates has been incorporated in an approximate three-dimensional inviscid technique. Surface streamlines are calculated from the inviscid solution, and the axisymmetric analog is then used along with a set of approximate convective-heating equations to compute the surface heat transfer. The method is applied to blunted axisymmetric and three-dimensional ellipsoidal cones at angle of attack for the laminar flow of a perfect gas. The method is also applicable to turbulent and equilibrium-air conditions. The present technique predicts surface heating rates that compare favorably with experimental (ground-test and flight) data and numerical solutions of the Navier-Stokes (NS) and viscous shock-layer (VSL) equations. The new technique represents a significant improvement over current engineering aerothermal methods with only a modest increase in computational effort.

Riley, Christopher J.↗

Flow Characteristics About Two Thin Wings of Low Aspect Ratio Determined from Surface Pressure Measurements Obtained in Flight at Mach Numbers from 0.73 to 1.90

Surface pressure measurements were obtained at three chordwise stations on the wings of the X-3 and X-lE airplanes at Mach numbers from 0.73 to 1.13 for the X-3, and from 0.82 to 1.90 for the X-IE. Leading-edge separation is present on the X-3 wing at a Mach number of about 0.73 and an angle of attack of about 6 deg. However., when the Mach number is increased to 0.88, the trailing-edge separation dominates the pressure distribution and no leading-edge separation is visible although it is anticipated at the higher angles of attack shown. Conversely, the X-lE wing shows no indication of leading-edge separation within the scope of this investigation, but an overexpansion immediately behind the leading edge is present at a Mach number of approximately 0.82. Two separate normal shocks are present on the X-3 wing at a Mach number of about 0.88 and at a low angle of attack as an effect of wing geometry. These shocks merge to form a single shock when the angle of attack is increased to about 6 deg. At supersonic speeds the upper-surface expansion on the X-lE wing is limited by the approach of the pressure coefficients to the pressure coefficient for a vacuum.

Taillon, Norman V.↗

STOVL CFD model test case

The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration have been measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60 deg cropped delta wing platform, blended fuselage and two circular in-line jets that exit perpendicularly from the flat lower surface. The measured flows have a maximum freestream Mach number of 0.2. Model angle of attack is varied between -10 and +20 deg. The flow is ambient temperature in both jet exits and the nozzle pressure ratios are varied between 1 and 3. The data presented includes forces and moments, pressures measured at 281 surface pressure ports and the pressures of the jets. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exits to guide boundary condition selection for the planned computations. Flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.

Roth, Karlin R.↗

Application of localized surface heating to actively control the boundary layer separation

The use of localized surface heating to improve the performance of the airfoil of a fixed wing or rotating blade at high angle of attack is considered. Separation control by active means can be accomplished by acoustic perturbation of the separated area above the airfoil or by localized surface heating in a pressure gradient region. Active control by sound in the vicinity of the leading edge of an airfoil is shown to greatly reduce the area of separation. Therefore, the overall circulation is sufficiently changed to result in an increase in lift and a reduction in drag.

El-Hady, Nabit M.↗

Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 0.8, 1.3, and 1.75

A typical missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 0.8, 1.3, and 1.75 and Reynolds number of 625,000 based on body diameter. Data were obtained at angles of attack ranging from 0 deg to 24 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). In addition, two different sets of canards and tail surfaces were investigated. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 7 deg with canard deflections of about 10 deg. Also, the tail arrangements studied provided ample pitch stability.

Kassner, D. L.↗

Characterization of Naturally-Occurring Particles in the Nasa Langley 0.3-m Transonic Cryogenic Tunnel Using PTV

Naturally-occurring particles within the NASA Langley 0.3-m Transonic Cryogenic Tunnel are characterized for their aerodynamic performance using particle tracking velocimetry. Two sets of experiments were conducted to observe different behaviors of the particles. The normal shockwave emanating from the top surface of a supercritical airfoil was used to induce velocity lag in the particles, and the subsequent spatial decay of the velocity was used to estimate the effective diameter of the particles. Mean particle diameters between 1.6 and 1.9 𝝁m were measured, with sizes ranging from 0.2 to 3.5 𝝁m over the entire ensemble. The response of particles to separated flow was investigated. By operating in “high-lift” (low Mach number, high angle of attack) conditions with a semi-span airfoil, the ability of particles to detect separated flow on the upper surface of the airfoil was assessed. Transition from fully attached flow to fully separated flow was observed on the top surface of the airfoil accompanying a variation of angle of attack from 8° to 12°. Examination of velocity distributions indicates less than 10 percent of particle trajectories did not respond to the regions of separated flow. These results are promising, but further facility-specific work is needed to answer the broader question of particle tracking reliability.

Transonic↗

Characterization of Naturally-Occurring Particles in the NASA Langley 0.3-m Transonic Cryogenic Tunnel Using PTV

Naturally-occurring particles within the NASA Langley 0.3-m Transonic Cryogenic Tunnel are characterized for their aerodynamic performance using particle tracking velocimetry. Two sets of experiments were conducted to observe different behaviors of the particles. The normal shockwave emanating from the top surface of a supercritical airfoil was used to induce velocity lag in the particles, and the subsequent spatial decay of the velocity was used to estimate the effective diameter of the particles. Mean particle diameters between 1.6 and 1.9 𝝁m were measured, with sizes ranging from 0.2 to 3.5 𝝁m over the entire ensemble. The response of particles to separated flow was investigated. By operating in “high-lift” (low Mach number, high angle of attack) conditions with a semi-span airfoil, the ability of particles to detect separated flow on the upper surface of the airfoil was assessed. Transition from fully attached flow to fully separated flow was observed on the top surface of the airfoil accompanying a variation of angle of attack from 8° to 12°. Examination of velocity distributions indicates less than 10 percent of particle trajectories did not respond to the regions of separated flow. These results are promising, but further facility-specific work is needed to answer the broader question of particle tracking reliability.

Transonic↗

Effect of wing aspect ratio and flap span on aerodynamic characteristics of an externally blown jet-flap STOL model

An investigation has been conducted to determine the effects of flap span and wing aspect ratio on the static longitudinal aerodynamic characteristics and chordwise and spanwise pressure distributions on the wing and trailing-edge flap of a straight-wing STOL model having an externally blown jet flap without vertical and horizontal tail surfaces. The force tests were made over an angle-of-attack range for several thrust coefficients and two flap deflections. The pressure data are presented as tabulated and plotted chordwise pressure-distribution coefficients for angles of attack of 1 and 16. Pressure-distribution measurements were made at several spanwise stations.

Smith, C. C., Jr.↗

Real gas scale effects on hypersonic laminar boundary-layer parameters including effects of entropy-layer swallowing

Inviscid and viscous (laminar boundary-layer) flow-field calculations under perfect gas hypersonic wind tunnel and equilibrium real gas flight conditions are presented for the windward centerline of the Rockwell International 139 Space Shuttle Orbiter at a 30-deg angle of attack. Correlation parameters for laminar boundary-layer edge quantities and surface heat transfer are developed which properly account for entropy-layer-swallowing effects under both real and perfect gas conditions. Some implications of the proposed correlation parameters on boundary-layer transition are discussed.

Adams, J. C., Jr.↗

A new flow model for highly separated airfoil flows at low speeds

An analytical model for separated airfoil flows is presented which is based on experimentally observed physical phenomena. These include a free stagnation point aft of the airfoil and a standing vortex in the separated region. A computer program is described which iteratively matches the outer potential flow, the airfoil turbulent boundary layer, the separated jet entrainment, mass conservation in the separated bubble, and the rear stagnation pressure. Separation location and pressure are not specified a priori. Results are presented for surface pressure coefficient and compared with experiment for three angles of attack for a GA(W)-1, 17% thick airfoil.

Zumwalt, G. W.↗

Acid Solutions for Etching Corrosion-Resistant Metals

New study characterized solutions for etching austenitic stainless steels, nickel-base alloys, and titanium alloys (annealed). Solutions recommended for use remove at least 0.4 mil of metal from surface in less than an hour. Solutions do not cause intergranular attack on metals for which they are effective, when used under specified conditions.

Simmons, J. R.↗

Flowfield computation of an aerobrake orbital transfer vehicle

Viscous real gas and ideal gas supersonic flowfields over the forebody of a aeroassisted orbital transfer vehicle are determined using a unsteady factored implicit algorithm. Air in chemical equilibrium is considered and its local thermodynamic properties are computed by an equilibrium composition method. Numerical solutions are obtained for both real and ideal gases at a Mach number of 30 and at angles of attack up to 20 degrees. Shock stand-off distances and surface pressure distributions are presented for the gas models with and without viscous effects. For the freestream conditions considered, viscous effects dominate the flow.

Balakrishnan, A.↗