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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 253 records · Page 14

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NASA NTRS1998

Inviscid Interactions Between Wake Vortices and Shear Layers

Aircraft trailing vortices can be influenced significantly by atmospheric conditions such as crosswind, turbulence, and stratification. According to the NASA 1994 and 1995 field measurement program in Memphis, Tennessee, the descending aircraft wake vortices could stall or be deflected at the top of low-level temperature inversions that usually produce pronounced shear zones. Numerical simulations of vortex/shear interactions with ground effects have been performed by several groups. Burnham used a series of evenly spaced line vortices at a particular altitude to model the ground shear layer of the cross- wind. He found that the wind shear was swept up around the downwind vortex and caused the downwind vortex to move upward, and claimed that the effect was actually produced by the vertical gradient in the wind shear rather than by the wind shear directly, because uniformly distributed wind-shear vortices would have no effect on the trailing vortex vertical motion. Recently, Proctor et al. numerically tested the effects of narrow shear zones on the behavior of the vortex pair, motivated by the observation of the Memphis field data. The shear-layer sensitivity tests indicated that the downwind vortex was more sensitive and deflected to a higher altitude than its upwind counterpart. The downstream vortex contained vorticity of opposite sign to that of the shear. There was no detectable preference for the downwind vortex (or upwind vortex) to weaken (or strengthen) at a greater rate.

Zheng, Z. C.↗

NASA NTRS2014

Computational Simulations of Convergent Nozzles for the AIAA 1st Propulsion Aerodynamics Workshop

Computational Fluid Dynamics (CFD) simulations were completed for a series of convergent nozzles in participation of the American Institute of Aeronautics and Astronautics (AIAA) 1st Propulsion Aerodynamics Workshop. The simulations were performed using the Wind-US flow solver. Discharge and thrust coefficients were computed for four axisymmetric nozzles with nozzle pressure ratios (NPR) ranging from 1.4 to 7.0. The computed discharge coefficients showed excellent agreement with available experimental data; the computed thrust coefficients captured trends observed in the experimental data, but over-predicted the thrust coefficient by 0.25 to 1.0 percent. Sonic lines were computed for cases with NPR >= 2.0 and agreed well with experimental data for NPR >= 2.5. Simulations were also performed for a 25 deg. conic nozzle bifurcated by a flat plate at NPR = 4.0. The jet plume shock structure was compared with and without the splitter plate to the experimental data. The Wind-US simulations predicted the shock structure well, though lack of grid resolution in the plume reduced the sharpness of the shock waves. Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulations and Detached Eddy Simulations (DES) were performed at NPR = 1.6 for the 25 deg conic nozzle with splitter plate. The simulations predicted vortex shedding from the trailing edge of the splitter plate. However, the vortices of URANS and DES solutions appeared to dissipate earlier than observed experimentally. It is believed that a lack of grid resolution in the region of the vortex shedding may have caused the vortices to break down too soon

DES↗

NASA NTRS1988

Flight tests of external modifications used to reduce blunt base drag

The effectiveness of a trailing disk (the trapped vortex concept) in reducing the blunt base drag of an 8-in diameter body of revolution was studied from measurements made both in flight and in full-scale wind-tunnel tests. The experiment demonstrated the significant base drag reduction capability of the trailing disk to Mach 0.93. The maximum base drag reduction obtained from a cavity tested on the flight body of revolution was not significant. The effectiveness of a splitter plate and a vented-wall cavity in reducing the base drag of a quasi-two-dimensional fuselage closure was studied from base pressure measurements made in flight. The fuselage closure was between the two engines of the F-111 airplane; therefore, the base pressures were in the presence of jet engine exhaust. For Mach numbers from 1.10 to 1.51, significant base drag reduction was provided by the vented-wall cavity configuration. The splitter plate was not considered effective in reducing base drag at any Mach number tested.

Powers, Sheryll Goecke↗

NASA NTRS1988

Flight tests of external modifications used to reduce blunt base drag

Powers, Sheryll Goecke↗

NASA NTRS1973

Aircraft wake dissipation by sinusoidal instability and vortex breakdown.

Sinusoidal instability of aircraft trailing vortices was induced by differentially oscillating inboard and outboard flaps on a model wing in a ship towing tank. Measured amplification rates qualitatively agree with theoretical predictions. Axial velocities in the vortex core were directed towards the wing and measured to be approximately 25% of the tow speed. Vortex breakdown was observed to occur along trailers undergoing sinusoidal instability near but ahead (towards the wing) of positions of maximum trailer separation. It is shown that axisymmetric pressure gradients are imposed along the vortex core by the other sinusoidally deformed trailer and are responsible for the observed changes in core diameter. A theoretical model predicts the short-time behavior of the vortex core and shows that the response of the vortex depends crucially on the axial velocity in the vortex.

Bilanin, A. J.↗

NASA NTRS1984

Prediction of vortex lift on interacting delta wings in incompressible flow

Vortex flow modeling is used to calculate the steady inviscid incompressible flow past one, two and three delta wing configurations. The wings are modeled with vortex lattices and the leading and trailing-edge sheets are modeled by segmented straight vortex filaments. Aerodynamic characteristics are obtained for a range of geometry and angle of attack.

Dodbele, S. S.↗

NASA NTRS2007

Flaperon Modification Effect on Jet-Flap Interaction Noise Reduction for Chevron Nozzles

Jet-flap interaction (JFI) noise can become an important component of far field noise when a flap is immersed in the engine propulsive stream or is in its entrained region, as in approach conditions for under-the-wing engine configurations. We experimentally study the effect of modifying the flaperon, which is a high speed aileron between the inboard and outboard flaps, at both approach and take-off conditions using scaled models in a free jet. The flaperon modifications were of two types: sawtooth trailing edge and mini vortex generators (vg s). Parametric variations of these two concepts were tested with a round coaxial nozzle and an advanced chevron nozzle, with azimuthally varying fan chevrons, using both far field microphone arrays and phased microphone arrays for source diagnostics purposes. In general, the phased array results corroborated the far field results in the upstream quadrant pointing to JFI near the flaperon trailing edge as the origin of the far field noise changes. Specific sawtooth trailing edges in conjunction with the round nozzle gave marginal reduction in JFI noise at approach, and parallel co-rotating mini-vg s were somewhat more beneficial over a wider range of angles, but both concepts were noisier at take-off conditions. These two concepts had generally an adverse JFI effect when used in conjunction with the advanced chevron nozzle at both approach and take-off conditions.

Thomas, Russell H.↗

NASA NTRS1986

Flight and wind-tunnel measurements showing base drag reduction provided by a trailing disk for high Reynolds number turbulent flow for subsonic and transonic Mach numbers

The effectiveness of a trailing disk, or trapped vortex concept, in reducing the base drag of a large body of revolution was studied from measurements made both in flight and in a wind tunnel. Pressure data obtained for the flight experiment, and both pressure and force balance data were obtained for the wind tunnel experiment. The flight test also included data obtained from a hemispherical base. The experiment demonstrated the significant base drag reduction capability of the trailing disk to Mach 0.93 and to Reynolds numbers up to 80 times greater than for earlier studies. For the trailing disk data from the flight experiment, the maximum decrease in base drag ranged form 0.08 to 0.07 as Mach number increased from 0.70 to 0.93. Aircraft angles of attack ranged from 3.9 to 6.6 deg for the flight data. For the trailing disk data from the wind tunnel experiment, the maximum decrease in base and total drag ranged from 0.08 to 0.05 for the approximately 0 deg angle of attack data as Mach number increased from 0.30 to 0.82.

Powers, Sheryll Goecke↗

NASA NTRS2013

Coupled Vortex-Lattice Flight Dynamic Model with Aeroelastic Finite-Element Model of Flexible Wing Transport Aircraft with Variable Camber Continuous Trailing Edge Flap for Drag Reduction

This paper presents a coupled vortex-lattice flight dynamic model with an aeroelastic finite-element model to predict dynamic characteristics of a flexible wing transport aircraft. The aircraft model is based on NASA Generic Transport Model (GTM) with representative mass and stiffness properties to achieve a wing tip deflection about twice that of a conventional transport aircraft (10% versus 5%). This flexible wing transport aircraft is referred to as an Elastically Shaped Aircraft Concept (ESAC) which is equipped with a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for active wing shaping control for drag reduction. A vortex-lattice aerodynamic model of the ESAC is developed and is coupled with an aeroelastic finite-element model via an automated geometry modeler. This coupled model is used to compute static and dynamic aeroelastic solutions. The deflection information from the finite-element model and the vortex-lattice model is used to compute unsteady contributions to the aerodynamic force and moment coefficients. A coupled aeroelastic-longitudinal flight dynamic model is developed by coupling the finite-element model with the rigid-body flight dynamic model of the GTM.

Flexible Aicraft↗

NASA NTRS2004

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

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.

Jenkins, Luther N.↗

NASA NTRS1986

Effect of blunt trailing edge on rotor broadband noise

The production of high-frequency broadband noise by turbulent vortex shedding from rotor blades with blunt trailing edges is investigated analytically. The derivation of the governing equations, analogous to that of Kim and George (1982) for boundary-layer/trailing-edge noise, is explained, and numerical results are compared with the experimental data of Hubbard et al. (1981) and Lowson et al. (1972) in graphs. It is shown that vortex-shedding noise is a significant component of blunt-trailing-edge rotor broadband noise and that the analytical method employed gives reasonable predictions. The need for a better empirical expression for the normalized spectrum and for more measurements of surface pressure fluctuations near blunt trailing edges is indicated.

Chou, S.-T.↗

NASA NTRS2005

The Effect of Pulse Length and Ejector Radius on Unsteady Ejector Performance

The thrust augmentation of a set of ejectors driven by a shrouded Hartmann-Sprenger tube has been measured at four different frequencies. Each frequency corresponded to a different length to diameter ratio of the pulse of air leaving the driver shroud. Two of the frequencies had length to diameter ratios below the formation number, and two above. The formation number is the value of length to diameter ratio below which the pulse converts to a vortex ring only, and above which the pulse becomes a vortex ring plus a trailing jet. A three level, three parameter Box-Behnken statistical design of experiment scheme was performed at each frequency, measuring the thrust augmentation generated by the appropriate ejectors from the set. The three parameters were ejector length, radius, and inlet radius. The results showed that there is an optimum ejector radius and length at each frequency. Using a polynomial fit to the data, the results were interpolated to different ejector radii and pulse length to diameter ratios. This showed that a peak in thrust augmentation occurs when the pulse length to diameter ratio equals the formation number, and that the optimum ejector radius is 0.87 times the sum of the vortex ring radius and the core radius.

Wilson, Jack↗

NASA NTRS1982

Secondary effects in combustion instabilities leading to flashback

The secondary effects in turbulent combustion instabilities leading to flashback are investigated, including those due to buoyancy and contraction at the combustor outlet. Experiments were conducted in an oblong, rectangular cross-section combustion tunnel, where the effects of a bluff-body flames holder were generated by a rear-facing step behind a streamlined inlet nozzle. The results of experiments leading to flashback with the step mounted at the bottom of the combustion chamber were compared to those of experiments in which it was located at the top. Irrespective of the flow obstructions introduced downstream, the critical equivalence ratio for flashback was consistently lower with the step at the bottom, indicating that buoyancy was enhancing the growth of the recirculation zone that pushed the flame upstream and caused flashback. The contraction at the end of the combustion chamber had a promoting influence on the process of vortex pairing, re-enforcing the influence of the trailing vortices over that of the recirculation vortex system, and thereby curbing the tendency to flashback. Provided that the flow velocity was low, however, the characteristic features of combustion instabilities leading to flashback in the absence of contraction could still be established in its presence.

Vaneveld, L.↗

NASA NTRS1983

Vortex flap flow reattachment line and subsonic longitudinal aerodynamic data on 50 deg to 74 deg Delta wings on common fuselage

Positions of the primary vortex flow reattachment line and longitudinal aerodynamic data were obtained at Mach number 0.3 for a systematic series of vortex flaps on delta wing body configurations with leading edge sweeps of 50, 58, 66, and 74 deg. The investigation was performed to study the parametric effects of wing sweep, vortex flap geometry and deflection, canards, and trailing edge flaps on the location of the primary vortex reattachment line relative to the flap hinge line. The vortex reattachment line was located via surface oil flow photographs taken at selected angles of attack. Force and moment measurements were taken over an angle of attack range of -1 deg to 22 deg at zero sideslip angle for many configurations to further establish the data base and to assess the aforementioned parametric effects on longitudinal aerodynamics. Both the flow reattachment and aerodynamic data are presented.

Frink, N. T.↗

NASA NTRS1985

Vortex-dominated flow with viscous core structure

Recent theoretical studies of vortex-dominated flows are reviewed with special emphasis on those for which the viscous core structures play an important role. The problems to be described are: The interaction and merging of two-dimensional vortices and of curved vortex filaments, the roll-up and decay of trailing far wakes, and the initiation of vortex breakdown. The analysis utilizes finite-difference solutions of the Navier-Stokes equations complemented by asymptotic expansion techniques.

Liu, C. H.↗

DOE OSTI2023

Evaluating Dynamic Stall-Onset Criteria for Mixed and Trailing-Edge Stall

We evaluate two leading-edge-based dynamic stall-onset criteria (namely, the maximum magnitudes of the leading-edge suction parameter and the boundary enstrophy flux) for mixed and trailing-edge stall. These criteria have been shown to successfully predict the onset of leading-edge stall at Reynolds numbers of [Formula: see text], where the leading-edge suction drops abruptly. However, for mixed/trailing-edge stall, leading-edge suction tends to persist even when there is significant trailing-edge reversed flow and stall is underway, necessitating further investigation into the effectiveness of these criteria. Using wall-resolved large-eddy simulations and the unsteady Reynolds-averaged Navier–Stokes method, we simulate one leading-edge stall and three mixed/trailing-edge stall cases at Reynolds numbers of [Formula: see text] and [Formula: see text]. We contrast the progression of flow features such as trailing-edge separation and vortex formation across different stall types and evaluate the stall-onset criteria relative to critical points in the flow. We find that the criteria nearly coincide with the instance of leading-edge suction collapse and are reached in advance of dynamic stall vortex formation and lift stall for all four cases. We conclude that the two criteria effectively signal dynamic stall onset in cases where the dynamic stall vortex plays a prominent role.

Engineering↗

NASA NTRS1973

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils.

High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 x 1,000,000 to 6.2 x 1,000,000 on a rectangular wing of NACA 63-009 airfoil section. A wide selection of leading-edge serrations were also added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large peak in rms pressure, which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related mathematically to the airfoil trailing-edge and boundary-layer thicknesses.

Schwind, R. G.↗

NASA NTRS1973

The effects of leading-edge serrations on reducing flow unsteadiness about airfoils, an experimental and analytical investigation

High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 times one million to 6.2 times one million on a rectangular wing of NACA 63-009 airfoil section. Measurements were also obtained with a wide selection of leading-edge serrations added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large apatial peak in rms pressure. Frequency analysis of the pressure signals in this region show a large, high-frequency energy peak which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related to the trailing-edge and boundary-layer thicknesses.

Schwind, R. G.↗