Search NASA⌕ Search

SEARCH · Search NASA

Results for “numbers”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

Noonan, K. W.↗
NASA NTRS1978

Simulated flight effects on noise characteristics of a fan inlet with high throat Mach number

An anechoic wind tunnel experiment was conducted to determine the effects of simulated flight on the noise characteristics of a high throat Mach number fan inlet. Comparisons were made with the performance of a conventional low throat Mach number inlet with the same 50.8 cm fan noise source. Simulated forward velocity of 41 m/sec reduced perceived noise levels for both inlets, the largest effect being more than 3 db for the high throat Mach number inlet. The high throat Mach number inlet was as much as 7.5 db quieter than the low throat Mach number inlet with tunnel airflow and about 6 db quieter without tunnel airflow. Effects of inlet flow angles up to 30 deg were seemingly irregular and difficult to characterize because of the complex flow fields and generally small noise variations. Some modifications of tones and directivity at blade passage harmonics resulting from inlet flow angle variation were noted.

Wesoky, H. L.↗
NASA NTRS1979

Investigation of the asymmetric aerodynamic characteristics of cylindrical bodies of revolution with variations in nose geometry and rotational orientation at angles of attack to 58 degrees and Mach numbers to 2

Wind-tunnel tests were conducted to investigate the side forces and yawing moments that can occur at high angles of attack and zero sideslip for cylindrical bodies of revolution. Two bodies having several tangent ogive forebodies with fineness ratios of 0.5, 1.5, 2.5, and 3.5 were tested. The forebodies with fineness ratios of 2.5 and 3.5 had several bluntnesses. The cylindrical afterbodies had fineness ratios of 7 and 13. The model components - tip, forebody, and afterbody - were tested in various rotational positions about their axes of symmetry. Most of the tests were conducted at a Mach number of 0.25, a Reynolds number of 0.32 x 10 to the 6th power, and with the afterbody that had a fineness ratio of 7 and with selected forebodies. The effect of Mach number was determined with the afterbody that had a fineness ratio of 13 and with selected forebodies at mach numbers from 0.25 to 2 at Reynolds number = 0.32 X 10 to the 6th power. Maximum angle of attack was 58 deg.

Kruse, R. L.↗
NASA NTRS1979

Low-Reynolds-number turbulent boundary layers

A test facility operating at substantially reduced pressures is described and shown to be suitable when used with an open-circuit turbulent boundary layer wind tunnel. The facility is capable of producing flows with momentum Reynolds number in the range 200-10,000 by varying the ambient pressure. At low-Reynolds-number flows the pressure is low enough to allow the tunnel to operate with large values of freestream speed. The tunnel's reliability is verified by examining flows where momentum Reynolds number is greater than 1200. The results skin friction coefficients, wake parameters, and basic velocity profiles - consistently agreed with well-established data. For momentum Reynolds number from 600 through 1200 the resulting data indicate that the 'universal velocity profile' is accurate and correctly describes the mean flow properties. For Reynolds number from 425 through 600, the viscous sublayer in this range can no longer maintain its characteristic size.

White, B. R.↗
NASA NTRS1980

Strouhal number influence on flight effects on jet noise radiated from convecting quadrupoles

The paper reports a complementary extension of previous work to include the high-frequency features reflected in the discussion of the higher Strouhal number influence on flight effects. It is found that, in addition to the usual features of flight effects on noise from ordinary flows, the high Strouhal number flows exhibit some more interesting features which are uniquely characteristic to them. The additional features are as follows: (1) Flight effects are more favorable to hot jets than to cold jets; (2) the higher the Strouhal number of the jet flow, the lesser the forward arc amplification due to flight; (3) as the Strouhal number increases, the peak amplification angle in the forward quadrant and the peak suppression angle in the aft quadrant move toward 90 deg and get closer, thus reducing the amplification exposure to a constricted angular region; (4) the silence zone is disturbed and displaced from its normal position parallel to the jet flow to give rise to multiple crossings of flight curves with the static line; and (5) the occurrence of multiple crossings is a strange phenomenon solely characteristic of high Strouhal number with high subcritical jet flows in flight.

Dash, R.↗
NASA NTRS1981

Comments on Reynolds number effects and the role of NTF in the development of air vehicles

The National Transonic Facility (NTF) capability to match the full scale Reynolds numbers of all but the largest airplanes is discussed. Conversion factors to enable calculation of Sl-unit equivalents for all U.S. units are listed. Using data from several facilities, analytic methods, and flight test data, a competetive aircraft in the relatively low Reynolds number was developed. The NTF offers the capability to obtain data at full scale Reynolds numbers in the cruise condition for most of the products, and will be much closer than previous tunnels to full scale Reynolds number for the operating envelopes. It is primarily on the operating envelope that Reynolds number effects are most important and least predictable.

Nagel, A. L.↗
NASA NTRS1981

Heat transfer on a flat plate in continuum to rarefied hypersonic flows at Mach numbers of 19.2 and 25.4

Local heat transfer rates were measured on a flat steel plate (10 in. wide and 16 in. long) with sharp and blunt leading edges (0.001 and 0.010 in.) in the transition from the strong interaction boundary layer regime with no slip at the surface to the free molecule regime. The tests were conducted in a combustion driven hypersonic shock tunnel, with the nominal free stream Mach numbers of 19.2 and 25.4, and a reflected stagnation temperature of approximately 2340 R. The twelve heat transfer gages were made of platinum sputtered on a Pyrex backing to a thickness of approximately 350 A, insulated by a silicone dioxide film. For both Mach numbers the heat transfer data agreed reasonably well with the strong interaction prediction of Li and Nagamatsu (1953, 1955) for unit Reynolds numbers greater than approximately 100,000 and leading edge Knudsen numbers less than approximately 4. At lower density conditions the rarefied flow effects began to dominate the flow phenomena near the leading edge region of the sharp flat plate. A systematic reduction in the heat transfer rate close to the leading edge was observed for both Mach number tests as the leading edge density was reduced and the mean free path was increased.

Nagamatsu, H. T.↗
NASA NTRS1981

Low Reynolds number airfoil survey, volume 1

The differences in flow behavior two dimensional airfoils in the critical chordlength Reynolds number compared with lower and higher Reynolds number are discussed. The large laminar separation bubble is discussed in view of its important influence on critical Reynolds number airfoil behavior. The shortcomings of application of theoretical boundary layer computations which are successful at higher Reynolds numbers to the critical regime are discussed. The large variation in experimental aerodynamic characteristic measurement due to small changes in ambient turbulence, vibration, and sound level is illustrated. The difficulties in obtaining accurate detailed measurements in free flight and dramatic performance improvements at critical Reynolds number, achieved with various types of boundary layer tripping devices are discussed.

Carmichael, B. H.↗
NASA NTRS1982

In-flight transition measurement on a 10 deg cone at Mach numbers from 0.5 to 2.0

Boundary layer transition measurements were made in flight on a 10 deg transition cone tested previously in 23 wind tunnels. The cone was mounted on the nose of an F-15 aircraft and flown at Mach numbers room 0.5 to 2.0 and altitudes from 1500 meters (5000 feet) to 15,000 meters (50,000 feet), overlapping the Mach number/Reynolds number envelope of the wind tunnel tests. Transition was detected using a traversing pitot probe in contact with the surface. Data were obtained near zero cone incidence and adiabatic wall temperature. Transition Reynolds number was found to be a function of Mach number and of the ratio of wall temperature to adiabatic all temperature. Microphones mounted flush with the cone surface measured free-stream disturbances imposed on the laminar boundary layer and identified Tollmien-Schlichting waves as the probable cause of transition. Transition Reynolds number also correlated with the disturbance levels as measured by the cone surface microphones under a laminar boundary layer as well as the free-stream impact.

Fisher, D. F.↗
NASA NTRS1982

A review of Reynolds number studies conducted in the Langley 0.3-m Transonic Cryogenic Tunnel

Initial proof-of-concept results obtained in a pilot transonic tunnel, led to a U.S. decision to apply the cryogenic tunnel concept to the National Transonic Facility (NTF). The pilot tunnel was reclassified as a 'permanent' facility and renamed the 0.3-m Transonic Cryogenic Tunnel (TCT). The 0.3-m TCT, with its ability to operate at temperatures ranging from about 78 K to 340 K, combined with a 6 atmospheres pressure capability, provides extremely high test Reynolds numbers at relatively low model loading. In addition, the ability to vary pressure and temperature independently of Mach number provides a unique capability to conduct studies at a constant dynamic pressure or at a constant Reynolds number with just one model. Attention is given to the simulation of flight conditions, the cryogenic-pressure tunnel concept, the characteristics of the 0.3-m TCT, Reynolds number studies in the three-dimensional test section including a Space Shuttle study, and Reynolds number studies in the two-dimensional test section.

Ray, E. J.↗
NASA NTRS1982

The applicability of Lyapunov characteristic numbers in the study of the stability of satellite orbits

It is pointed out that the Lyapunov Characteristic Numbers constitute a new tool for determining stability of trajectories of dynamical systems, or, even more generally, of solutions of systems of ordinary differential equations. In contrast with the characteristic exponents, which apply only to periodic solutions, the Lyapunov Characteristic Numbers apply to arbitrary nonperiodic solutions as well. A description is presented of the numerical experiments which have been made in order to investigate the practical value of the Lyapunov Characteristic Number and the Kolmogorov Entropy for the purpose of estimating the stability of trajectories and/or numerical integration methods in celestial mechanics. It is found that the Lyapunov Characteristic Numbers are extremely useful for the classification of the solutions of nonintegrable dynamical systems, especially in order to distinguish between quasi-periodic and chaotic solutions. However, the Lyapunov Characteristics Numbers do not appear to be useful for the purpose of evaluating numerical integration methods.

Broucke, R.↗
NASA NTRS1983

Effect of low Reynolds number turbulence amplification on the Galileo probe flowfield

The amplification of turbulence at low Reynolds numbers is analyzed as it affects the Jupiter-entry flowfield, surface-heating rate, and mass-loss rate of the 335-kg Galileo probe now being designed. The constant k(2) in the Clauser-Klebanoff outer law of the two-layer algebraic eddy-viscosity model of Cebeci (1970), as used in earlier models of the Galileo flowfield (Moss and Simmonds, 1982), is modified to increase at low Reynolds numbers, as found experimentally by Varner and Adams (1980). Calculations were performed for peak heating conditions (at 49.13 sec of entry), using a turbulent Prandtl number of 0.9 and turbulent Lewis numbers (LeT) of 0.8, 1.0, and 1.2. The low-Reynolds-number effect is found to produce mass-loss-ratio increases of from 4 to 50 percent, while an LeT of 1.2 produced an increase of from 4 to 10 percent as compared to an LeT of 1.0. While these findings are based on experimental data obtained under conditions somewhat different from those probably present on Jupiter, their importance for a conservative probe design is stressed.

Gupta, R. N.↗
NASA NTRS1983

Reynolds number tests of an NPL 9510 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

An investigation of the NPL 9510 airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel over the following ranges of test conditions: Mach number of 0.35 to 0.82, total temperature of 94 K to 300 K, total pressure of 1.20 to 5.81 atm, Reynolds number based on airfoil chord of 1.34 x 10 to the 6th power to 48.23 x 10 to the 6th power, and angle of attack of 0 deg to 6 deg. The drag creep previously reported by the British National Physics Laboratory at low Reynolds numbers was also found to be present at high Reynolds numbers; the section drag coefficient continued to decrease even at the highest Reynolds number tested. Tests made close to free-stream saturation did not produce altered aerodynamic coefficients due to condensation effects.

Jenkins, R. V.↗
NASA NTRS1982

High Reynolds number tests of a Boeing BAC I airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind tunnel investigation of an advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents the first in a series of NASA/U.X. industry two dimensional airfoil studies to be completed in the Advanced Technology Airfoil Test program. Test temperature was varied from ambient to about 100 K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from about 0.40 to 0.80. These variables provided a Reynolds number (based on airfoil chord) range from about .0000044 to .00005. This investigation was specifically designed to: (1) test a Boeing advanced airfoil from low to flight-equivalent Reynolds numbers; (2) provide the industry participant (Boeing) with experience in cryogenic wind-tunnel model design and testing techniques; and (3) demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. All the objectives of the cooperative test were met. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics of the airfoil. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

Johnson, W. G., Jr.↗
NASA NTRS1984

The influence of laminar separation and transition on low Reynolds number airfoil hysteresis

An experimental study of the Lissaman 7769 and Miley MO6-13-128 airfoils at low chord Reynolds numbers is presented. Although both airfoils perform well near their design Reynolds number of about 600,000, they each produce a different type of hysteresis loop in the lift and drag forces when operated below chord Reynolds numbers of 300,000. The type of hysteresis loop was found to depend upon the relative location of laminar separation and transition. The influence of disturbance environment and experimental procedure on the low Reynolds number airfoil boundary layer behavior is also presented. The use of potential flow solutions to help predict how a given airfoil will behave at low Reynolds numbers is also discussed.

Mueller, T. J.↗
NASA NTRS1985

High Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.60 to 0.80. These variables provided a Reynolds number range from 4,400,000 to 40,000,000 based on a 15.24-cm (6.0-in.) airfoil chord. This investigation was designed to test a NASA advanced-technology airfoil from low to flight-equivalent Reynolds numbers, provide experience in cryogenic wind tunnel model design and testing techniques, and demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

Johnson, W. G., Jr.↗
NASA NTRS1985

Coherent large-scale structures in high Reynolds number supersonic jets

The flow structure of a 50.8 mm (2 in) diameter jet operated at a full expanded Mach number of 1.37, with Reynolds numbers in the range 1.7 to 2.35 million, was examined for the first 20 jet diameters. To facilitate the study of the large scale structure, and determine any coherence, a discrete tone acoustic excitation method was used. Phase locked flow visualization as well as laser velocimeter quantitative measurements were made. The main conclusions derived from this study are: (1) large scale coherent like turbulence structures do exist in large Reynolds number supersonic jets, and they prevail even beyond the potential core; (2) the most preferential Strouhal number for these structures is in the vicinity of 0.4; and (3) quantitatively, the peak amplitudes of these structures are rather low, and are about 1% of the jet exit velocity. Finally, since a number of unique problems related to LV measurements in supersonic jets were encountered, a summary of these problems and lessons learned therefrom are also reported.

Lepicovsky, J.↗
NASA NTRS1986

Experimental determination of the laminar separation bubble characteristics on an airfoil at low Reynolds numbers

An experimental investigation was conducted in order to document the structure and behavior of laminar separation bubbles at low Reynolds numbers. Data of this type is necessary if the currently insufficient analytical and numerical models are to be improved. The laminar separation bubble which forms on a NACA 66(3)-018 airfoil model was surveyed at chord Reynolds numbers ranging from 50,000 to 200,000 at angles of attack from 8 to 12 degrees. The effects of the various testing conditions on the separation bubble were isolated, and the data was analyzed in relation to existing separation bubble correlations in order to test their low Reynolds number applicability. This analysis indicated that the chord Reynolds number and the disturbance environment strongly influence the experimental pressure distributions. These effects must be included in any analytic prediction technique applied to the low Reynolds number flight regime.

Omeara, M. M.↗