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  • Animals  (417)
  • AERODYNAMICS  (342)
  • 1980-1984  (759)
  • 1925-1929
  • 1980  (759)
  • 1
    Publication Date: 2013-08-31
    Description: These basic characteristics of critical wings included wing area, aspect ratio, average thickness, and sweep as well as practical constraints on the planform and thickness near the wing root to allow for the landing gear. Within these constraints, a large matrix of wing designs was studied with spanwise variations in the types of airfoils and distribution of lift as well as some small planform changes. The criteria by which the five candidate wings were chosen for testing were the cruise and buffet characteristics in the transonic regime and the compatibility of the design with low speed (high-lift) requirements. Five wing-wide-body configurations were tested in the NASA Ames 11-foot transonic wind tunnel. Nacelles and pylons, flap support fairings, tail surfaces, and an outboard aileron were also tested on selected configurations.
    Keywords: AERODYNAMICS
    Type: ACEE-06-FR-9894 , NASA-CR-159332 , NAS 1.26:159332
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  • 2
    Publication Date: 2013-08-31
    Description: Wind tunnel tests were conducted to determine the low speed two dimensional aerodynamic characteristics of a 17 percent thick medium speed airfoil (MS(1)-0317) designed for general aviation applications. The results were compared with data for the 17 percent thick low speed airfoil (LS(1)-0417) and the 13 percent thick medium speed airfoil (MS(1)-0313). Theoretical predictions of the drag rise characteristics of this airfoil are also provided. The tests were conducted in the Langley low turbulence pressure tunnel over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2 million to 12 million, and an angle of attack range from about -8 to 20 deg.
    Keywords: AERODYNAMICS
    Type: L-13900 , NASA-TP-1786 , NAS 1.60:1786
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  • 3
    Publication Date: 2013-08-31
    Description: Changes to be made that provide increased accuracy and increased user flexibility in prediction of unsteady loadings caused by control surface motions are described. Analysis flexibility is increased by reducing the restrictions on the location of the downwash stations relative to the leading edge and the edges of the control surface boundaries. Analysis accuracy is increased in predicting unsteady loading for high Mach number analysis conditions through use of additional chordwise downwash stations. User guideline are presented to enlarge analysis capabilities of unusual wing control surface configurations. Comparative results indicate that the revised procedures provide accurate predictions of unsteady loadings as well as providing reductions of 40 to 75 percent in computer usage cost required by previous versions of this program.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145354-1 , NAS 1.26:145354-1
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  • 4
    Publication Date: 2013-08-31
    Description: An implicit finite difference scheme for an efficient computation of unsteady potential flow about airfoils is presented. The formulation uses density and velocity potential as dependent variables, and is cast in conservation form to assure the theoretically correct determination of shockwave location and speed. To enable boundary conditions to be imposed directly on the airfoil surface, a time varying sheared rectilinear coordinate transformation is employed. Calculated time history solutions on a pulsating airfoil are compared with the results of another unsteady transonic code. It is concluded that the method has excellent numerical stability and gives accurate solutions with sharply resolved shocks.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166152
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  • 5
    Publication Date: 2013-08-31
    Description: Water tunnel studies were performed to define the changes that occur in vortex flow fields above the wing due to spanwise blowing over the inboard and outboard wing panels and over the trailing-edge flaps. Flow visualization photographs were obtained for angles of attack up to 30 deg and sideslip angles up to 10 deg. The sensitivity of the vortex flows to changes in flap deflection angle, nozzle position, and jet momentum coefficient was determined. Deflection of the leading edge flap delayed flow separation and the formation of the wing vortex to higher angles of attack. Spanwise blowing delayed the breakdown of the wing vortex to farther outboard and to higher angles of attack. Spanwise blowing over the trailing edge flap entrained flow downward, producing a lift increase over a wide range of angles of attack. The sweep angle of the windward wing was effectively reduced in sideslip. This decreased the stability of the wing vortex, and it burst farther inboard. Reduced wing sweep required a higher blowing rate to maintain a stable vortex. A vortex could be stabilized on the outboard wing panel when an outboard blowing nozzle was used. Blowing from both an inboard and an outboard nozzle was found to have a favorable interaction.
    Keywords: AERODYNAMICS
    Type: NOR-80-138 , NASA-CR-163096
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  • 6
    Publication Date: 2013-08-31
    Description: Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/10 scale single engine agricultural airplane model. The configurations tested include the basic airplane, various wing leading edge and wing tip devices, elevator, aileron, and rudder control settings, and other modifications. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg, and clockwise and counter-clockwise rotations covering a spin coefficient range from 0 to .9.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3311
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  • 7
    Publication Date: 2013-08-31
    Description: The time dependent Navier-Stokes equations in mass averaged variables are solved for transonic flow over axisymmetric boattail plume simulator configurations. Numerical solution of these equations is accomplished with the unsplit explict finite difference algorithm of MacCormack. A grid subcycling procedure and computer code vectorization are used to improve computational efficiency. The two layer algebraic turbulence models of Cebeci-Smith and Baldwin-Lomax are employed for investigating turbulence closure. Two relaxation models based on these baseline models are also considered. Results in the form of surface pressure distribution for three different circular arc boattails at two free stream Mach numbers are compared with experimental data. The pressures in the recirculating flow region for all separated cases are poorly predicted with the baseline turbulence models. Significant improvements in the predictions are usually obtained by using the relaxation models.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1784 , L-13826
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  • 8
    Publication Date: 2013-08-31
    Description: The longitudinal and lateral directional aerodynamic characteristics for two Mach 5 cruise aircraft concepts were determined for test Mach numbers of 2.96, 3.96, and 4.63. Estimates from hypersonic impact theory and first order supersonic linearized theory were compared with data to indicate the usefulness of these methods. The method which applied tangent cone empirical theory to the body and tangent wedge theory to the wings and to the horizontal and vertical tails provided the best estimates. The tangent cone empirical theory applied to all components showed poor agreement with data, and the linear theory estimates were accurate only for lift coefficient and drag coefficient at low angles of attack.
    Keywords: AERODYNAMICS
    Type: L-13868 , NASA-TP-1767
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  • 9
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2013-08-31
    Description: The potential benefits were determined for the variable camber of commercial transport airplanes designed for intercontinental and domestic missions. A variable camber concept was developed and incorporated into airplanes designed for the two missions. Benefits were evaluated by comparing the mission performance and direct operating costs for the variable camber airplanes with those for reference airplanes designed for the same missions but having fixed geometry high speed wings. Several technical uncertainties associated with implementing variable camber were also examined.
    Keywords: AERODYNAMICS
    Type: NASA-CR-158930
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  • 10
    Publication Date: 2013-08-31
    Description: Two complementary methods of describing the high speed rotor noise problem are discussed. The first method uses the second order transonic potential equation to define and characterize the nature of the aerodynamic and acoustic fields and to explain the appearance of radiating shock waves. The second employs the Ffowcs Williams and Hawkings equation to successfully calculate the acoustic far field. Good agreement between theoretical and experimental waveforms is shown for transonic hover tip Mach numbers from 0.8 to 0.9.
    Keywords: AERODYNAMICS
    Type: AVRADCOM-TR-80-A-12 , NASA-TM-81236 , A-8342
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  • 11
    Publication Date: 2013-08-31
    Description: A stability test program was conducted to determine the effects of airspeed, collective pitch, rotor speed and shaft angle on stability and loads at speeds beyond that attained in the BMR/BO-105 flight test program. Loads and performance data were gathered at forward speeds up to 165 knots. The effect of cyclic pitch perturbations on rotor response was investigated at simulated level flight conditions. Two configuration variations were tested for their effect on stability. One variable was the control system stiffness. An axially softer pitch link was installed in place of the standard BO-105 pitch link. The second variation was the addition of elastomeric damper strips to increase the structural damping. The BMR was stable at all conditions tested. At fixed collective pitch, shaft angle and rotor speed, damping generally increased between hover and 60 knots, remained relatively constant from 60 to 90 knots, then decreased above 90 knots. Analytical predictions are in good agreement with test data up to 90 knots, but the trend of decreasing damping above 90 knots is contrary to the theory.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152373 , D210-11659-1
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  • 12
    Publication Date: 2013-08-31
    Description: The inviscid and viscid effects existing within the passages of a three bladed axial flow inducer operating at a flow coefficient of 0.065 are investigated. The blade static pressure and blade limiting streamline angle distributions were determined and the three components of mean velocity, turbulence intensities, and turbulence stresses were measured at locations inside the inducer blade passage utilizing a rotating three sensor hotwire probe. Applicable equations were derived for the hotwire data reduction analysis and solved numerically to obtain the appropriate flow parameters. The three dimensional inviscid flow in the inducer was predicted by numerically solving the exact equations of motion, and the three dimensional viscid flow was predicted by incorporating the dominant viscous terms into the exact equations. The analytical results are compared with the experimental measurements and design values where appropriate. Radial velocities are found to be of the same order as axial velocities within the inducer passage, confirming the highly three dimensional characteristic of inducer flow. Total relative velocity distribution indicate a substantial velocity deficiency near the tip at mid-passage which expands significantly as the flow proceeds toward the inducer trailing edge. High turbulence intensities and turbulence stresses are concentrated within this core region. Considerable wake diffusion occurs immediately downstream of the inducer trailing edge to decay this loss core. Evidence of boundary layer interactions, blade blockage effects, radially inward flows, annulus wall effects, and backflows are all found to exist within the long, narrow passages of the inducer.
    Keywords: AERODYNAMICS
    Type: PSU-AERSP-74-2 , NASA-CR-3333
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  • 13
    Publication Date: 2013-08-31
    Description: A computer program is presented which numerically solves an exact, full potential equation (FPE) for three dimensional, steady, inviscid flow through an isolated wind turbine rotor. The program automatically generates a three dimensional, boundary conforming grid and iteratively solves the FPE while fully accounting for both the rotating cascade and Coriolis effects. The numerical techniques incorporated involve rotated, type dependent finite differencing, a finite volume method, artificial viscosity in conservative form, and a successive line overrelaxation combined with the sequential grid refinement procedure to accelerate the iterative convergence rate. Consequently, the WIND program is capable of accurately analyzing incompressible and compressible flows, including those that are locally transonic and terminated by weak shocks. The program can also be used to analyze the flow around isolated aircraft propellers and helicopter rotors in hover as long as the total relative Mach number of the oncoming flow is subsonic.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1729 , E-474
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  • 14
    Publication Date: 2013-08-31
    Description: A finite difference code for predicting the high speed flow over the advancing helicopter rotor is presented. The code solves the low frequency, transonic small disturbance equation and is suitable for modeling the effects of advancing blade unsteadiness on blades of nearly arbitrary planform. The method employs a quasi-conservative mixed differencing scheme and solves the resulting difference equations by an alternating direction scheme. Computed results showed good agreement with experimental blade pressure data and illustrate some of the effects of varying the rotor planform. The flow unsteadiness is shown to be an indispensible part of a transonic solution. Close to the tip at high advance ratio, cross flow effects can significantly affect the solution.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1721 , A-8024
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  • 15
    Publication Date: 2013-08-31
    Description: Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/6 scale, single engine, low wing, general aviation model (model C). The configurations tested included the basic airplane and control deflections, wing leading edge and fuselage modification devices, tail designs and airplane components. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg and clockwise and counter clockwise rotations covering an omega b/2v range from 0 to .9.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3200
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  • 16
    Publication Date: 2013-08-31
    Description: A computer implemented numerical method for predicting the flow in and about an isolated three dimensional jet exhaust nozzle is summarized. The approach is based on an implicit numerical method to solve the unsteady Navier-Stokes equations in a boundary conforming curvilinear coordinate system. Recent improvements to the original numerical algorithm are summarized. Equations are given for evaluating nozzle thrust and discharge coefficient in terms of computed flowfield data. The final formulation of models that are used to simulate flow turbulence effect is presented. Results are presented from numerical experiments to explore the effect of various quantities on the rate of convergence to steady state and on the final flowfield solution. Detailed flowfield predictions for several two and three dimensional nozzle configurations are presented and compared with wind tunnel experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3264 , LMSC-D678888-PT-2
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  • 17
    Publication Date: 2013-08-31
    Description: Calculated and measured values of helicopter rotor flapping angles in forward flight are compared for a model rotor in a wind tunnel and an autogiro in gliding flight. The lateral flapping angles can be accurately predicted when a calculation of the nonuniform wake-induced velocity is used. At low advance ratios, it is also necessary to use a free wake geometry calculation. For the cases considered, the tip vortices in the rotor wake remain very close to the tip-path plane, so the calculated values of the flapping motion are sensitive to the fine details of the wake structure, specifically the viscous core radius of the tip vortices.
    Keywords: AERODYNAMICS
    Type: A-8239 , NASA-TM-81213 , AVRADCOM-TR-80-A-11
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  • 18
    Publication Date: 2013-08-31
    Description: A flight investigation produced data on performance and rotor loads for a teetering rotor, AH-1G helicopter flown with a main rotor that had the NLR-1T airfoil as the blade section contour. The test envelope included hover, forward flight speeds from 34 to 83 m/sec (65 to 162 knots), and collective fixed maneuvers at about 0.25 tip speed ratio. The data set for each test point describes vehicle flight state, control positions, rotor loads, power requirements, and blade motions. Rotor loads are reviewed primarily in terms of peak to peak and harmonic content. Lower frequency components predominated for most loads and generally increased with increased airspeed, but not necessarily with increased maneuver load factor. Detailed data for an advanced airfoil on an AH-1G are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81871 , AVRADCOM-TM-80-B-2
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  • 19
    Publication Date: 2013-08-31
    Description: In order to provide experimental data for comparison with newly developed finite difference methods for computing supersonic flows over aircraft configurations, wind tunnel tests were conducted on four arrow wing models. The models were machined under numeric control to precisely duplicate analytically defined shapes. They were heavily instrumented with pressure orifices at several cross sections ahead of and in the region where there is a gap between the body and the wing trailing edge. The test Mach numbers were 2.36, 2.96, and 4.63. Tabulated pressure data for the complete test series are presented along with selected oil flow photographs. Comparisons of some preliminary numerical results at zero angle of attack show good to excellent agreement with the experimental pressure distributions.
    Keywords: AERODYNAMICS
    Type: L-13703 , NASA-TM-81835
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  • 20
    Publication Date: 2013-08-31
    Description: Digitally acquired and processed results from an experimental investigation of grid generated turbulence of various scales through and downstream of nine matched cubic contour contractions ranging in area ratio from 2 to 36, and in length to inlet diameter ratio from 0.25 to 1.50 are reported. An additional contraction with a fifth order contour was also utilized for studying the shape effect. Thirteen homogeneous and nearly isotropic test flow conditions with a range of turbulence intensities, length scales and Reynolds numbers were generated and used to examine the sensitivity of the contractions to upstream turbulence. The extent to which the turbulence is altered by the contraction depends on the incoming turbulence scales, the total strain experienced by the fluid, as well as the contraction ratio and the strain rate. Varying the turbulence integral scale influences the transverse turbulence components more than the streamwise component. In general, the larger the turbulence scale, the lesser the reduction in the turbulence intensity of the transverse components. Best agreement with rapid distortion theory was obtained for large scale turbulence, where viscous decay over the contraction length was negligible, or when a first order correction for viscous decay was applied to the results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165136 , R80-1
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  • 21
    Publication Date: 2013-08-31
    Description: A theoretical analysis is presented yielding sets of partial differential equations for determination of turbulent aerodynamic flowfields in the vicinity of an airfoil trailing edge. A four phase interaction algorithm is derived to complete the analysis. Following input, the first computational phase is an elementary viscous corrected two dimensional potential flow solution yielding an estimate of the inviscid-flow induced pressure distribution. Phase C involves solution of the turbulent two dimensional boundary layer equations over the trailing edge, with transition to a two dimensional parabolic Navier-Stokes equation system describing the near-wake merging of the upper and lower surface boundary layers. An iteration provides refinement of the potential flow induced pressure coupling to the viscous flow solutions. The final phase is a complete two dimensional Navier-Stokes analysis of the wake flow in the vicinity of a blunt-bases airfoil. A finite element numerical algorithm is presented which is applicable to solution of all partial differential equation sets of inviscid-viscous aerodynamic interaction algorithm. Numerical results are discussed.
    Keywords: AERODYNAMICS
    Type: COMCO-80-TR-1.0 , NASA-CR-3301
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  • 22
    Publication Date: 2013-08-31
    Description: A computational method for simulating the aerodynamics of wing-fuselage configurations at transonic speeds is developed. The finite difference scheme is characterized by a multiple embedded mesh system coupled with a modified or extended small disturbance flow equation. This approach permits a high degree of computational resolution in addition to coordinate system flexibility for treating complex realistic aircraft shapes. To augment the analysis method and permit applications to a wide range of practical engineering design problems, an arbitrary fuselage geometry modeling system is incorporated as well as methodology for computing wing viscous effects. Configuration drag is broken down into its friction, wave, and lift induced components. Typical computed results for isolated bodies, isolated wings, and wing-body combinations are presented. The results are correlated with experimental data. A computer code which employs this methodology is described.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3243
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  • 23
    Publication Date: 2013-08-31
    Description: A cryogenic wind tunnel is based on the twofold idea of lowering drive power and increasing Reynolds number by operating with nitrogen near its boiling point. There are two possible types of condensation problems involved in this mode of wind tunnel operation. They concern the expansion from the nozzle supply to the test section at relatively low cooling rates, and secondly the expansion around models in the test section. This secondary expansion involves higher cooling rates and shorter time scales. In addition to these two condensation problems it is not certain what purity of nitrogen can be achieved in a large facility. Therefore, one cannot rule out condensation processes other than those of homogeneous nucleation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163217
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  • 24
    Publication Date: 2013-08-31
    Description: The nonlinear propagation equations for sound generated by a constant speed blade tip are presented. Propagation from a subsonic tip is treated as well as the various cases that can occur at supersonic speeds. Some computed examples indicate that the nonlinear theory correlates with experimental results better than linear theory for large amplitude waves. For swept tips that generate a wave with large amplitude leading expansion, the nonlinear theory predicts a cancellation effect that results in a significant reduction of both amplitude and impulse.
    Keywords: AERODYNAMICS
    Type: L-13388 , NASA-TP-1660
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  • 25
    Publication Date: 2013-08-31
    Description: An experimental investigation of inviscid real-gas effects on the pressure distribution along the Space Shuttle Orbiter nose center line up to an angle of attack of 32 deg was performed in support of the Shuttle Entry Air Data System (SEADS). Free-stream velocities from 4.8 to 6.6 kn/s were generated at hypersonic conditions with helium, air, and CO2, resulting in normal-shock density ratios from 3.7 to 18.4. The experimental results for pressure distribution agreed closely with numerical results. Modified Newtonian theory deviates from both experiment and the numerical results as angle of attack increases or shock density ratio decreases. An evaluation of the use of modified Newtonian theory for predicting SEADS pressure distributions in actual flight conditions was made through comparison with numerical predictions.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1627 , L-13341
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  • 26
    Publication Date: 2013-08-31
    Description: Turbulent flow within turbomachines having arbitrary blade geometries is examined. Effects of turbulence are modeled using two equations, one expressing the development of the turbulence kinetic energy and the other its dissipation rate. To account for complicated blade geometries, the flow equations are formulated in terms of a nonorthogonal boundary fitted coordinate system. The analysis is applied to a radial inflow turbine. The solution obtained indicates the severity of the complex interaction mechanism that occurs between the different flow regimes (i.e., boundary layers, recirculating eddies, separation zones, etc.). Comparison with nonviscous flow solutions tend to justify strongly the inadequacy of using the latter with standard boundary layer techniques to obtain viscous flow details within turbomachine rotors. Capabilities and limitations of the present method of analysis are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159636
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  • 27
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2013-08-31
    Description: The curved flow capability of a stability wind tunnel was used to investigate the lateral directional characteristics of an F-18 aircraft. The model is described and the procedures used to obtain and correct the data and a graphical presentation of the results are presented. The results include graphs of lateral directional derivatives versus sideslip or static plots, the lateral directional static stability derivatives versus angle of attack, and finally the lateral directional derivatives versus nondimensional yaw rate for different angles of attack and sideslip. Results are presented for several configurations including complete, complete without vertical tails, complete without horizontal tails, fuselage wing and fuselage alone. Each of these were tested with and without wing leading edge extensions.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:169345 , VPI-AERO-108 , NASA-CR-169345
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  • 28
    Publication Date: 2013-08-31
    Description: A user's guide is provided for a computer code which calculates the laminar and turbulent hypersonic flows about blunt axisymmetric bodies, such as spherically blunted cones, hyperboloids, etc., at zero and small angles of attack. The code is written in STAR FORTRAN language for the CDC-STAR-100 computer. Time-dependent, viscous-shock-layer-type equations are used to describe the flow field. These equations are solved by an explicit, two-step, time asymptotic, finite-difference method. For the turbulent flow, a two-layer, eddy-viscosity model is used. The code provides complete flow-field properties including shock location, surface pressure distribution, surface heating rates, and skin-friction coefficients. This report contains descriptions of the input and output, the listing of the program, and a sample flow-field solution.
    Keywords: AERODYNAMICS
    Type: NASA-TM-80202 , NAS 1.15:80202
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  • 29
    Publication Date: 2013-08-31
    Description: The results are presented of a wind tunnel test utilizing a 4.7-percent-scale semispan model of the DC-10 in the Calspan 8-foot transonic wind tunnel. The effect of a revised long-duct nacelle shape on the channel velocities, the incremental drag relative to the baseline long-duct nacelle, and channel velocities for the baseline long-duct nacelle were determined and compared with data obtained at Ames. The baseline and the revised long-duct nacelles are representative of a CF6-50 mixed-flow configuration and were evaluated on a model of a proposed DC-10 stretched-fuselage configuration. The results showed that the revised long-duct nacelle has an appreciable effect on the inboard channel velocities, resulting in an increased channel Mach number. However, the pressure recovery on the nacelle afterbody was about the same for both nacelles. The lift curves for both long-duct nacelle configurations were the same. The channel pressures measured at Calspan were in good agreement with those measured at Ames for the baseline long-duct nacelle. The incremental drag for the revised nacelle was measured as two to four counts (three counts is approximately equal to one percent of the airplane drag) higher than that of the baseline long-duct nacelle.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159271 , ACEE-17-FR-9005 , NAS 1.26:159271
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  • 30
    Publication Date: 2013-08-31
    Description: As a means to achieve a minimum interference correction wind tunnel, a partially actively controlled test section was experimentally examined. A jet flapped wing with 0.91 m (36 in) span and R = 4.05 was used as a model to create moderately high lift coefficients. The partially controlled test section was simulated using an insert, a rectangular box 0.96 x 1.44 m (3.14 x 4.71 ft) open on both ends in the direction of the tunnel air flow, placed in the University of Washington Aeronautical Laboratories (UWAL) 2.44 x 3.66 m (8 x 12 ft) wind tunnel. A tail located three chords behind the wing was used to measure the downwash at the tail region. The experimental data indicates that, within the range of momentum coefficient examined, it appears to be unnecessary to actively control all four sides of the test section walls in order to achieve the near interference free flow field environment in a small wind tunnel. The remaining wall interference can be satisfactorily corrected by the vortex lattice method.
    Keywords: AERODYNAMICS
    Type: NASA-CR-164439
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  • 31
    Publication Date: 2013-08-31
    Description: The investigation was conducted at static conditions and over a Mach number range from 0.6 to 1.2. Angle of attack was held constant at 0 deg. High pressure air was used to simulate jet exhaust flow at ratios of jet total pressure to free-stream static pressure from 1 (jet off) to approximately 10. Sidewall cutback appears to be a viable way of reducing nozzle weight and cooling requirements without compromising installed performance.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1771 , L-13826
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  • 32
    Publication Date: 2013-08-31
    Description: A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic translational model flutter. This five bladed cascade had a solidity of 1.52 and a setting angle of 0.90 rad. Unique graphite epoxy airfoils were fabricated to achieve the realistic high reduced frequency level of 0.15. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time steady and time unsteady flow field surrounding the center cascade airfoil were investigated.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165166 , EDR-10361-VOL-2
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  • 33
    Publication Date: 2013-08-31
    Description: Three new options were incorporated into an existing computer program for the design and analysis of low speed airfoils. These options permit the analysis of airfoils having variable chord (variable geometry), a boundary layer displacement iteration, and the analysis of the effect of single roughness elements. All three options are described in detail and are included in the FORTRAN IV computer program.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81862
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  • 34
    Publication Date: 2013-08-31
    Description: Six interchangeable tip shapes were tested: a square (baseline) tip, an ogee tip, a subwing tip, a swept tip, a winglet tip, and a short ogee tip. In hover at the lower rotational speeds the swept, ogee, and short ogee tips had about the same torque coefficient, and the subwing and winglet tips had a larger torque coefficient than the baseline square tip blades. The ogee and swept tip blades required less torque coefficient at lower rotational speeds and roughly equivalent torque coefficient at higher rotational speeds compared with the baseline square tip blades in forward flight. The short ogee tip required higher torque coefficient at higher lift coefficients than the baseline square tip blade in the forward flight test condition.
    Keywords: AERODYNAMICS
    Type: AVRADCOM-TR-79-49 , L-12774 , NASA-TM-80080
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  • 35
    Publication Date: 2013-08-31
    Description: The wind tunnel tests were conducted both with and without boundary layer trips at Mach 3 and nominal free stream Reynolds numbers per meter ranging from 3.3 x 10 the 6th power. Instrumentation consisted of pressure orifices, thermocouples, a boundary layer pitot pressure rake, and a floating element skin friction balance. Measurements from both wind tunnel and flight were compared with existing engineering prediction methods.
    Keywords: AERODYNAMICS
    Type: L-14044 , NASA-TP-1789
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  • 36
    Publication Date: 2013-08-31
    Description: A computational system for estimation of nonlinear aerodynamic characteristics of wings at supersonic speeds was developed and was incorporated in a computer program. This corrected linearized theory method accounts for nonlinearities in the variation of basic pressure loadings with local surface slopes, predicts the degree of attainment of theoretical leading edge thrust, and provides an estimate of detached leading edge vortex loadings that result when the theoretical thrust forces are not fully realized.
    Keywords: AERODYNAMICS
    Type: L-13589 , NASA-TP-1718
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  • 37
    Publication Date: 2013-08-31
    Description: The question of the effect of distribution and magnitude of spanwise circulation and shed vorticity from an airplane wing on the distribution pattern of agricultural products distributed from an airplane was studied. The first step in an analysis of this question is the determination of the actual distribution of lift along an airplane wing, from which the pattern of shed vorticity can be determined. A procedure is developed to calculate the span loading for flapped and unflapped wings of arbitrary aspect ratio and taper ratio. The procedure was programmed on a small programmable calculator, the Hewlett Packard HP-97, and also was programmed in BASIC language. They could be used to explore the variations in span loading that can be secured by variable flap deflections or the effect of flying at varying air speeds at different airplane gross weights. Either an absolute evaluation of span loading can be secured or comparative span loading can be evaluated to determine their effect on swath width and swath distribution pattern. The programs are intended to assist the user in evaluating the effect of a given spanload distribution.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159329
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  • 38
    Publication Date: 2013-08-31
    Description: A conformal mapping method for the design of airfoils with prescribed velocity distribution characteristics, a panel method for the analysis of the potential flow about given airfoils, and a boundary layer method have been combined. With this combined method, airfoils with prescribed boundary layer characteristics can be designed and airfoils with prescribed shapes can be analyzed. All three methods are described briefly. The program and its input options are described. A complete listing is given as an appendix.
    Keywords: AERODYNAMICS
    Type: NASA-TM-80210 , L-12937
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  • 39
    Publication Date: 2013-08-31
    Description: A general design method was developed for steady, three dimensional, potential, incompressible or subsonic-compressible flow. In this design method, the flow field, including the shape of its boundary, was determined for arbitrarily specified, continuous distributions of velocity as a function of arc length along the boundary streamlines. The method applied to the design of both internal and external flow fields, including, in both cases, fields with planar symmetry. The analytic problems associated with stagnation points, closure of bodies in external flow fields, and prediction of turning angles in three dimensional ducts were reviewed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3288
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  • 40
    Publication Date: 2013-08-31
    Description: An experiment was conducted at static conditions to determine the internal performance effects of nozzle throat contouring, the result of increasing the circular-arc throat radius. Five nonaxisymmetric converging-diverging nozzles were tested at nozzle pressure ratios up to 9.0. Data are presented as internal thrust ratios, discharge coefficients, and static-pressure distributions. Comparisons of internal performance data for the five nozzles show that throat contouring increases the value of discharge coefficient but has no significant effect on internal thrust ratio except in cases of internal flow separation. To illustrate the use of the two dimensional converging-diverging (2-D C-D) nozzle data base, a two dimensional inviscid theory was applied to the five configurations. The generally good agreement of data with theoretical results indicates that two-dimensional inviscid theory can be successfully applied to the prediction of 2-D C-D nozzle internal flow.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1704 , L-13591
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  • 41
    Publication Date: 2013-08-31
    Description: Two dimensional incompressible flow over wavy surfaces are analyzed numerically by spectral methods. Algorithms for periodic flows (Fourier modes in the periodic flow direction and Chebycheff modes in the normal direction), and inflow-outflow boundary conditions (Chebycheff modes used in both directions) are described. Results obtained using both codes are reported for laminar flows. Comparisons with known theoretical and experimental results are made.
    Keywords: AERODYNAMICS
    Type: CHI-41 , NASA-CR-159305
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  • 42
    Publication Date: 2013-08-31
    Description: The data and major conclusions obtained from an experimental/analytical study of upper-surface blown (USB) configurations at cruise are summarized. The high-speed (subsonic) experimental work, studying the aerodynamic effects of wing-nacelle geometric variations, was conducted around semi-span model configurations composed of diversified, interchangeable components. Power simulation was provided by high pressure air ducted through closed forebody nacelles. Nozzle geometry was varied across size, exit aspect ratio, exit position and boattail angle. Both 3-D force and 2-D pressure measurements were obtained at cruise Mach numbers from 0.5 to 0.8 and at nozzle pressure ratios up to about 3.0. The experimental investigation was supported by an analytical synthesis of the system using a vortex lattice representation with first-order power effects. Results are also presented from a compatibility study in which a short-haul transport is designed on the basis of the aerodynamic findings in the experimental study as well as acoustical data obtained in a concurrent program. High-lift test data are used to substantiate the projected performance of the selected transport design.
    Keywords: AERODYNAMICS
    Type: LG77ER0028 , NASA-CR-3193
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  • 43
    Publication Date: 2013-08-31
    Description: Recent developments of the Green's function method and the computer program SOUSSA (Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics) are reviewed and summarized. Applying the Green's function method to the fully unsteady (transient) potential equation yields an integro-differential-delay equation. With spatial discretization by the finite-element method, this equation is approximated by a set of differential-delay equations in time. Time solution by Laplace transform yields a matrix relating the velocity potential to the normal wash. Premultiplying and postmultiplying by the matrices relating generalized forces to the potential and the normal wash to the generalized coordinates one obtains the matrix of the generalized aerodynamic forces. The frequency and mode-shape dependence of this matrix makes the program SOUSSA useful for multiple frequency and repeated mode-shape evaluations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159130 , ASI-TR-78-45-VOL-1
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  • 44
    Publication Date: 2013-08-31
    Description: A method for generating two dimensional finite difference grids about airfoils and other shapes by the use of the Poisson differential equation is developed. The inhomogeneous terms are automatically chosen such that two important effects are imposed on the grid at both the inner and outer boundaries. The first effect is control of the spacing between mesh points along mesh lines intersecting the boundaries. The second effect is control of the angles with which mesh lines intersect the boundaries. A FORTRAN computer program has been written to use this method. A description of the program, a discussion of the control parameters, and a set of sample cases are included.
    Keywords: AERODYNAMICS
    Type: A-8178 , NASA-TM-81198
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  • 45
    Publication Date: 2013-08-31
    Description: The model hardware, test facilities and instrumentation utilized in an experimental study of upper surface blown configurations at cruise is described. The high speed (subsonic) experimental work, studying the aerodynamic effects of wing nacelle geometric variations, was conducted around semispan model configurations composed of diversified, interchangeable components. Power simulation was provided by high pressure air ducted through closed forebody nacelles. Nozzle geometry was varied across size, exit aspect ratio, exit position and boattail angle. Three dimensional force and two dimensional pressure measurements were obtained at cruise Mach numbers from 0.5 to 0.8 and at nozzle pressure ratios up to about 3.0. The experimental investigation was supported by an analytical synthesis of the system using a vortex lattice representation with first order power effects. Results are also presented from a compatibility study in which a short haul transport is designed on the basis of the aerodynamic findings in the experimental study as well as acoustical data obtained in a concurrent program. High lift test data are used to substantiate the projected performance of the selected transport design.
    Keywords: AERODYNAMICS
    Type: LG77FR0028 , NASA-CR-3192
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  • 46
    Publication Date: 2013-08-31
    Description: A potential-flow panel method was modified to calculate the effects of a rotor wake on the time-averaged surface pressure and velocity distributions on a helicopter fuselage. The rotor-induced velocities are calculated by using a vortex-tube wake model. The calculated pressure distributions are found to compare well with experimental data obtained from tests of a wind-tunnel model.
    Keywords: AERODYNAMICS
    Type: AVRADCOM-TR-80-B-3 , NASA-TP-1656 , L-13363
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  • 47
    Publication Date: 2013-08-31
    Description: A method for solving the linear integral equations of incompressible potential flow in three dimensions is presented. Both analysis (Neumann) and design (Dirichlet) boundary conditions are treated in a unified approach to the general flow problem. The method is an influence coefficient scheme which employs source and doublet panels as boundary surfaces. Curved panels possessing singularity strengths, which vary as polynomials are used, and all influence coefficients are derived in closed form. These and other features combine to produce an efficient scheme which is not only versatile but eminently suited to the practical realities of a user-oriented environment. A wide variety of numerical results demonstrating the method is presented.
    Keywords: AERODYNAMICS
    Type: D6-43808 , NASA-CR-3079
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  • 48
    Publication Date: 2013-08-31
    Description: An experimental evaluation of the effects of free stream turbulence on the performance of a subsonic two dimensional diffuser was made. The diffuser's static pressure recovery coefficient was increased 11.4 and 21.1 percent at total. Divergence angles of 12 and 20 degrees respectively were obtained when the value of the inlet integral free stream scale of turbulence in the flow direction was at least 7.5 times larger than the inlet boundary layer displacement thickness, and when the inlet total free stream turbulence intensity was at least 3.5 percent. It is hypothesized that a larger scale of turbulence transmits the free stream energy to the wall more effectively and when coupled with large turbulence intensities, acts to decrease the distortion and delay separation within the diffuser.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163194
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  • 49
    Publication Date: 2013-08-31
    Description: A survey was made of the parameters affecting the development of the leeward symmetric separated flow over slender delta wings immersed in a supersonic stream. The parameters included Mach number, Reynolds number, angle of attack, leading-edge sweep angle, and body cross-sectional shape, such that subsonic and supersonic leading-edge flows are encountered. It was seen that the boundaries between the various flow regimes existing about the leeward surface may conveniently be represented on a diagram with the components of angle of attack and Mach number normal to the leading edge as governing parameters.
    Keywords: AERODYNAMICS
    Type: A-8117 , NASA-TM-81187
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  • 50
    Publication Date: 2013-08-31
    Description: A viscous-inviscid interaction model was developed to account for jet entrainment effects in the prediction of the subsonic flow over nozzle afterbodies. The model is based on the concept of a weakly interacting shear layer in which the local streamline deflections due to entrainment are accounted for by a displacement-thickness type of correction to the inviscid plume boundary. The entire flow field is solved in an iterative manner to account for the effects on the inviscid external flow of the turbulent boundary layer, turbulent mixing and chemical reactions in the shear layer, and the inviscid jet exhaust flow. The components of the computational model are described, and numerical results are presented to illustrate the interactive effects of entrainment on the overall flow structure. The validity of the model is assessed by comparisons with data obtained form flow-field measurements on cold-air jet exhausts. Numerical results and experimental data are also given to show the entrainment effects on nozzle boattail drag under various jet exhaust and free-stream flow conditions.
    Keywords: AERODYNAMICS
    Type: L-13362 , NASA-TP-1626
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  • 51
    Publication Date: 2013-08-31
    Description: Experimental/theoretical correlations are presented which show that significant levels of leading-edge thrust are possible at supersonic speeds for certain planforms having the geometry to support the theoretical thrust-distribution potential. The new analytical process employed provides not only the level of leading-edge thrust attainable but also the spanwise distribution of both it and that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.
    Keywords: AERODYNAMICS
    Type: L-13316 , NASA-TP-1632
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  • 52
    Publication Date: 2013-08-31
    Description: A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane.
    Keywords: AERODYNAMICS
    Type: H-1079 , NASA-TP-1516
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  • 53
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2013-08-31
    Description: For slow flapping motions it is found that the minimum energy loss occurs when the vortex wake moves as a rigid surface that rotates about the wing root - a condition analogous to that determined for a slow-turning propeller. The optimum circulation distribution determined by this condition differs from the elliptic distribution, showing a greater concentration of lift toward the tips. It appears that very high propulsive efficiencies are obtained by flapping.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81174 , A-8076
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  • 54
    Publication Date: 2013-08-31
    Description: The results of calculations necessary to extrapolate performance data on an XB-70-1 wind tunnel model to full scale at Mach numbers from 0.76 to 2.53 are presented. The extrapolation was part of a joint program to evaluate performance prediction techniques for large flexible supersonic airplanes similar to a supersonic transport. The extrapolation procedure included: interpolation of the wind tunnel data at the specific conditions of the flight test points; determination of the drag increments to be applied to the wind tunnel data, such as spillage drag, boundary layer trip drag, and skin friction increments; and estimates of the drag items not represented on the wind tunnel model, such as bypass doors, roughness, protuberances, and leakage drag. In addition, estimates of the effects of flexibility of the airplane were determined.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1515 , L-12688
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  • 55
    Publication Date: 2013-08-31
    Description: A computer code is described which may be used to calculate the steady, supersonic, three-dimensional, inviscid flow over blunt bodies. The theoretical and numerical formulation of the problem is given (shock-capturing, downstream marching), including exposition of the boundary and initial conditions. The overall flow logic of the program, its usage, accuracy, and limitations are discussed.
    Keywords: AERODYNAMICS
    Type: NEAR-TR-189 , NASA-CR-3223
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  • 56
    Publication Date: 2013-08-31
    Description: An implicit, shock-capturing finite-difference code which is used to calculate two-dimensional inlet flow fields in a supersonic free stream is explained. The Euler equations are subjected to general nonorthogonal transformation and a body-fitted coordinate system is employed. The mathematical formulation of the problem is given along with the numerical algorithm. Initial and boundary conditions, numerical stability, program limitations, and accuracy is discussed. An overall program logic as well as instructions for program use and operation are also furnished.
    Keywords: AERODYNAMICS
    Type: NEAR-TR-193 , NASA-CR-3222
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  • 57
    Publication Date: 2013-08-31
    Description: A supercritical wing with an aspect ratio of 10.76 and with two trailing-edge oscillating control surfaces is described. The semispan wing is instrumented with 252 static orifices and 164 in situ dynamic-pressure gages for studying the effects of control-surface position and motion on steady- and unsteady-pressures at transonic speeds. Results from initial tests conducted in the Langley Transonic Dynamics Tunnel at two Reynolds numbers are presented in tabular form.
    Keywords: AERODYNAMICS
    Type: L-13964 , NASA-TM-81888 , NAS 1.15:81888
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  • 58
    Publication Date: 2013-08-31
    Description: A method and program called TRANSEP is presented that can be used for the analysis of the flow about a low speed airfoil under high lift, massive separation conditions. Since the present program is a modification of the direct-inverse TRANDES code, it can also be used for the design and analysis of transonic airfoils, including the effects of weak viscous interaction. Interactions on program usage, program modifications to convert TRANDES to TRANSEP, and sample cases and results are given.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3376 , NAS 1.26:3376
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  • 59
    Publication Date: 2013-08-31
    Description: The low speed aerodynamic characteristics of a 14 percent thick supercritical airfoil are documented. The wind tunnel test was conducted in the Low Turbulence Pressure Tunnel. The effects of varying chord Reynolds number from 2,000,000 to 18,000,000 at a Mach number of 0.15 and the effects of varying Mach number from 0.10 to 0.32 at a Reynolds number of 6,000,000 are included.
    Keywords: AERODYNAMICS
    Type: NAS 1.15:81912 , NASA-TM-81912 , L-13891
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  • 60
    Publication Date: 2013-08-31
    Description: The listings, user's instructions, sample inputs, and sample outputs of two computer programs which are especially useful in obtaining an approximate solution of the viscous flow over an arbitrary nonlifting three dimensional body are provided. The first program performs a potential flow solution by a well known panel method and readjusts this initial solution to account for the effects of the boundary layer displacement thickness, a nonuniform but unidirectional onset flow field, and the presence of air intakes and exhausts. The second program is effectually a geometry package which allows the user to change or refine the shape of a body to satisfy particular needs without a significant amount of human intervention. An effort to reduce the cruise drag of light aircraft through an analytical study of the contributions to the drag arising from the engine cowl shape and the foward fuselage area and also that resulting from the cooling air mass flowing through intake and exhaust sites on the nacelle is presented. The programs may be effectively used to determine the appropriate body modifications or flow port locations to reduce the cruise drag as well as to provide sufficient air flow for cooling the engine.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159380 , NAS 1.26:159380
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  • 61
    Publication Date: 2013-08-31
    Description: The contributions to the cruise drag of light aircraft arising from the shape of the engine cowl and the forward fuselage area and also that resulting from the cooling air mass flow through intake and exhaust sites on the nacelle were analyzed. The methods employed for the calculation of the potential flow about an arbitrary three dimensional body are described with modifications to include the effects of boundary layer displacement thickness, a nonuniform onset flow field (such as that due to a rotating propeller), and the presence of air intakes and exhausts. A simple, reliable, largely automated scheme to better define or change the shape of a body is also presented. A technique was developed which can yield physically acceptable skin friction and pressure drag coefficients for isolated light aircraft bodies. For test cases on a blunt nose Cessna 182 fuselage, the technique predicted drag reductions as much as 28.5% by body recontouring and proper placements and sizing of the cooling air intakes and exhausts.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:159379 , NASA-CR-159379
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  • 62
    facet.materialart.
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    In:  CASI
    Publication Date: 2013-08-31
    Description: The lateral directional characteristics of an F-18 aircraft was investigated. Aerodynamic derivatives associated with pure roll rate, or the 'p' derivatives were obtained. The model is described and the procedures used to obtain and correct the data, and a graphical presentation of the results are presented. These results include graphs of the lateral directional static stability derivatives versus angle of attack, and the lateral directional force and moment coefficients versus nondimensional roll rate. Results are presented for several configurations including complete, complete without vertical tails, complete without horizontal tails, fuselage wing and fuselage alone. Each of these configuations was tested with and without wing leading edge extensions. The basic control surfaces were deflected and the results were investigated.
    Keywords: AERODYNAMICS
    Type: NASA-CR-169344 , NAS 1.26:169344
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  • 63
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2013-08-31
    Description: The effects of blade planform and tip speed on noise and performance for a Hughes 500 C rotor system were studied. A cursory examination of the effects of such planform shapes as regular, inverse, and no taper on the noise and performance of the rotor was conducted. It was found that a constant width wide chord planform at tower tip speed provided the best performance and lowest noise. The tapered planforms had lower performance figures due to the reduced solidity. However, some noise reductions were achieved.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166256 , T-35584
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  • 64
    Publication Date: 2013-08-31
    Description: Modifications were made to the model to improve longitudinal acceleration capability during transition from hovering to wing borne flight. A rearward deflection of the fuselage augmentor thrust vector is shown to be beneficial in this regard. Other agmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also shows negligible influence on the performance of the wing and of the fuselage augmentor.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152380 , DHC-DND-80-1
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  • 65
    Publication Date: 2013-08-31
    Description: The method of vortex discretization is used to analyze the interaction of the vorticity generated by a strake, with the flow over a delta wing. The validity of the approach is first established by making comparisons with established methods for dealing with delta wings, after which compound delta planforms are discussed. An understanding of the favorable interference effects normally associated with this type of configuration is obtained and results are presented to quantify the expected lift increments resulting from the strake interaction.
    Keywords: AERODYNAMICS
    Type: SU-JIAA-TR-30 , NASA-CR-166183
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  • 66
    Publication Date: 2013-08-31
    Description: The rolling up of the trailing vortex sheet produced by a wing of finite span was calculated as a series expansion in time. For a vorticity distribution corresponding to a wing with cusped tips, the shape of the sheet was found by summing the series using Pade approximants. The sheet remains analytic for some time but ultimately develops an exponential spiral at the tips. The centroid of vorticity was conserved to high accuracy.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166182 , SU-JIAA-TR-32
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  • 67
    Publication Date: 2013-08-31
    Description: A simplified model is used to describe the interaction between a propeller slipstream and a wing in the transonic regime. The undisturbed slipstream boundary is assumed to coincide with an infinite circular cylinder. The undisturbed slipstream velocity is rotational and is a function of the radius only. In general, the velocity perturbation caused by introducing a wing into the slipstream is also rotational. By making small disturbance assumptions, however, the perturbation velocity becomes nearly potential, and an approximation for the flow is obtained by solving a potential equation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152351
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  • 68
    Publication Date: 2013-08-31
    Description: A wind tunnel test was conducted to measure the aerodynamic characteristics of two horizontal attitude takeoff and landing V/STOL fighter/attack aircraft concepts. In one concept, a jet diffuser ejector was used for the vertical lift system; the other used a remote augmentation lift system (RALS). Wind tunnel tests to investigate the aerodynamic uncertainties and to establish a data base for these types of concepts were conducted over a Mach number range from 0.2 to 2.0. The present report covers tests, conducted in the 11 foot transonic wind tunnel, for Mach numbers from 0.4 to 1.4. Detailed effects of varying the angle of attack (up to 27 deg), angle of sideslip (-4 deg to +8 deg), Mach number, Reynolds number, and configuration buildup were investigated. In addition, the effects of wing trailing edge flap deflections, canard incidence, and vertical tail deflections were explored. Variable canard longitudinal location and different shapes of the inboard nacelle body strakes were also investigated.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81234 , A-8338
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  • 69
    Publication Date: 2013-08-31
    Description: Calibration data for the two dimensional test section of the Langley 0.3 Meter Transonic Cryogenic Tunnel were used to develop a Mach number-Reynolds number correlation for the fan pressure ratio in terms of test section conditions. Well established engineering relationships combined to form an equation which is functionally analogous to the correlation. A geometric loss coefficient which is independent of Reynolds number or Mach number was determined. Present and anticipated uses of this concept include improvement of tunnel control schemes, comparison of efficiencies for operationally similar wind tunnels, prediction of tunnel test conditions and associated energy usage, and determination of Reynolds number scaling laws for similar fluid flow systems.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1752 , L-13713
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  • 70
    Publication Date: 2013-08-31
    Description: The effectiveness of leading edge concepts for minimizing or controlling leading edge flow separation was studied. Emphasis was placed on low speed performance, stability, and control characteristics of configurations with highly swept wings. Simple deflection of the leading edge, a variable camber leading edge system, and a leading edge vortex flow system were among the concepts studied. The data are presented without analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-80180
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  • 71
    Publication Date: 2013-08-31
    Description: Experimental data on the unsteady aerodynamics of oscillating airfoils in transonic flow are presented. Two 0.5 m-chord airfoil models - an NACA 64A010 and an NLR 7301 - were tested in the NASA-Ames 11 by 11 foot Transonic Wind Tunnel at Mach numbers to 0.85, at chord Reynolds numbers to 12 million and at mean angles of attack to 4 deg. The airfoils were subjected to both pitching and plunging motions at reduced frequencies to 0.3 (physical frequencies to 53 Hz). The new hardware and the extensive use of computer-experiment integration developed for this test are described. The geometrical configuration of the model and the test arrangement are described in detail. Mean and first harmonic data are presented in both tabular and graphical form to aid in comparisons with other data and with numerical computations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81221 , A-8294
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  • 72
    Publication Date: 2013-08-31
    Description: Automatic flare and decrab control laws for conventional takeoff and landing aircraft were adapted to the unique requirements of the powered lift short takeoff and landing airplane. Three longitudinal autoland control laws were developed. Direct lift and direct drag control were used in the longitudinal axis. A fast time simulation was used for the control law synthesis, with emphasis on stochastic performance prediction and evaluation. Good correlation with flight test results was obtained.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152365
    Format: application/pdf
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  • 73
    Publication Date: 2013-08-31
    Description: Theodorsen's circulation function relates lift to downwash in unsteady two dimensional incompressible flow. A continued fraction representation for the circulation function is described. The continued fraction converges and has a particularly simple coefficient pattern.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81838
    Format: application/pdf
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  • 74
    Publication Date: 2013-08-31
    Description: An investigation was conducted in the Langely 6 by 28 inch transonic tunnel to determine the two dimensional aerodynamic characteristics of three helicopter rotor airfoils at Reynolds numbers from typical model scale to full scale at Mach numbers from about 0.35 to 0.90. The model scale Reynolds numbers ranged from about 700,00 to 1,500,000 and the full scale Reynolds numbers ranged from about 3,000,000 to 6,600,000. The airfoils tested were the NACA 0012 (0 deg Tab), the SC 1095 R8, and the SC 1095. Both the SC 1095 and the SC 1095 R8 airfoils had trailing edge tabs. The results of this investigation indicate that Reynolds number effects can be significant on the maximum normal force coefficient and all drag related parameters; namely, drag at zero normal force, maximum normal force drag ratio, and drag divergence Mach number. The increments in these parameters at a given Mach number owing to the model scale to full scale Reynolds number change are different for each of the airfoils.
    Keywords: AERODYNAMICS
    Type: AVRADCOM-TR-80-B-5 , NASA-TP-1701 , L-13139
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  • 75
    Publication Date: 2013-08-31
    Description: Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin g tunnel are presented in plotted form for a 1/6 scale, single engine, high wing, general aviation model. The configurations tested included the basic airplane and control deflections, wing leading edge devices, tail designs, and airplane components. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg and clockwise and counter clockwise rotations covering a spin coefficient range from 0 to 0.9.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3201
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  • 76
    Publication Date: 2013-08-31
    Description: An analytical technique for the prediction of fan blade flutter was evaluated by utilizing first stage fan flutter data from tests on an advanced high performance engine. The formulation includes both aerodynamic and mechanical coupling among all the blades of the assembly. Mistuning is accounted for in the analysis so that individual blade inertias, frequencies, or damping can be considered. Airfoil stability was predicted by calculating a flutter determinant, the eigenvalues of which indicate the extent of susceptibility to flutter. When blade to blade differences in frequencies are considered, a stable system is predicted for the test points examined. For a tuned system, it was found that torsional flutter can be predicted at a limited number of interblade phase angles. Examination of these phase angles indicated that they were "close" to the condition of acoustic resonance. For the range of Mach numbers and reduced frequencies considered, the so called subcritical flutter cannot be predicted. The essential influence of mechanical coupling among the blades is to change the frequencies of the system with little or no change in damping; however, aerodynamic coupling together with mechanical coupling could change not only frequencies, but also damping in the system, with a trend toward instability.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165137 , R80-914545-16
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  • 77
    Publication Date: 2013-08-31
    Description: Static and forward speed tests were made in a 40 multiplied by 80 foot wind tunnel of a large-scale, ejector-powered V/STOL aircraft model. Modifications were made to the model following earlier tests primarily to improve longitudinal acceleration capability during transition from hovering to wingborne flight. A rearward deflection of the fuselage augmentor thrust vector was shown to be beneficial in this regard. Other augmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also showed negligible influence on the performance of the wing and of the fuselage augmentor.
    Keywords: AERODYNAMICS
    Type: DHC-DND-80-1 , NASA-CR-163578
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  • 78
    Publication Date: 2013-08-31
    Description: Pressure distributions on a wing body at Mach 4.63 are calculated. The combined theory is shown to give improved predictions over either linear theory or impact theory alone. The combined theory is also applied in the inverse design mode to calculate optimum camber slopes at Mach 4.63. Comparisons with optimum camber slopes obtained from unmodified linear theory show large differences. Analysis of the results indicate that the combined theory correctly predicts the effect of thickness on the loading distributions at high Mach numbers, and that finite thickness wings optimized at high Mach numbers using unmodified linear theory will not achieve the minimum drag characteristics for which they are designed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3314
    Format: application/pdf
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  • 79
    Publication Date: 2013-08-31
    Description: The potential advantages of bank-to-turn control are summarized. Recent and current programs actively investigating bank-to-turn steering are reviewed and critical technology areas concerned with bank-to-turn control are assessed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3325
    Format: application/pdf
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  • 80
    Publication Date: 2013-08-31
    Description: The implementation of the approximate factorization algorithm and its ability to efficiently and accurately describe transonic flow about an NACA 64A010 airfoil section is examined. The approximate factorization algorithm is developed from the nondimensional, conservative, vectorized Navier-Stokes equations expressed in curvilinear coordinates. Equations of state and transport coefficient relations appropriate to atmospheric air are appended to close the system of partial differential equations. An algebraic turbulence model is also incorporated into the equation set. This algorithm was verified by investigating the flow about an NACA 64A010 airfoil at 0, 2, and 3.5 deg angle of attack for free-stream conditions of 2,000,000 Reynolds number and 0.8 Mach number. Overall results were in good qualitative agreement with wind tunnel data sets. However, while nondimensional times of six were attained, numerical difficulties prevented any case from reaching a true steady state.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163376
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  • 81
    Publication Date: 2013-08-31
    Description: Test results are presented for a large scale, external augmentor V/STOL model in a 40 ft by 80 ft wind tunnel. The model was powered by a GE J97 engine and featured longitudinal ejectors alongside and external to the fuselage together with an augmentor flap on the low aspect ratio, double-delta wing. A static thrust augmentation ratio of 1.60 was measured for the fuselage augmentor at a nozzle pressure ratio of 3.0 and a nozzle exhaust gas temperature of 700 C. At forward speed the model showed a strong positive lift interference due to the augmentor flap, and a marked absence of negative lift interference due to the fuselage augmentor jet system. The nose-up moment of the fuselage augmentor inlet flow was approximately cancelled by a 60 deg deflection of the augmentor flap. An assessment of the thrust and drag components to allow the prediction of transition performance of aircraft designs based on the present conceptual model was made. Lateral tests showed strong but well ordered effects of power.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152255 , DHC-DND-79-4
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  • 82
    Publication Date: 2013-08-31
    Description: The investigation was carried out using the rotating hot wire technique. Measurements were taken inside the end wall boundary layer to discern the effect of annulus and hub wall boundary layer, secondary flow, and tip leakage on the wake structure. Static pressure gradients across the wake were measured using a static stagnation pressure probe insensitive to flow direction changes. The axial and the tangential velocity defects, the radial component of velocity, and turbulence intensities were found to be very large as compared to the near and far wake regions. The radial velocities in the trailing edge region exhibited characteristics prevalent in a trailing vortex system. Flow near the blade tips found to be highly complex due to interaction of the end wall boundary layers, secondary flows, and tip leakage flow with the wake. The streamwise curvature was found to be appreciable near the blade trailing edge. Flow properties in the trailing edge region are quite different compared to that in the near and far wake regions with respect to their decay characteristics, similarity, etc. Fourier decomposition of the rotor wake revealed that for a normalized wake only the first three coefficients are dominant.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159518 , PSU-TURBO-R-80-4
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  • 83
    Publication Date: 2013-08-31
    Description: Experimental data are presented on the effect of Reynolds number on unsteady pressures induced by the pitching motion of an oscillating airfoil. Scale effects are discussed with reference to a conventional airfoil (NACA 64A010) and a supercritical airfoil (NLR 7301) at mean-flow conditions that support both weak and strong shock waves. During the experiment the Reynolds number was varied from 3,000,000 to 12,000,000 at a Mach number and incidence necessary to induce the required flow. Both fundamental frequency and complete time history data are presented over the range of reduced frequencies that is important in aeroelastic applications. The experimental data show that viscous effects are important in the case of the supercritical airfoil at all flow conditions and in the case of the conventional airfoil under strong shock-wave conditions. Some frequency-dependent viscous effects were also observed.
    Keywords: AERODYNAMICS
    Type: A-8259 , NASA-TM-81216
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  • 84
    Publication Date: 2013-08-31
    Description: An experimental investigation was conducted to determine the effect of Reynolds number on the stability characteristics of a body with cruciform wings