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NASA NTRS1993

Effect of tabs on the flow and noise field of an axisymmetric jet

The effect of vortex generators, in the form of small tabs projecting normally into the flow at the nozzle exit, on the characteristics of an axisymmetric jet is investigated experimentally over the jet Mach number range of 0.3-1.81. The tabs eliminate screech noise from supersonic jets and alter the shock structure drastically. They distort the jet cross section and increase the jet spread rate significantly. The distortion produced is essentially the same at subsonic and underexpanded supersonic conditions. Thus, the underlying mechanism must be independent of compressibility effects. A tab with a height as small as 2 percent of the jet diameter, but larger than the efflux boundary-layer thickness, is found to produce a significant effect. Flow visualization reveals that each tab introduces an 'indentation' into the high speed side of the shear layer via the action of streamwise vortices. These vortices are inferred to be of the 'trailing vortex' type rather than of the 'necklace vortex' type. It is apparent that a substantial pressure differential must exist between the upstream and the downstream sides of the tab to effectively produce these trailing vortices. This explains why the tabs are ineffective in the overexpanded flow, as in that case an adverse pressure gradient exists near the nozzle exit which reduces the pressure differential produced by the tab.

Samimy, M.↗

NASA NTRS1999

Boundary Layer Transition in the NTF: HSR Experience and Plans

Efforts towards understanding boundary layer transition characteristics on a High Speed Civil Transport (HSCT)-class configuration in the National Transonic Facility (NTF) are ongoing. The majority of the High Speed Research (HSR) data base in the NTF has free transition on the wing, even at low Reynolds numbers (Rn) attainable in conventional facilities. Limited data has been obtained and is described herein showing the effects of a conventional, Braslow method based wing boundary-layer trip on drag. Comparisons are made using force data polars and surface flow visualization at selected angles-of-attack and Mach number. Minimum drag data obtained in this study suggest that boundary layer transition occurred very near the wing leading edge by a chord Rn of 30 million. Sublimating chemicals were used in the air mode of operation only at low Rn and low angles-of-attack with no flap deflections; sublimation results suggest that the forebody and outboard wing panel are the only regions with significant laminar flow. The process and issues related to the sublimating chemical technique as applied in the NTF are discussed. Beyond the existing experience, status of efforts to develop a production transition detection system applicable to both air and cryogenic nitrogen environments is presented.

Owens, Lewis R., Jr.↗

NASA NTRS1972

Asymmetric lateral-directional characteristics of pointed bodies of revolution at high angles of attack

A low-speed wind-tunnel investigation was conducted in order to determine the cause of asymmetric yawing moments produced by long pointed fuselage nose shapes at high angles of attack. Force tests were conducted with a cone, tangent-ogive model, and a paraboloid of revolution over a range of Reynolds numbers from 0.15 million to 0.35 million for an angle-of-attack range from 0 deg to 75 deg and an angle-of-sideslip range of plus or minus 30 deg. Tuft and smoke flow-visualization tests were also conducted to aid in the analysis. Large asymmetric yawing moments were obtained for the cone and tangent-ogive body at high angles of attack (of the order of 40 to 60 deg). These large moments were caused by asymmetric shedding of vortex sheets off the long pointed nose. The asymmetric moments could be eliminated by use of symmetrically arranged strikes on the nose. The paraboloid of revolution did not produce a strong asymmetric flow field at high angles of attack and did not exhibit asymmetric moments.

Coe, P. L., Jr.↗

NASA NTRS2000

Development of a Large Field of View Shadowgraph System for a 16 Ft. Transonic Wind Tunnel

A large field of view shadowgraph flow visualization system for the Langley 16 ft. Transonic Tunnel (16 ft.TT) has been developed to provide fast, low cost, aerodynamic design concept evaluation capability to support the development of the next generation of commercial and military aircraft and space launch vehicles. Key features of the 16 ft. TT shadowgraph system are: (1) high resolution (1280 X 1024) digital snap shots and sequences; (2) video recording of shadowgraph at 30 frames per second; (3) pan, tilt, & zoom to find and observe flow features; (4) one microsecond flash for freeze frame images; (5) large field of view approximately 12 X 6 ft; and (6) a low maintenance, high signal/noise ratio, retro-reflective screen to allow shadowgraph imaging while test section lights are on.

Talley, Michael A.↗

NASA NTRS1992

Forebody vortex control with miniature, rotatable nose-boom strakes

Experiments have been conducted in water and wind tunnel facilities to investigate the effectiveness of using rotatable miniaturized forebody nose-boom strakes to manipulate and control fighter aircraft forebody vortices at moderate to high angles of attack to provide potential aerodynamic control power in yaw. Water tunnel tests including flow visualization and yaw moment measurements were conducted on several different models of the F-16 fighter. Wind tunnel force and moment tests investigated the effects of several nose-boom strake configurations on an isolated 1/8th-scale X-29A forebody model. The primary variables of interest were strake size, configuration and roll angle orientation on the nose-boom. Both experiments showed clearly that small nose-boom strakes can be very effective in controlling the forebody vortex flowfield. Different degrees of flow asymmetry can be obtained by rotating a single or a pair of small strakes located on the nose-boom. With large fixed forebody strakes, such as those on the X-29A configuration, the effectiveness of the nose-boom strakes is significantly reduced.

Suarez, Carlos J.↗

NASA NTRS2004

Experimental, Theoretical, and Computational Investigation of Separated Nozzle Flows

A detailed experimental, theoretical, and computational study of separated nozzle flows has been conducted. Experimental testing was performed at the NASA Langley 16-Foot Transonic Tunnel Complex. As part of a comprehensive static performance investigation, force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a sub-scale, non-axisymmetric, two-dimensional, convergent- divergent nozzle. In addition, two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and algebraic Reynolds stress modeling. For reference, experimental and computational results were compared with theoretical predictions based on one-dimensional gas dynamics and an approximate integral momentum boundary layer method. Experimental results from this study indicate that off-design overexpanded nozzle flow was dominated by shock induced boundary layer separation, which was divided into two distinct flow regimes; three- dimensional separation with partial reattachment, and fully detached two-dimensional separation. The test nozzle was observed to go through a marked transition in passing from one regime to the other. In all cases, separation provided a significant increase in static thrust efficiency compared to the ideal prediction. Results indicate that with controlled separation, the entire overexpanded range of nozzle performance would be within 10% of the peak thrust efficiency. By offering savings in weight and complexity over a conventional mechanical exhaust system, this may allow a fixed geometry nozzle to cover an entire flight envelope. The computational simulation was in excellent agreement with experimental data over most of the test range, and did a good job of modeling internal flow and thrust performance. An exception occurred at low nozzle pressure ratios, where the two-dimensional computational model was inconsistent with the three-dimensional separation observed in the experiment. In general, the computation captured the physics of the shock boundary layer interaction and shock induced boundary layer separation in the nozzle, though there were some differences in shock structure compared to experiment. Though minor, these differences could be important for studies involving flow control or thrust vectoring of separated nozzles. Combined with other observations, this indicates that more detailed, three-dimensional computational modeling needs to be conducted to more realistically simulate shock-separated nozzle flows.

Hunter, Craig A.↗

NASA NTRS2016

A Study of Heat Transfer and Flow Characteristics of Rising Taylor Bubbles

Practical application of flow boiling to ground- and space-based thermal management systems hinges on the ability to predict the systems heat removal capabilities under expected operating conditions. Research in this field has shown that the heat transfer coefficient within two-phase heat exchangers can be largely dependent on the experienced flow regime. This finding has inspired an effort to develop mechanistic heat transfer models for each flow pattern which are likely to outperform traditional empirical correlations. As a contribution to the effort, this work aimed to identify the heat transfer mechanisms for the slug flow regime through analysis of individual Taylor bubbles.An experimental apparatus was developed to inject single vapor Taylor bubbles into co-currently flowing liquid HFE 7100. The heat transfer was measured as the bubble rose through a 6 mm inner diameter heated tube using an infrared thermography technique. High-speed flow visualization was obtained and the bubble film thickness measured in an adiabatic section. Experiments were conducted at various liquid mass fluxes (43-200 kgm2s) and gravity levels (0.01g-1.8g) to characterize the effect of bubble drift velocityon the heat transfer mechanisms. Variable gravity testing was conducted during a NASA parabolic flight campaign.Results from the experiments showed that the drift velocity strongly affects the hydrodynamics and heat transfer of single elongated bubbles. At low gravity levels, bubbles exhibited shapes characteristic of capillary flows and the heat transfer enhancement due to the bubble was dominated by conduction through the thin film. At moderate to high gravity, traditional Taylor bubbles provided small values of enhancement within the film, but large peaks in the wake heat transfer occurred due to turbulent vortices induced by the film plunging into the trailing liquid slug. Characteristics of the wake heat transfer profiles were analyzed and related to the predicted velocity field. Results were compared and shown to agree with numerical simulations of colleagues from EPFL, Switzerland.In addition, a preliminary study was completed on the effect of a Taylor bubble passing through nucleate flow boiling, showing that the thinning thermal boundary layer within the film suppressed nucleation, thereby decreasing the heat transfer coefficient.

capillary flow↗

NASA NTRS1997

Three Dimensional Viscous Flow Field in an Axial Flow Turbine Nozzle Passage

The objective of this investigation is experimental and computational study of three dimensional viscous flow field in the nozzle passage of an axial flow turbine stage. The nozzle passage flow field has been measured using a two sensor hot-wire probe at various axial and radial stations. In addition, two component LDV measurements at one axial station (x/c(sum m) = 0.56) were performed to measure the velocity field. Static pressure measurements and flow visualization, using a fluorescent oil technique, were also performed to obtain the location of transition and the endwall limiting streamlines. A three dimensional boundary layer code, with a simple intermittency transition model, was used to predict the viscous layers along the blade and endwall surfaces. The boundary layers on the blade surface were found to be very thin and mostly laminar, except on the suction surface downstream of 70% axial chord. Strong radial pressure gradient, especially close to the suction surface, induces strong cross flow components in the trailing edge regions of the blade. On the end-walls the boundary layers were much thicker, especially near the suction corner of the casing surface, caused by secondary flow. The secondary flow region near the suction-casing surface corner indicates the presence of the passage vortex detached from the blade surface. The corner vortex is found to be very weak. The presence of a closely spaced rotor downstream (20% of the nozzle vane chord) introduces unsteadiness in the blade passage. The measured instantaneous velocity signal was filtered using FFT square window to remove the periodic unsteadiness introduced by the downstream rotor and fans. The filtering decreased the free stream turbulence level from 2.1% to 0.9% but had no influence on the computed turbulence length scale. The computation of the three dimensional boundary layers is found to be accurate on the nozzle passage blade surfaces, away from the end-walls and the secondary flow region. On the nozzle passage endwall surfaces the presence of strong pressure gradients and secondary flow limit the validity of the boundary layer code.

Ristic, D.↗

NASA NTRSDate not supplied

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper presents and discusses the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of the test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model angle of incidence. Laser Doppler velocimetry (LDV) measurements were made at three model angles of incidence (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data were obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data on the test article and Reynolds-averaged Navier-Stokes CFD results are presented and discussed.

Juncture Flow↗

NASA NTRS1993

Transition on a sharp cone at high enthalpy - New measurements in the shock tunnel T5 at GALCIT

An exploratory study of the laminar, transitional and turbulent boundary layer on a slender cone in hypervelocity flow was conducted in the high-enthalpy shock tunnel T5. A novel flow visualization technique using sodium seeding to increase the sensitivity of conventional interferometric techniques by resonant enhancement of the refractivity of the medium was developed to study its structure. The experiments were designed to cover a large range of specific reservoir enthalpy, ranging from the perfect-gas region to the range where significant oxygen and some nitrogen dissociation and recombination effects may be expected in the boundary layer. Surface heat flux measurements indicate that the transition Reynolds number is approximately the same as that observed in cold hypersonic flow in conventional continuous flow facilities, but significantly lower than that observed in 'quiet' tunnels. No significant effect of reservoir specific enthalpy on transition was observed. The heat flux distribution showed differences between the air and the nitrogen experiments in the laminar regime. These are consistent with the presence of recombination of oxygen at the wall. This view is supported by the fact that this effect is absent in the low-enthalpy region, where the heat flux distribution agrees with the similarity theory.

Germain, P.↗

NASA NTRS1994

Evaluation of flip-flop jet nozzles for use as practical excitation devices

This paper describes the flowfield characteristics of the flip-flop jet nozzle and the potential for using this nozzle as a practical excitation device. It appears from the existing body of published information that there is a lack of data on the parameters affecting the operation of such nozzles and on the mechanism of operation of these nozzles. An attempt is made in the present work to study the important parameters affecting the operation and performance of a flip-flop jet nozzle. Measurements were carried out to systematically assess the effect of varying the nozzle pressure ratio (NPR) as well as the length and volume of the feedback tube on the frequency of oscillation of this device. Flow visualization was used to obtain a better understanding of the jet flowfield and of the processes occurring within the feedback tube. The frequency of oscillation of the flip-flop jet depended significantly on the feedback tube length and volume as well as on the nozzle pressure ratio. In contrast, the coherent velocity perturbation levels did not depend on the above mentioned parameters. The data presented in this paper would be useful for modeling such flip-flop excitation devices that are potentially useful for controlling practical shear flows.

Raman, Ganesh↗

NASA NTRS2016

Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

An experimental investigation of tip vortex flow from a NACA0012 airfoil, pitched periodically at various frequencies, is conducted in a low-speed wind tunnel. Initially, data for stationary airfoil held fixed at various angles-of-attack are gathered. Flow visualization pictures as well as detailed cross-sectional properties areobtained at various streamwise locations using hot-wire anemometry. Data include mean velocity, streamwise vorticity as well as various turbulent stresses. Preliminary data are also acquired for periodically pitched airfoil. These results are briefly presented in this extended abstract.

wing aerodynamics↗

NASA NTRS1972

Flow field measurements for space-shuttle-related cylindrical configurations in hypersonic streams.

Use of surface pressures, heat-transfer rates, and flow-visualization photographs to define the flowfields for simulated infinite cylinders. Four cylindrical cross sections were studied in the experimental program, which was conducted at free-stream Mach numbers of 5 to 15, free-stream unit Reynolds numbers from 2 to 20 million per foot, and angles of attack from 30 to 90 deg. Although swept-cylinder theory provides a reasonable correlation of the measured laminar heat-transfer rates in the plane of symmetry, improved correlation with the measurements is obtained if the three-dimensional nature of the flow is taken into account. In cases where the boundary layer on the windward forebody was turbulent, the large velocity gradients near the chines produced relaminarization of the boundary layer prior to separation. The average heat-transfer-rate in the separated region, which is weakly dependent on the free-stream conditions, was found to be a linear function of the angle of attack. For a given flow condition the effect of the leeward geometry on heat transfer can be correlated by a single geometric parameter which is dependent on the location of separation.

Bertin, J. J.↗

NASA NTRS1984

Assessment of preliminary prediction techniques for wing leading-edge vortex flows at supersonic speeds

The applicability of existing theoretical models for predicting wing leading edge separated flow characteristics for high-lift supersonic flight was examined with regard to previous experimental force, pressure and flow visualization data. Correlations of data on uncambered delta wings revealed that the upper surface normal force and minimum pressure coefficients decreased nonlinearly with increasing angles of attack. The attainable vacuum pressure decreased with increasing Mach number, while the lower surface normal force increased nonlinearly with increasing speed. The LISTAR and VORCAM linear theory aerodynamic codes generated predictions for comparison with the data. LISTAR displayed better agreement with measured vortex strength and position and lifting characteristics than did VORCAM. The Euler code SWINT was ill-suited to calculating wing performance in separated flows at high lift and low supersonic speeds.

Wood, R. M.↗

NASA NTRS2013

Aerodynamic Characterization of a Thin, High-Performance Airfoil for Use in Ground Fluids Testing

The FAA has worked with Transport Canada and others to develop allowance times for aircraft operating in ice-pellet precipitation. Wind-tunnel testing has been carried out to better understand the flowoff characteristics and resulting aerodynamic effects of anti-icing fluids contaminated with ice pellets using a thin, high-performance wing section at the National Research Council of Canada Propulsion and Icing Wind Tunnel. The objective of this paper is to characterize the aerodynamic behavior of this wing section in order to better understand the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination. Aerodynamic performance data, boundary-layer surveys and flow visualization were conducted at a Reynolds number of approximately 6.0×10(exp 6) and a Mach number of 0.12. The clean, baseline model exhibited leading-edge stall characteristics including a leading-edge laminar separation bubble and minimal or no separation on the trailing edge of the main element or flap. These results were consistent with expected 2-D aerodynamics and showed no anomalies that could adversely affect the evaluation of anti-icing fluids and ice-pellet contamination on the wing. Tests conducted with roughness and leading-edge flow disturbances helped to explain the aerodynamic impact of the anti-icing fluids and contamination. The stalling characteristics of the wing section with fluid and contamination appear to be driven at least partially by the effects of a secondary wave of fluid that forms near the leading edge as the wing is rotated in the simulated takeoff profile. These results have provided a much more complete understanding of the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination on this wing section. This is important since these results are used, in part, to develop the ice-pellet allowance times that are applicable to many different airplanes.

Broeren, Andy P.↗

NASA NTRS1989

The structure of separated flow regions occurring near the leading edge of airfoils - including transition

Laser Doppler Velocimeter data, static pressure data, and smoke flow visualization data was obtained and analyzed to correlate with separation bubble data. The Eppler 387 airfoil was focused on at a chord Reynolds number of 100,000 and an angle of attack of 2 deg. Additional data was also obtained from the NACA 663-018 airfoil at a chord Reynolds number of 160,000 and an angle of attack of 12 deg. The structure and behavior of the transition separation bubble was documented along with the redeveloping boundary layer after reattachment over an airfoil at low Reynolds numbers. The understanding of the complex flow phenomena was examined so that analytic methods for predicting their formation and development can be improved. These analytic techniques have applications in the design and performance prediction of airfoils operating in the low Reynolds number flight regime.

Source record↗

NASA NTRS1991

A laser velocimeter investigation of the normal shock-wave boundary layer interaction

Nonintrusive 3-D measurements were made of a normal shock wave/turbulent boundary layer interaction. The measurements were made through a quadrant of a square test section of a continuous flow supersonic wind tunnel in which a normal shock wave had been stabilized. Two dimensional measurements were made throughout the interaction region while 3-D measurements were made near the corner in the vicinity of the shock where 3-D flow effects were expected to be significant. Laser Doppler velocimetry, surface static pressure measurements, and flow visualization techniques were used for two freestream nominal Mach number test cases: 1.6 and 1.3. No turbulence information was obtained. The mean velocity measurements were converted to Mach number by recording the wind tunnel total temperature. Some shock oscillation was present during both of the test cases. After startup of the wind tunnel, the shock oscillated with an amplitude of approx. + or - 1 cm, however, after reaching steady condition, the shock oscillation amplitude was greatly reduced, as evidenced by the laser velocimeter results. The Mach 1.3 test case resulted in a nearly uniform flow without secondary shock waves and with no or very isolated corner separation.

Chriss, Randall M.↗

NASA NTRS2013

Aerodynamic Characterization of a Thin, High-Performance Airfoil for Use in Ground Fluids Testing

The FAA has worked with Transport Canada and others to develop allowance times for aircraft operating in ice-pellet precipitation. Wind-tunnel testing has been carried out to better understand the flowoff characteristics and resulting aerodynamic effects of anti-icing fluids contaminated with ice pellets using a thin, high-performance wing section at the National Research Council of Canada Propulsion and Icing Wind Tunnel. The objective of this paper is to characterize the aerodynamic behavior of this wing section in order to better understand the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination. Aerodynamic performance data, boundary-layer surveys and flow visualization were conducted at a Reynolds number of approximately 6.0 x 10(exp 6) and a Mach number of 0.12. The clean, baseline model exhibited leading-edge stall characteristics including a leading-edge laminar separation bubble and minimal or no separation on the trailing edge of the main element or flap. These results were consistent with expected 2-D aerodynamics and showed no anomalies that could adversely affect the evaluation of anti-icing fluids and ice-pellet contamination on the wing. Tests conducted with roughness and leading-edge flow disturbances helped to explain the aerodynamic impact of the anti-icing fluids and contamination. The stalling characteristics of the wing section with fluid and contamination appear to be driven at least partially by the effects of a secondary wave of fluid that forms near the leading edge as the wing is rotated in the simulated takeoff profile. These results have provided a much more complete understanding of the adverse aerodynamic effects of anti-icing fluids and ice-pellet contamination on this wing section. This is important since these results are used, in part, to develop the ice-pellet allowance times that are applicable to many different airplanes.

Broeren, Andy P.↗