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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 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Evaluation of high Reynolds number flow in a 180 degree turn-around-duct

Mean and turbulent velocities were measured for the flow in a 180 degree turn-around-duct over a Reynolds number range from 600,000 to greater than 900,000. The measurements were made in water using a forward scattering laser velocimeter. A duct of 100 x 10 cm constant cross-section, with a mean radius of curvature (centerline) of 10 cm was employed for the study. The measurements are in agreement with previous studies in that the use of local bulk velocity to nondimensionalize the mean and turbulent velocities reduce the Reynolds number variations. The basic phenomenon of relaminarization along the inner surface at the exit of the turn are similar to the flow observed at low Reynolds numbers. The separation bubble region shows a systematic variation with Reynolds number, however the Reynolds number effect may be of second order in the calculation of the overall flow. Large tangential, radial, and lateral turbulent velocities are measured along the outer surface of the turn.

Sandborn, V. A.↗
NASA NTRS1992

Longitudinal and lateral-directional aerodynamic characteristics of a wing-cone configuration at Mach numbers from 2.5 to 4.5

A study was conducted to determine the longitudinal and lateral-directional aerodynamic characteristics of a generic wing-cone configuration at supersonic speeds. The fuselage has a 5 deg half-angle cone forebody, cylindrical midbody, and a 9 deg truncated cone afterbody. The delta wing (aspect ratio 1.0) had a 4 pct. thick diamond airfoil section. The tests were made in the Langley Unitary Plan Wind Tunnel at Mach numbers from 2.50 and 4.50. Angle of attack was varied from -4 to 28 degs, and angle of sideslip varied from -8 to 8 degs. Several configurations were studied to determine the effects of variations in wing longitudinal position, wing incidence, vertical tail configuration, canard shape, and nose bluntness. Typical effects of Reynolds number and Mach number on the longitudinal characteristics were observed. The incremental effects of the configuration variables were generally unaffected by Mach number. The directional stability characteristics of the large and small centerline mounted vertical tail configurations were significantly degraded with increasing angle of attack and Mach number.

Covell, Peter F.↗
NASA NTRS1992

Parametric investigation of single-expansion-ramp nozzles at Mach numbers from 0.60 to 1.20

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of varying six nozzle geometric parameters on the internal and aeropropulsive performance characteristics of single-expansion-ramp nozzles. This investigation was conducted at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.5 to 12, and angles of attack of 0 deg +/- 6 deg. Maximum aeropropulsive performance at a particular Mach number was highly dependent on the operating nozzle pressure ratio. For example, as the nozzle upper ramp length or angle increased, some nozzles had higher performance at a Mach number of 0.90 because of the nozzle design pressure was the same as the operating pressure ratio. Thus, selection of the various nozzle geometric parameters should be based on the mission requirements of the aircraft. A combination of large upper ramp and large lower flap boattail angles produced greater nozzle drag coefficients at Mach number greater than 0.80, primarily from shock-induced separation on the lower flap of the nozzle. A static conditions, the convergent nozzle had high and nearly constant values of resultant thrust ratio over the entire range of nozzle pressure ratios tested. However, these nozzles had much lower aeropropulsive performance than the convergent-divergent nozzle at Mach number greater than 0.60.

Capone, Francis J.↗
NASA NTRS1992

Reynolds number dependence of the freestream turbulence effects on turbulent boundary layers

Published experimental data on the influence of freestream turbulence on turbulent boundary layers were examined to determine the effect of Reynolds number on such influence. Two manifestations of the effect of low Reynolds numbers on the outer layer were observed: (1) the dependence of Clauser's shape parameter G on Reynolds number at very low Reynolds numbers and (2) the reduction in the wake component due to freestream turbulence undergoing a reversal in the Reynolds number dependence. These observations were used to modify Hancock's (1980) freestream turbulence parameter.

Bandyopadhyay, Promode R.↗
NASA NTRS1990

High-Reynolds-Number Test of a 5-Percent-Thick Low-Aspect-Ratio Semispan Wing in the Langley 0.3-Meter Transonic Cryogenic Tunnel: Wing Pressure Distributions

A high Reynolds number test of a 5 percent thick low aspect ratio semispan wing was conducted in the adaptive wall test section of the Langley 0.3 m Transonic Cryogenic Tunnel. The model tested had a planform and a NACA 64A-105 airfoil section that is similar to that of the pressure instrumented canard on the X-29 experimental aircraft. Chordwise pressure data for Mach numbers of 0.3, 0.7, and 0.9 were measured for an angle-of-attack range of -4 to 15 deg. The associated Reynolds numbers, based on the geometric mean chord, encompass most of the flight regime of the canard. This test was a free transition investigation. A summary of the wing pressures are presented without analysis as well as adapted test section top and bottom wall pressure signatures. However, the presented graphical data indicate Reynolds number dependent complex leading edge separation phenomena. This data set supplements the existing high Reynolds number database and are useful for computational codes comparison.

Chu, Julio↗
NASA NTRS1992

A finite Reynolds-number approach for the prediction of boundary-layer receptivity in localized regions

Previous theoretical work on the boundary layer receptivity problem has utilized large Reynolds number asymptotic theories, thus being limited to a narrow part of the frequency-Reynolds number domain. An alternative approach is presented for the prediction of localized instability generation which has a general applicability, and also accounts for finite Reynolds number effects. This approach is illustrated for the case of Tollmien-Schlichting wave generation in a Blasius boundary layer due to the interaction of a free stream acoustic wave with a region of short scale variation in the surface boundary condition. The specific types of wall inhomogeneities studied are: regions of short scale variations in wall suction, wall admittance, and wall geometry (roughness). Extensive comparison is made between the results of the finite Reynolds number approach and previous asymptotic predictions, which also suggests an alternative way of using the latter at Reynolds numbers of interest in practice.

Choudhari, Meelan↗
NASA NTRS1993

Direct simulations of premixed turbulent flames with nonunity Lewis numbers

A principal effect of turbulence on premixed flames in the flamelet regime is to wrinkle the flame fronts. For nonunity Lewis numbers, Le is not equal to 1, the local flame structure is altered in curved regions. This effect is examined using direct numerical simulations of 3D isotropic turbulence with constant density, single-step Arrhenius kinetics chemistry. Simulations of Lewis numbers 0.8, 1.0, and 1.2 are compared. At the local level, curvature effects dominated changes to the flame structure while strain effects were insignificant. A strong Lewis-number-dependent correlation was found between surface curvature and the local flame speed. The correlation was positive for Le less than 1 and negative for Le greater than 1. At the global level, strain-related effects were more significant than curvature effects. The turbulent flame speed changed significantly with Lewis number, increasing as Le decreased. This was found to be due to strain effects that have a nonzero mean over the flame surface, rather than to curvature effects that have a nearly zero mean. The mean product temperature was also found to vary with Lewis number, being higher for Le greater than 1 and lower for Le less than 1.

Rutland, C. J.↗
NASA NTRS1993

Prediction of stall and post-stall behavior of airfoils at low and high Reynolds numbers

An interactive boundary-layer method, together with the e(super n)-approach to the calculation of transition, has been used to predict the stall and post-stall behavior of airfoils at low and high Reynolds numbers. The turbulence model is based on the Cebeci-Smith algebraic eddy-viscosity formulation with improvements for strong pressure gradient effects and transitional flows at low Reynolds numbers. Comparison of calculated results for incompressible flows indicate good agreement with experiment for a wide range of Reynolds numbers. Preliminary calculations for low Mach number flows with this interactive method with compressibility corrections to the panel method indicate that, at a Mach number of 0.3, the compressibility effect on (C sub Q)max is not negligible.

Cebeci, T.↗
NASA NTRS1993

Low-Reynolds-number k-epsilon model for unsteady turbulent boundary-layer flows

An assessment of the near-wall and low-Reynolds-number functions used in low-Reynolds-number k-epsilon models suggests that they are not suitable for the near-wall region of unsteady turbulent boundary layers, where the flow is characterized by rapid changes in phase. An improved low-Reynolds-number k-epsilon model is developed in this paper. The near-wall and low-Reynolds-number functions in this model are formulated as functions of the local turbulent Reynolds numbers instead of the inner variable y(+). The present model also has the correct asymptotic behavior in the near-wall region. The turbulence model has been incorporated in an unsteady boundary-layer code and validated for unsteady turbulent boundary layers with and without adverse pressure gradients. The predictions agree well with the experimental data and the theoretical analysis. For the cases tested, the present model correctly predicts the unsteady near-wall flow and the unsteady shin friction at various frequencies.

Fan, Sixin↗
NASA NTRS1952

Effect of Mach Number on the Flow and Application of Compressibility Corrections in a Two-Dimensional Subsonic-Transonic Compressor Cascade Having Varied Porous-Wall Suction at the Blade Tips

A cascade of 65-(12)10 compressor blades was tested at one geometric setting over a range of inlet Mach number from 0.12 to 0.89. Two groups of data are presented and compared: the first from the cascade operating conventionally with no boundary-layer control, and the second with the boundary layer controlled by a combination of upstream slot suction and porous-wall suction at the blade tips. A criterion for two-dimensionality was used to specify the degree of boundary-layer control by suction to be applied. The data are presented and an analysis is made to show the effect of Mach number on turning angle, blade wake, pressure distribution about the blade profile and static-pressure rise. The influence of boundary-layer control on these parameters as well as on the secondary losses is illustrated. A system of correlating the measured static-pressure rise through the cascade with the theoretical isentropic values is presented which gives good agreement with the data. The pressure distribution about the blade profile for an inlet Mach number of 0.21 is corrected with the Prandtl-Glauert, Karman-Tsien, and vector-mean velocity - contraction coefficient compressibility correction factors to inlet Mach numbers of 0.6 and 0.7. The resulting curves are compared with the experimental pressure distributions for inlet Mach numbers of 0.6 and 0.7 so that the validity of applying the three corrections can be evaluated.

Briggs, William B.↗
NASA NTRS1951

Analytical investigation of ram-jet-engine performance in flight Mach number range from 3 to 7

An analytical investigation was made of the performance of isolated ram-jet engines in the flight Mach number range from 3 to 7 for two types of diffuser, a high-efficiency diffuser, and a normal-shock diffuser. The fuel was assumed to be a hydrocarbon similar to gasoline. The conclusions reached are: (1) a design altitude of about 100,000 feet is desirable for a high-efficiency high Mach number ram jet on the basis of engine construction and performance; and (2) although greater thrust could be obtained with other fuels, gasoline provides sufficient energy release for maximum engine efficiency in the flight Mach number range investigated. The maximum engine efficiency calculated was 0.47, which occurred at a Mach number of 5. At a Mach number of 7, the maximum propulsive-thrust coefficient was 0.57.

ENGINES, RAM JET↗
NASA NTRS1953

Preliminary Drag and Heat-transfer Data Obtained from Air-launched Cone-cylinder Test Vehicle over Mach Number Range from 1.5 to 5.18

An air-launched cone-cylinder test vehicle designed to obtain data at Mach numbers above 4.0 was rocket boosted from a release Mach number of 5.18. The vehicle was launched at an altitude of 35,000 feet and reached peak velocity of 5150 feet per second at 28,500 feet. The total-drag coefficient (based on maximum cross-sectional area) decreased gradually from 0.31 at a Mach number of 1.75 to 0.145 at a Mach number of 5.18, while the Reynold's number (based on body length) increased from 31 x 10 to the 6th power to 107 x 10 to the 6th power. The skin friction coefficients, in general, were slightly lower than Van Driest's theoretical values for similar wall-temperature conditions. Convective heat-transfer coefficients were obtained from a single skin-thermocouple measurement. The maximum wall temperature recorded was 1240 degrees r.

Messing, Wesley E↗
NASA NTRS1957

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

A full- scale rocket-powered model of a cruciform canard missile configuration with a low- aspect - ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed- control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift - curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift- .curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta= -0.3 deg . The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic- center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number.The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Brown, Clarence A , Jr↗
NASA NTRS1958

Investigation of an underslung half-cone inlet with compression surface mounted outboard from fuselage at Mach numbers of 1.5, 1.8, and 2.0

An investigation was conducted to determine the performance of an underslung half-cone inlet mounted on a missile forebody model with the compression surface outboard from the fuselage. The inlet was designed for shock-on-lip operation at Mach number 2.0 with 25 degree half-angle spike. The cowling was attached to the fuselage through the boundary-layer plow and served as part of the fuselage boundary-layer diverter system. The performance of the half-cone inlet was compared with that of a scoop-type inlet and a normal-wedge inlet on a maximum-thrust-minus-drag basis. The increase in pressure recovery obtained with the half-cone inlet over that obtained with the reference inlets offset the slightly higher drags observed over the Mach number range for the half-cone so that the performance of this configuration was equal to that of the other inlets at Mach number 2.0 and was slightly superior at the lower Mach numbers. For a particular configuration, a peak pressure recovery of 0.879 was obtained at Mach number 2.0, zero angle of attack, and 4-percent throat bleed; the subcritical stability was 16 percent. Use of a fuselage-mounted boundary-layer splitter plate ahead of the inlet did not improve the stability. Subcritical distortion values were below 10 percent for all configurations. (author)

Yeager, Richard A↗
NASA NTRS1946

Variation with Mach Number of Static and Total Pressures Through Various Screens

Tests were conducted in the Langley 24-inch highspeed tunnel to ascertain the static-pressure and total-pressure losses through screens ranging in mesh from 3 to 12 wires per inch and in wire diameter from 0.023 to 0.041 inch. Data were obtained from a Mach number of approximately 0.20 up to the maximum (choking) Mach number obtainable for each screen. The results of this investigation indicate that the pressure losses increase with increasing Mach number until the choking Mach number, which can be computed, is reached. Since choking imposes a restriction on the mass rate of flow and maximum losses are incurred at this condition, great care must be taken in selecting the screen mesh and wire dimmeter for an installation so that the choking Mach number is

Adler, Alfred A↗
NASA NTRS1950

The Effect of the Inlet Mach Number and Inlet-boundary-layer Thickness on the Performance of a 23 Degree Conical-diffuser-tail-pipe Combination

An investigation was conducted to determine the effect of the inlet Mach number and entrance-boundary-layer thickness on the performance of a 23 degree 21-inch conical-diffuser - tail-pipe combination with a 2:1 area ratio. The air flows used in this investigation covered an inlet Mach number range from 0.17 to 0.89 and corresponding Reynolds numbers of 1,700,000 to 7,070,000. Results are reported for two inlet-boundary-layer thicknesses. Over the entire range of flows, the mean value of the inlet displacement thickness is about 0.034 inch for the thinner inlet boundary layer and about 0.170 inch for the case of the thicker inlet boundary layer. The performance of the diffuser - tail-pipe combination is presented together with examples of longitudinal static-pressure distribution and the results of boundary-layer pressure surveys made at six points along the diffuser wall. The results indicated a progressive diminution of the static-pressure recovery and a steady increase in the total-pressure losses as the inlet Mach number was increased for both inlet-boundary-layer thicknesses. The ratio of actual static-pressure rise to that theoretically possible was much less and the total-pressure losses were greater for the case of the thicker inlet boundary layer throughout the speed range investigated. With the thinner inlet boundary layer, flow separation occurred at the diffuser exit at all inlet Mach numbers.Unseparated flow alternating with separated flow was observed near the inlet at the higher velocities. For the case of the thicker inlet boundary layer, the origin of the separated region occurred in the vicinity of the inlet-duct-diffuser junction section at all Mach numbers.

Persh, Jerome↗
NASA NTRS1955

Aerodynamic Loads on an External Store Adjacent to a 45 Degree Sweptback Wing at Mach Numbers from 0.70 to 1.96, Including an Evaluation of Techniques Used

Aerodynamic forces and moments have been obtained in the Langley 9- by 12-inch blowdown tunnel on an external store and on a 45 degree swept-back wing-body combination measured separately at Mach numbers from 0.70 to 1.96. The wing was cantilevered and had an aspect ratio of 4.0; the store was independently sting-mounted and had a Douglas Aircraft Co. (DAC) store shape. The angle of attack range was from -3 degrees to 12 degrees and the Reynolds number (based on wing mean aerodynamic chord) varied from 1.2 x10(6) to 1.7 x 10(6). Wing-body transonic forces and moments have been compared with data of a geometrically similar full-scale model tested in the Langley 16-foot and 8-foot transonic tunnels in order to aid in the evaluation of transonic-tunnel interference. The principal effect of the store, for the position tested, was that of delaying the wing-fuselage pitch-up tendency to higher angles of attack at Mach numbers from 0.70 to 0.90 in a manner similar to that of a wing chord extension. The most critical loading condition on the store was that due to side force, not only because the loads were of large magnitude but also because they were in the direction of least structural strength of the supporting pylon. These side loads were greatest at high angles of attack in the supersonic speed range. Removal of the supporting pylon (or increasing the gap between the store and wing) reduced the values of the variation of side-force coefficientwith angle of attack by about 50 percent at all test Mach numbers, indicating that important reductions in store side force may be realized by proper design or location of the necessary supporting pylon. A change of the store skew angle (nose inboard) was found to relieve the excessive store side loads throughout the Mach number range. It was also determined that the relative position of the fuselage nose to the store can appreciably affect the store side forces at supersonic speeds.

Guy, Lawrence D↗
NASA NTRS1993

High Reynolds Number Configuration Development of a High-Lift Airfoil

An experimental program has been conducted to assess performance of a transport multielement airfoil at flight Reynolds numbers. The studies were performed at chord Reynolds numbers as high as 16 million in the NASA Langley Low Turbulence Pressure Tunnel. Sidewall boundary-layer control to enforce flow two dimensionality was provided via an endplate suction system. The basic airfoil was an 11.55 percent thick supercritical airfoil representative of the stall critical station of a new-generation transport aircraft wing. The multielement airfoil was configured as a three-element airfoil with slat and flap chord ratios of 14.48 percent and 30 percent respectively. Testing focused on the development of landing configurations with high maximum lift capability and the assessment of Reynolds and Mach number effects. Also assessed were high-lift performance effects due to devices such as drooped spoilers and trailing-edge wedges. The present experimental studies revealed significant effects on high-lift airfoil performance due to Reynolds and Mach number variations and favorable lift increments at approach angles of attack due to the use of drooped spoilers or trailing-edge wedges. However, no substantial improvements in maximum lift capability were identified. A recently developed high performance single-segment flap was also tested and results indicated considerable improvements in lift and drag performance over existing airfoils. Additionally, it was found that this new flap shape at its optimum rigging was less sensitive to Reynolds number variations than previous designs.

Walter O Valarezo↗