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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

The effects of trailing vorticity on the flow through highly loaded cascades

This paper presents a procedure whereby three-dimensional inviscid flow through a highly loaded turbomachinery cascade of lifting lines can be treated by methods corresponding to classical aerodynamic theory. In contrast to earlier linearized (thin airfoil) three-dimensional theory, the present study allows analysis of the flow corresponding to the large turning and/or large pressure ratios induced by practical rotors or stators. For the sake of simplicity, the present paper is limited to incompressible flow through a highly loaded rectilinear cascade and to the design problem; i.e., given blade loading. Formulas are derived for both the mean and the three-dimensional components of the flow; in particular, the velocities at the blades induced by the trailing vorticity associated with nonuniform blade circulation are determined.

Mccune, J. E.↗
NASA NTRS1971

Preliminary wind-tunnel investigation of the effects of engine nacelles on a transport configuration with high lift drag ratios to a Mach number of 1.00

Wind tunnel tests to determine the effect of engine nacelles added to a low wing fuselage vertical tail configuration utilizing the NASA supercritical airfoil and a refined area ruled fuselage are discussed. The engine arrangement consisted of two aft fuselage, side mounted flow through nacelles and a solid body of revolution mounted above the fuselage in a manner similar to the Boeing 727. A preliminary analysis of the wind tunnel data shows that favorable interference drag can be obtained with the proper longitudinal locations of the nacelles, by canting the nacelle inlets, and by cusping the rearward region of the nacelle.

Flechner, S. G.↗
NASA NTRS1986

PROFILE user's guide

User information for program PROFILE, an aerodynamics design utility for refining, plotting, and tabulating airfoil profiles is provided. The theory and implementation details for two of the more complex options are also presented. These are the REFINE option, for smoothing curvature in selected regions while retaining or seeking some specified thickness ratio, and the OPTIMIZE option, which seeks a specified curvature distribution. REFINE uses linear techniques to manipulate ordinates via the central difference approximation to second derivatives, while OPTIMIZE works directly with curvature using nonlinear least squares techniques. Use of programs QPLOT and BPLOT is also described, since all of the plots provided by PROFILE (airfoil coordinates, curvature distributions) are achieved via the general purpose QPLOT utility. BPLOT illustrates (again, via QPLOT) the shape functions used by two of PROFILE's options. The programs were designed and implemented for the Applied Aerodynamics Branch at NASA Ames Research Center, Moffett Field, California, and written in FORTRAN and run on a VAX-11/780 under VMS.

Collins, L.↗
NASA NTRS1987

Local heat/mass transfer and pressure drop in a two-pass rib-roughened channel for turbine airfoil cooling

The heat transfer characteristics of turbulent air flow in a multipass channel were studied via the naphthalene sublimation technique. The naphthalene-coated test section, consisting of two straight, square channels joined by a 180 deg turn, resembled the internal cooling passages of gas turbine airfoils. The top and bottom surfaces of the test channel were roughened by rib turbulators. The rib height-to-hydraulic diameter ratio (e/D) were 0.063 and 0.094, and the rib pitch-to-height ratio (P/e) were 10 and 20. The local heat/mass transfer coefficients on the roughened top wall and on the smooth divider and side walls of the test channel were determined for three Reynolds numbers of 15, 30, and 60, thousand, and for three angles of attack (alpha) of 90, 60, and 45 deg. Results showed that the local Sherwood numbers on the ribbed walls were 1.5 to 6.5 times those for a fully developed flow in a smooth square duct. The average ribbed-wall Sherwood numbers were 2.5 to 3.5 times higher than the fully developed values, depending on the rib angle of attack and the Reynolds number. The results also indicated that, before the turn, the heat/mass transfer coefficients in the cases of alpha = 60 and 45 deg were higher than those in the case of alpha=90 deg. However, after the turn, the heat/mass transfer coefficients in the oblique-rib cases were lower than those in the transverse rib case. Correlations for the average Sherwood number ratios for individual channel surfaces and for the overall Sherwood number ratios are reported. Correlations for the fully developed friction factors and for the loss coefficients are also provided.

Han, J. C.↗
NASA NTRS1989

Effect of 90 degree flap on the aerodynamics of a two-element airfoil

The aerodynamic performance of a two-element airfoil with a 90-deg trailing edge flap was experimentally investigated. The 5 percent-chord long flap, significantly increased the lift of the baseline airfoil, throughout a wide range of angles of attack. The maximum lift coefficient of the flapped wing increased too, whereas the lift/drag ratio decreased.

Katz, J.↗
NASA NTRS1947

A theory of unstaggered airfoil cascades in compressible flow

By use of the methods of thin airfoil theory, which include effects of compressibility, relations are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord/gap ratios. For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low-speed flows. Choking cascades, a phenomenon of particular importance in compressor design, is considered.

Spurr, Robert A↗
NASA NTRS1981

Experimental study of the interaction between the wing of a subsonic aircraft and a nacelle of a high by-pass ratio engine

The oncoming of a new generation of subsonic transport aircraft (with supercritical wing and high by-pass ratio turbofans) led to an experimental study of wing nacelle jet pylon interference in transonic flow. To this end, a test set-up was developed at the ONERA S3Ch wind tunnel. The nacelle models represent a turbofan by means of two compressed air jets. The scale is 1/18.5. The nacelles are fixed on a thrust balance measuring afterbody thrust and discharge coefficients. The wing is located between the sidewalls of the test section. Pressures are measured through 456 holes located on 8 airfoils. Drag coefficient of the wing is obtained by wake survey. The following parameters can vary (1) wing/nacelle position; (2) upstream Mach number (from 0.3 to 0.8); (3) jet pressure ratio; (4) with/without pylon and (5) type of nacelle. Wing nacelle interference can be studied by means of total thrust drag analysis as a functon of the various parameters. The test set-up is described and examples of results are presented.

Levart, P.↗
DOE OSTI2021

Design of high-deflection foils MHK applications - CFD models - Helical turbines

The Ocean Renewable Power Company's (ORPC's) goal is to design, develop, and test hydrofoils with large deflections. The effects of the deflections on cross-flow turbine performance would be evaluated in order to inform design considerations for full-scale water turbines and other marine hydrokinetic devices. CFD models of helical model scale turbines tested at UNH OpenFOAM v1912 Tip Speed Ratio (TSR) = 3.00 Different strut configurations

16 TIDAL AND WAVE POWER↗
NASA NTRS1983

The performance of a circulation control airfoil at transonic speeds

A two-dimensional wind tunnel test has been performed on a small circulation control airfoil section equipped with trailing edge blowing. The model was mounted in the NASA-Ames 2 x 2 ft transonic wind tunnel and tested at speeds up to free-stream Mach equals 0.75 for a range of incidences. Jet pressure ratios up to 3 (relative to tunnel static conditions) were evaluated together with the effects of Reynolds number. Normal force and pitching moment coefficients were calculated from surface pressures using a Scanivalve pressure measuring system. Drag force coefficients were calculated from wake rake pressures. The results obtained indicated that this airfoil was capable of producing useful lift at high subsonic Mach numbers. Some changes in the stall characteristics were apparent at above free-stream Mach equals 0.4 and some dependence between lift augmentation and incidence was observed. There also appeared to be a significant Reynolds number effect on the airfoil drag performance.

Wood, N. J.↗
NASA NTRS1973

Low speed aerodynamic characteristics of a 17 percent thick airfoil section designed for general aviation applications

Wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-1). The results were compared with predictions based on a theoretical method for calculating the viscous flow about the airfoil. The tests were conducted over a Mach number range from 0.10 to 0.28. Reynolds numbers based on airfoil chord varied from 2.0 million to 20.0 million. Maximum section lift coefficients greater than 2.0 were obtained and section lift-drag ratio at a lift coefficient of 1.0 (climb condition) varied from about 65 to 85 as the Reynolds number increased from about 2.0 million to 6.0 million.

Mcghee, R. J.↗
NASA NTRS1942

High-speed Tests of a Ducted Body with Various Air-outlet Openings

Test of a ducted body with Internal flow were made in the 8-foot high-speed wind tunnel for the purpose of studying the effects on external drag and an critical speed of the addition of efficient inlet and outlet openings to a basic streamline shape. Drag tests of a 13.6- inch-diameter streamline body of fineness ratio 6.14 were made at Mach numbers ranging from 0.20 to 0.75. The model was centrally mounted on a 9-percent-thick airfoil and was designed to have an efficient airfoil-body juncture and a high critical speed. An air inlet at the nose and various outlets at the tail were added: drag and internal-flow data were obtained over the given speed range. The critical speed of the ducted bodies was found to be as high as that of the streamline body. The external - drag with air flow through the body did not exceed the drag of the basic streamline shape. No appreciable variation in the efficiency of the diffuser section of the internal duct occurred throughout the Mach number range of the tests.

Becker, John V.↗
NASA NTRS1947

A Theory of Unstaggered Airfoil Cascades in Compressible Flow

By use of the methods of thin airfoil theory, which include effects of compressibility, rela.tio^as are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord gap ratios, For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low speed flows. Choking of cascades, a phenomenon of particular importance in compressor design, is considered.

Spurr, Robert A.↗
NASA NTRS1945

Completed Tabulation in the United States of Tests of 24 Airfoils at High Mach Numbers (Derived from Interrupted Work at Guidonia, Italy in the 1.31- by 1.74-Foot High-Speed Tunnel)

Two-dimensional data were obtained in Mach range of from 0.40 to 0.94 and Reynolds Number range of (3.4 - 4.2) X 10 Degrees. Results indicate that thickness ratio is dominating shape parameter at high Mach numbers and that aerodynamic advantages are attainable by using thinnest possible sections. Effects of jet boundaries, Reynolds Number, and Data presented are free from jet-boundary and humidity effects.

AIR FLOW VELOCITY, HIGH - SUBSONIC↗
NASA NTRS1982

Advanced turbine study

Experiments to determine the available increase in turbine horsepower achieved by increasing turbine inlet temperature over a range of 1800 to 2600 R, while applying current gas turbine airfoil cling technology are discussed. Four cases of rocket turbine operating conditions were investigated. Two of the cases used O2/H2 propellant, one with a fuel flowrate of 160 pps, the other 80 pps. Two cases used O2/CH4 propellant, each having different fuel flowrates, pressure ratios, and inlet pressures. Film cooling was found to be the required scheme for these rocket turbine applications because of the high heat flux environments. Conventional convective or impingement cooling, used in jet engines, is inadequate in a rocket turbine environment because of the resulting high temperature gradients in the airfoil wall, causing high strains and low cyclic life. The hydrogen-rich turbine environment experienced a loss, or no gain, in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The effects of film cooling with regard to reduced flow available for turbine work, dilution of mainstream gas temperature and cooling reentry losses, offset the relatively low specific work capability of hydrogen when increasing turbine inlet temperature over the 1800 to 2600 R range. However, the methane-rich environment experienced an increase in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The results of a materials survey and heat transfer and durability analysis are discussed.

Source record↗
NASA NTRS1981

Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section

Wind-tunnel tests were conducted in the Langley low-turbulence pressure tunnel to evaluate the effects on performance of modifying a 17-percent-thick low-speed airfoil. The airfoil contour was altered to reduce the pitching-moment coefficient by increasing the forward loading and to increase the climb lift-drag ratio by decreasing the aft upper surface pressure gradient. The tests were conducted over a Mach number range from 0.07 to 0.32, a chord Reynolds number range 1.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle-of-attack range from about -10 deg to 20 deg.

Mcghee, R. J.↗
NASA NTRS1952

Airfoil profiles for minimum pressure drag at supersonic velocities -- general analysis with application to linearized supersonic flow

A theoretical investigation is made of the airfoil profile for minimum pressure drag at zero lift in supersonic flow. In the first part of the report a general method is developed for calculating the profile having the least pressure drag for a given auxiliary condition, such as a given structural requirement or a given thickness ratio. The various structural requirements considered include bending strength, bending stiffness, torsional strength, and torsional stiffness. No assumption is made regarding the trailing-edge thickness; the optimum value is determined in the calculations as a function of the base pressure. To illustrate the general method, the optimum airfoil, defined as the airfoil having minimum pressure drag for a given auxiliary condition, is calculated in a second part of the report using the equations of linearized supersonic flow.

Chapman, Dean R↗
NASA NTRS1987

Design and test of a natural laminar flow/large Reynolds number airfoil with a high design cruise lift coefficient

Research activity on an airfoil designed for a large airplane capable of very long endurance times at a low Mach number of 0.22 is examined. Airplane mission objectives and design optimization resulted in requirements for a very high design lift coefficient and a large amount of laminar flow at high Reynolds number to increase the lift/drag ratio and reduce the loiter lift coefficient. Natural laminar flow was selected instead of distributed mechanical suction for the measurement technique. A design lift coefficient of 1.5 was identified as the highest which could be achieved with a large extent of laminar flow. A single element airfoil was designed using an inverse boundary layer solution and inverse airfoil design computer codes to create an airfoil section that would achieve performance goals. The design process and results, including airfoil shape, pressure distributions, and aerodynamic characteristics are presented. A two dimensional wind tunnel model was constructed and tested in a NASA Low Turbulence Pressure Tunnel which enabled testing at full scale design Reynolds number. A comparison is made between theoretical and measured results to establish accuracy and quality of the airfoil design technique.

Kolesar, C. E.↗
NASA NTRS1972

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 5: Analysis and design of stages D and E

A conventional and a tandem bladed stage were designed for a comparative experimental evaluation in a 0.8 hub/tip ratio single-stage compressor. Based on a preliminary design study, a radially constant work input distribution was selected for the rotor designs. Velocity diagrams and blade leading and trailing edge angles selected for the conventional rotor and stator were used in the design of the tandem blading. The effects of axial velocity ratio and secondary flow on turning were included in the selection of blade leading and trailing edge angles. Design values of rotor tip velocity and stage pressure ratio were 757 ft/sec and 1.26, respectively.

Brent, J. A.↗