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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Space Shuttle Orbiter nose cap and wing leading edge certification test program

A reinforced carbon-carbon thermal protection system is used on the space shuttle orbiter vehicle's nose cap and wing leading edge regions where temperatures reach 1538 C (2800 F). To verify the analyses used to certify reinforced carbon-carbon for the first flight and operational missions, a multi-environment incremental test program was developed and implemented through the combined efforts of Rockwell International and the National Aeronautics and Space Administration. Three separate facilities at the Johnson Space Center were used to subject full-scale nose cap and wing leading edge test articles to simulated critical launch, on-orbit, and atmospheric entry environments.

Suppanz, M. J.↗
NASA NTRS1992

Measurements in a leading-edge separation bubble due to a simulated airfoil ice accretion

The separation bubble formed on an airfoil at low Reynolds number behind a simulated leading-edge glaze ice accretion is studied experimentally. Surface pressure and split hot-film measurements as well as flow visualization studies of the bubble reattachment point are reported. The simulated ice generates an adverse pressure gradient that causes a laminar separation bubble of the long bubble type to form. The boundary layer separates at a location on the ice accretion that is independent of angle of attack and reattaches at a downstream location 5-40 percent chord behind the leading edge, depending on the angle of attack. Velocity profiles show a large region of reverse flow that extends up from the airfoil surface as much as 2.5 percent chord. After reattachment, a thick distorted turbulent boundary layer exists. The separation bubble growth and reattachment are clearly seen in the plots of boundary-layer momentum thickness vs surface distance. Local minima and maxima in the boundary-layer momentum thickness development compare well with the shear layer transition point as indicated by the surface pressures and the reattachment point as measured from surface oil flow, respectively.

Bragg, M. B.↗
NASA NTRS1984

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

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

Walitt, L.↗
NASA NTRS1978

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

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.

Arena, A. V.↗
NASA NTRS1983

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

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

Manro, M. E.↗
NASA NTRS2023

Applications of Flow Control to Wing High-Lift Leading Edge Devices on a Commercial Aircraft

Active flow control was applied to the leading edge region of a representative future short/medium-range twin-engine airplane to improve aerodynamic performance during high-lift operations. The study is aimed at enhanced lift over the practical angle of attack range, including stall, and at reduced drag. These benefits translate to airplane performance improvements, such as longer range or larger payload. Various flow control applications were explored using Computational Fluid Dynamics and the aerodynamic performance enhancements were benchmarked against the baseline configuration. The computational analyses are used to quantify aerodynamic benefits, as well as the input required for actuation. The results were used in a system integration study for identifying potential practical implementations, which are described in a companion paper. Combined with the integration analysis, the objective of this project is to identify the most promising flow control candidates that potentially provide material net airplane level enhancements using onboard fluidic sources. Depending on the implementation of active flow control, the current study indicates that up to 1.5% net improvement in L/D at takeoff and 4% increase in maximum lift during landing are potentially achievable, after accounting for factors of system integration.

CFD↗
NASA NTRS2005

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

In support of recommendations by the Columbia Accident Investigation Board, a team has been studying the effect of debris impacting the reinforced carbon-carbon panels of the shuttle leading edge. The objective of this study was to examine the effect of varying parameters of the debris trajectory on the damage tolerance. Impacts at the upper and lower surface and the apex of the leading edge were examined. For each location, trajectory variances included both the alpha and beta directions. The results of the analysis indicated in all cases the beta sweep decreased the amount of damage to the panel. The increases in alpha resulted in a significant increase in damage to the RCC panel. In particular, for the lower surface, where the alpha can increase by 10 degrees, there was a nearly 40% increase in the impulse. As a result, it is recommended that for future analyses, a 10 degree offset in alpha from the nominal trajectory is included for impacts on the lower surface. It is also recommended to assume a straight aft, or zero beta, trajectory for a more conservative analysis.

Spellman, Regina L.↗
NASA NTRS1979

Exploratory study of the influence of wing leading-edge modifications on the spin characteristics of a low-wing single-engine general aviation airplane

A wide variety of model and airplane tests were conducted to explore the effects of wing leading-edge modifications on the stalling and spinning characteristics of a low-wing general aviation research airplane configuration. The results presented herein discuss the effects of leading-edge airfoil modifications applied to the full wing span and to partial-span locations. The results obtained in the study indicate that the wing modifications can produce large effects (both favorable or unfavorable) on spin resistance, developed spin characteristics, and spin recovery.

Dicarlo, D. J.↗
NASA NTRS1973

Wind tunnel investigation of effects of variations in Reynolds number and leading-edge treatment on the aerodynamic characteristics of an externally blown jet-flap configuration

An investigation has been conducted in a full-scale tunnel to determine the effects of variations in Reynolds number and leading-edge treatment on the aerodynamic characteristics of an externally blown jet-flap transport configuration. The model had a double-slotted trailing-edge flap and was powered by four high-bypass-ratio turbofan engines. Tests were performed by using each of three leading-edge devices (a 30-percent-chord flap and 15- and 25-percent-chord slats) at Reynolds numbers from 0.47 x one million to 1.36 x one million thrust coefficients up to 3.5. The use of a 25-percent-chord slat was found to be more effective than a 15-percent-chord slat or a 30-percent-chord flap in extending the stall angle of attack and in minimizing the loss of lift after the stall. The large slat was also effective in reducing the rolling moments that occurred when the engine-out wing stalled first.

Parlett, L. P.↗
NASA NTRS1980

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

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

Dicarlo, D. J.↗
NASA NTRS1961

Heat-Transfer Measurements at a Mach Number of 4.95 on Two 60 deg Swept Delta Wings with Blunt Leading Edges and Dihedral Angles of 0 deg and 45 deg

An experimental investigation was conducted to evaluate the heat-transfer characteristics of two 60 deg swept delta wings with cylindrical leading edges of 0.25-inch radii and dihedral angles of 0 deg and 45 deg. The tests were conducted at a Mach number of 4.95 and a stagnation temperature of 400 F. The The test-section unit Reynolds number was varied from 1.95 x 10(exp 6) to 12.24 x 10(exp 6) per foot. The results of the investigation indicated that, in a plane normal to the leading edge, the laminar-flow heat-transfer distribution was in good agreement with two-dimensional blunt-body theory. The stagnation-line heat-transfer level could be predicted from two-dimensional blunt-body theory provided the stagnation-line heat-transfer coefficient was assumed to vary as the cosine of the effective sweep. A comparison of the heating rates to the 0 deg dihedral wing (planform sweep of 60 deg) and the 45 deg dihedral wing (planform sweep of 69.3 deg) with equal panel sweep and panel area indicated that the stagnation-line heat-transfer coefficient for the 45 deg dihedral wing could be as much as 40 percent less than the stagnation-line heat-transfer coefficient for the 0 deg dihedral wing at both equal angles of attack and equal lifts. The laminar-flow heat-transfer rate to both wings outside the vicinity of the stagnation line was essentially equal.

Stainback, P. Calvin↗
NASA NTRS1978

Effects of wing leading-edge flap deflections on subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing

An investigation was conducted to determine the effects of wing leading-edge flap deflections on the subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing. The tests were conducted at Mach numbers from 0.40 to 0.85, corresponding to Reynolds numbers (based on wing mean geometric chord) of 2.37 x 1,000,000 to 4.59 x 1,000,000 and at angles of attack from -3 deg to 22 deg. The configurations under study included a wing-fuselage configuration and a wing-fuselage-strake configuration. Each configuration had multisegmented, constant-chord leading-edge flaps which could be deflected independently or in various combinations.

Henderson, W. P.↗
NASA NTRS1987

Numerical analysis of a NACA0012 airfoil with leading edge ice accretions

Analysis of a NACA0012 airfoil with leading edge ice has been performed using a Navier-Stokes code coupled with a grid generation code. The computed results were compared to experimental information obtained for an airfoil with a well defined artificial ice shape. The computations were performed at angles of attack ranging from zero to ten degrees. This range is sufficient to show the development of the separation bubble aft of the ice shape on both the upper and lower surfaces. Velocity profile plots in the separation bubble are examined in order to determine if recirculation patterns are predicted properly and if separation and reattachment points are found within the resolution of the experimental information. Also, the massive separation near the point of stall is examined in order to more accurately evaluate the lift coefficient curve in that region. Lift, drag, and moment coefficients are computed and compared to experiment.

Potapczuk, Mark G.↗
NASA NTRS1993

Active control of wing rock of a delta wing at post-stall using tangential leading edge blowing

Post-stall roll control utilizing tangential leading edge blowing is demonstrated in a wind tunnel on a delta wing model that exhibited wing rock. The dampening effect of symmetric blowing alone on wing rock is found to be effective up to a certain maximum amount of blowing. A moderate amount of symmetric blowing was shown to be effective in linearizing the asymmetric blowing static rolling moment responses.

Wong, G. S.↗