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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 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.↗
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.↗
NASA NTRS1992

Fully integrated aerodynamic/dynamic optimization of helicopter rotor blades

A fully integrated aerodynamic/dynamic optimization procedure is described for helicopter rotor blades. The procedure combines performance and dynamic analyses with a general purpose optimizer. The procedure minimizes a linear combination of power required (in hover, forward flight, and maneuver) and vibratory hub shear. The design variables include pretwist, taper initiation, taper ratio, root chord, blade stiffnesses, tuning masses, and tuning mass locations. Aerodynamic constraints consist of limits on power required in hover, forward flight and maneuvers; airfoil section stall; drag divergence Mach number; minimum tip chord; and trim. Dynamic constraints are on frequencies, minimum autorotational inertia, and maximum blade weight. The procedure is demonstrated for two cases. In the first case, the objective function involves power required (in hover, forward flight and maneuver) and dynamics. The second case involves only hover power and dynamics. The designs from the integrated procedure are compared with designs from a sequential optimization approach in which the blade is first optimized for performance and then for dynamics. In both cases, the integrated approach is superior.

Walsh, Joanne L.↗
NASA NTRS1992

Fully integrated aerodynamic/dynamic optimization of helicopter rotor blades

This paper describes a fully integrated aerodynamic/dynamic optimization procedure for helicopter rotor blades. The procedure combines performance and dynamics analyses with a general purpose optimizer. The procedure minimizes a linear combination of power required (in hover, forward flight, and maneuver) and vibratory hub shear. The design variables include pretwist, taper initiation, taper ratio, root chord, blade stiffnesses, tuning masses, and tuning mass locations. Aerodynamic constraints consist of limits on power required in hover, forward flight and maneuver; airfoil section stall; drag divergence Mach number; minimum tip chord; and trim. Dynamic constraints are on frequencies, minimum autorotational inertia, and maximum blade weight. The procedure is demonstrated for two cases. In the first case the objective function involves power required (in hover, forward flight, and maneuver) and dynamics. The second case involves only hover power and dynamics. The designs from the integrated procedure are compared with designs from a sequential optimization approach in which the blade is first optimized for performance and then for dynamics. In both cases, the integrated approach is superior.

Walsh, Joanne L.↗
NASA NTRS1944

The Wall Interference of a Wind Tunnel of Elliptic Cross Section

The wall interference is obtained for a wind tunnel of elliptic section for the two cases of closed and open working sections. The approximate and exact methods used gave results in practically good agreement. Corresponding to the result given by Glauert for the case of the closed rectangular section, the interference is found to be a minimum for a ratio of minor to major axis of 1:square root of 6 This, however, is true only for the case where the span of the airfoil is small in comparison with the width of the tunnel. For a longer airfoil the favorable ellipse is flatter. In the case of the open working section the circular shape gives the minimum interference.

Tani, Itiro↗
NASA NTRS1953

Summary of Rocket-Model Tests at Zero Lift of the Northrop MX-775B Missile Configuration from Mach Numbers of 0.9 to 1.8

Flight tests were conducted between Mach numbers of 0.9 and 1.8 over a Reynolds number range of 9(exp 6) to 30(exp 6) to determine the zero-lift drag and some rolling-effectiveness characteristics of the Northrop MX -775B missile with small and large body. The MX-775B is a proposed long range, supersonic, ground-to-ground missile having an arrow wing with 67.5 degree leading-edge sweep, 15 deg trailing-edge sweep, and a modified NACA 0004 airfoil section. The configuration has no horizontal tail but has wing trailing-edge elevons which serve a dual purpose as elevators and ailerons. The ratio of body frontal area to wing plan-form area is 0.0127 for the small-body configuration and 0.0330 for the large-body configuration. Five 1/4-scale models were flown permitting determination of the drag coefficient for the basic small-body configuration, the incremental drag due to the large body, the incremental drag resulting from a blunt wing trailing edge, the wing-plus-interference drag, and some rolling-effectiveness data. Results indicated that the MX-775B has low supersonic zero-lift drag, the maximum zero-lift drag coefficients being respectively 0.0125 and 0.0155 at a Mach number of M = 1803 for the small- and large-body configurations. The effect of a blunt wing trailing edge, obtained by cutting off 10 percent of the wing chord, was to increase the zero-lift drag by 13 to 21 percent. Wing-plus-interference drag accounted for 78 percent of the total drag at M = 0.9 and 70 percent at M = 195 for the small-body configuration. The ailerons produced positive rolling effectiveness for the wing stiffness of the test models and the dynamic pressures of the test.

Arbic, Richard G.↗
NASA NTRS2012

Evaluation of Icing Scaling on Swept NACA 0012 Airfoil Models

Icing scaling tests in the NASA Glenn Icing Research Tunnel (IRT) were performed on swept wing models using existing recommended scaling methods that were originally developed for straight wing. Some needed modifications on the stagnation-point local collection efficiency (i.e., beta(sub 0) calculation and the corresponding convective heat transfer coefficient for swept NACA 0012 airfoil models have been studied and reported in 2009, and the correlations will be used in the current study. The reference tests used a 91.4-cm chord, 152.4-cm span, adjustable sweep airfoil model of NACA 0012 profile at velocities of 100 and 150 knot and MVD of 44 and 93 mm. Scale-to-reference model size ratio was 1:2.4. All tests were conducted at 0deg angle of attack (AoA) and 45deg sweep angle. Ice shape comparison results were presented for stagnation-point freezing fractions in the range of 0.4 to 1.0. Preliminary results showed that good scaling was achieved for the conditions test by using the modified scaling methods developed for swept wing icing.

Tsao, Jen-Ching↗
NASA NTRS2015

Rotor with Flattened Exit Pressure Profile

A rotor blade comprises an airfoil extending radially from a root section to a tip section and axially from a leading edge to a trailing edge, the leading and trailing edges defining a curvature therebetween. The curvature determines a relative exit angle at a relative span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height. In operation of the rotor blade, the relative exit angle determines a substantially flat exit pressure ratio profile for relative span heights from 75% to 95%, wherein the exit pressure ratio profile is constant within a tolerance of 10% of a maximum value of the exit pressure ratio profile.

Baltas, Constantine↗