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  • AERODYNAMICS  (366)
  • 1980-1984  (366)
  • 1965-1969
  • 1945-1949
  • 1981  (366)
  • 101
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    A potential-flow/boundary-layer method for calculating subsonic and transonic airfoil flow with trailing-edge separation (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The development of a potential-flow/boundary-layer method for calculating subsonic and transonic turbulent flow past airfoils with trailing-edge separation is reported. A moment-of-momentum integral boundary-layer method is used which employs the law-of-the-wall/law-of-the-wake velocity profile and a two-layer eddy-viscosity model and ignores the laminar sublayer. All integrals across the boundary layer are obtained in closed form. Separation is assumed to occur when the shearing-stress velocity vanishes. A closed-form solution is derived for separated-flow regions where the shearing stress is negligible. In the potential-flow method, the exact form of the airfoil boundary condition is used, but it is applied at the chord line rather than the airfoil surface. This allows the accurate computation of flow about airfoils at large angles of attack but permits the use of body-oriented Cartesian computational grids. The governing equation for the perturbation velocity potential contains several terms in addition to the classical small-disturbance terms.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81850 , L-14255
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  • 102
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    A procedure for designing forebodies with constraints on cross-section shape and axial area distribution (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method is described for designing a forebody with cross sections which vary smoothly from an initial prescribed nose shape to a different prescribed base shape in such a way that the cross-section areas conform to a preassigned axial area distribution. It is shown that these conditions can be satisfied with a remaining degree of freedon, which can be used to accomplish a modest amount of geometric or pressure tailoring of the forebody. An example is provided which involves modifying the pressure distribution along a given meridian line of the forebody.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1881 , L-14516
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  • 103
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    TAIR: A transonic airfoil analysis computer code (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81296 , A-8594
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  • 104
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    Simple approximations for the asymptotic description of the interaction between a normal shock wave and a turbulent boundary layer at transonic speeds (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The asymptotic description of the interaction between a normal shock wave and a turbulent boundary layer is reviewed. The layers necessary in a rational analysis of the interaction are discussed with emphasis on the differences from an interaction with a laminar boundary layer, the uncoupling of solutions for the distribution of pressure and skin friction at the wall, and the role of the Reynolds shear stress in these solutions. The accuracy of asymptotic solutions in flows at Reynolds numbers of technical interest is discussed. Solutions for the distribution of pressure and skin friction at the wall and the shape of the shock are considered for the case where the flow is near separation. For the pressure and skin friction, it is possible to write two simplified partial solutions, one valid at the beginning of the interaction and one valid somewhat downstream of the shock wave. A solution composed of these two parts and a linear interpolation between them appears to give good comparison with experiment; one unknown constant, independent of the parameters of the interaction, must be found from experiment. The simplified relations are presented. Comparison of numerical computations with experimental data indicates a possible value for the constant and shows quite satisfactory results.
    Keywords: AERODYNAMICS
    Type: AGARD Computation of Viscous-Inviscid Interactions; 14 p
    Format: text
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  • 105
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    Icing tunnel tests of a glycol-exuding porous leading edge ice protection system on a general aviation airfoil (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A glycol-exuding porous leading edge ice protection system was tested. Results show that the system is very effective in preventing ice accretion (anti-ice mode) or removing ice from an airfoil. Minimum glycol flow rates required for anti-icing are a function of velocity, liquid water content in the air, ambient temperature, and droplet size. Large ice caps were removed in only a few minutes using anti-ice flow rates. It was found that the shed time is a function of the type of ice, size of the ice cap, angle of attack, and glycol flow rate. Wake survey measurements show that there is no significant drag penalty for the installation or operation of the system tested.
    Keywords: AERODYNAMICS
    Type: NASA-CR-164377 , KU-FRL-464-1
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  • 106
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    Experimental and analytical study of the longitudinal aerodynamic characteristics of analytically and empirically designed Strake-wing configurations at subcritical speeds (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Sixteen analytically and empirically designed strakes have been tested experimentally on a wing-body at three subcritical speeds in such a way as to isolate the strake-forebody loads from the wing-afterbody loads. Analytical estimates for these longitudinal results are made using the suction analogy and the augmented vortex lift concepts. The synergistic data are reasonably well estimated or bracketed by the high- and low-angle-of-attack vortex lift theories over the Mach number range and up to maximum lift or strake-vortex breakdown over the wing. Also, the strake geometry is very important in the maximum lift value generated and the lift efficiency of a given additional area. Increasing size and slenderness ratios are important is generating lift efficiently, but similar efficiency can also be achieved by designing a strake with approximately half the area of the largest gothic strake tested. These results correlate well with strake-vortex-breakdown observations in the water tunnel.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1803 , L-14041
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  • 107
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    Comparison of selected lift and sideslip characteristics of the Ayres Thrush S2R-800, winglets off and winglets on, to full-scale wind-tunnel data (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: All calculations were done in the stability axes system. The winglets used were constructed of modified GA(w)-2 airfoils. Aerodynamic characteristics discussed include: angle of attack; lift-curve slope; side force; yawing moments; rolling moments.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165710 , KU-FRL-399-3
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  • 108
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    Aerodynamics, aeroelasticity, and stability of hang gliders. Experimental results (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: One-fifth-scale models of three basic ultralight glider designs were constructed to simulate the elastic properties of full scale gliders and were tested at Reynolds numbers close to full scale values. Twenty-four minor modifications were made to the basic configurations in order to evaluate the effects of twist, reflex, dihedral, and various stability enhancement devices. Longitudinal and lateral data were obtained at several speeds through an angle of attack range of -30 deg to +45 deg with sideslip angles of up to 20 deg. The importance of vertical center of gravity displacement is discussed. Lateral data indicate that effective dihedral is lost at low angles of attack for nearly all of the configurations tested. Drag data suggest that lift-dependent viscous drag is a large part of the glider's total drag as is expected for thin, cambered sections at these relatively low Reynolds numbers.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81269
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  • 109
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    Design and Experimental Results for a Natural-Laminar-Flow Airfoil for General Aviation Applications (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A natural-laminar-flow airfoil for general aviation applications, the NLF(1)-0416, was designed and analyzed theoretically and verified experimentally in the Langley Low-Turbulence Pressure Tunnel. The basic objective of combining the high maximum lift of the NASA low-speed airfoils with the low cruise drag of the NACA 6-series airfoils was achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge was also met. Comparisons of the theoretical and experimental results show excellent agreement. Comparisons with other airfoils, both laminar flow and turbulent flow, confirm the achievement of the basic objective.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1861 , L-14117
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  • 110
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    Analysis of turbulent flow about an isolated airfoil using a time dependent Navier-Stokes procedure (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The procedure solves the Navier-Stokes equations by the consistently split linearized block implicit method of Briley and McDonald in a body fitted coordinate system. The procedure is described and results are presented for flow about an airfoil whose incidence changes from 6 degrees to 19 degrees at a Reynolds number of one million and Mach number of 0.2. In addition, the unsteady flow about an airfoil held at a constant 19 degree incidence is examined and compared to data.
    Keywords: AERODYNAMICS
    Type: AGARD Boundary Layer Effects on Unsteady Airfoils; 14 p
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  • 111
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    An investigation of drag reduction for a standard truck with various modifications (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A wind tunnel investigation was conducted to determine the influence of several physical variables on the aerodynamic drag of a standard truck model. The physical variables included: a cab mounted air deflector; a boattail on the rear of the cargo compartment; flow-vanes on the front of the cargo compartment; and a forebody fairing over the cab. Tests were conducted at yaw angles (relative wind angle) of 0, 5, 10, 20, and 30 degrees and Reynolds numbers of 3.4 x 100,000 to 6.1 x 100,000 based upon the equivalent diameter of the vehicles. The forebody fairing and the flow-vane with the closed bottom were very effective in improving the flow over the forward part of the cargo compartment. The forebody fairing provided a calculated fuel saving of 5.6 liters per hour (1.5 gallons per hour) over the baseline configuration for a ground speed of 88.6 km/hr (55 mph) in national average winds.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163107 , KU-FRL-406-2
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  • 112
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    Prediction of the turbulent wake with second-order closure (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A turbulence was envisioned whose energy containing scales would be Gaussian in the absence of inhomogeneity, gravity, etc. An equation was constructed for a function equivalent to the probability density, the second moment of which corresponded to the accepted modeled form of the Reynolds stress equation. The third moment equations obtained from this were simplified by the assumption of weak inhomogeneity. Calculations are presented with this model as well as interpretations of the results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-164294 , FDA-81-03
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  • 113
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    Full-scale aerodynamic characteristics of a propellar installed on a small twin-engine aircraft wing panel (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Full-scale measurements of shaft thrust and torque were made. Wind-tunnel speeds and blade angles were set for full-scale flight conditions. Excellent quality measurements were obtained of the thrust coefficient, the power coefficient, and the propeller efficiency for various values of the advance ratio and the blade incidence angle at 3/4-blade radius. A conventional propeller theory found in the literature was applied to the present results. Although thrust, power, and efficiency were somewhat overpredicted, the advance ratio for maximum efficiency was predicted quite accurately. It was found that, for some conditions, spinner drag could be significant. A simple correction that was based on the spinner base pressure substantially accounted for the changes in efficiency that resulted from this cause.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81285 , A-8478
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  • 114
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    Two-dimensional aerodynamic characteristics of the NACA 0012 airfoil in the Langley 8 foot transonic pressure tunnel (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Data are presented for lift coefficients from near zero through maximum values at Mach numbers from 0.30 to 0.86 and Reynolds numbers of 3.0 x 10 to the sixth power with transition fixed. A limited amount of data is presented near zero and maximum lift for a Reynolds number of 6.0 x 10 to the sixth power with transition fixed. In addition, transition free data is presented through the Mach number range from 0.30 to 0.86 for near zero lift and a Reynolds number of 3.0 x 10 to the sixth power.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81927
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  • 115
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    Topology of three-dimensional separated flows (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Based on the hypothesis that patterns of skin-friction lines and external streamlines reflect the properties of continuous vector fields, topology rules define a small number of singular points (nodes, saddle points, and foci) that characterize the patterns on the surface and on particular projections of the flow (e.g., the crossflow plane). The restricted number of singular points and the rules that they obey are considered as an organizing principle whose finite number of elements can be combined in various ways to connect together the properties common to all steady three dimensional viscous flows. Introduction of a distinction between local and global properties of the flow resolves an ambiguity in the proper definition of a three dimensional separated flow. Adoption of the notions of topological structure, structural stability, and bifurcation provides a framework to describe how three dimensional separated flows originate and succeed each other as the relevant parameters of the problem are varied.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81294 , A-8554
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  • 116
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    An asymptotic unsteady lifting-line theory with energetics and optimum motion of thrust-producing lifting surfaces (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A low frequency unsteady lifting-line theory is developed for a harmonically oscillating wing of large aspect ratio. The wing is assumed to be chordwise rigid but completely flexible in the span direction. The theory is developed by use of the method of matched asymptotic expansions which reduces the problem from a singular integral equation to quadrature. The wing displacements are prescribed and the pressure field, airloads, and unsteady induced downwash are obtained in closed form. The influence of reduced frequency, aspect ratio, planform shape, and mode of oscillation on wing aerodynamics is demonstrated through numerical examples. Compared with lifting-surface theory, computation time is reduced significantly. Using the present theory, the energetic quantities associated with the propulsive performance of a finite wing oscillating in combined pitch and heave are obtained in closed form. Numerical examples are presented for an elliptic wing.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165679
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  • 117
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    Improved method for calculating transonic velocities on blade-to-blade stream surfaces of a turbomachine (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method was developed to improve the accuracy of an existing computer program used to calculate transonic velocities on a blade-to-blade surface of a turbomachine. The method eliminates problems encountered in obtaining solutions with the velocity gradient equation when large gradients in velocity occur through the blade row. With the improved method, results indicate that the transonic solution can be obtained by scaling the velocities obtained at the reduced mass flow rate where all velocities are subsonic thereby eliminating the need for a solution of the velocity gradient equation. Solutions obtained with the scaling method on a two dimensional compressor cascade and an axial turbine stator show good agreement with experimental data. The results obtained for the stationary blade rows and comparison of analytical results obtained with and without the present method suggest that the method will yield an improved solution for centrifugal compressor impellers.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1772 , E-128
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  • 118
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    Computation of wing-vortex interaction in transonic flow using implicit finite difference algorithm (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An implicit delta form finite difference algorithm for Euler equations in conservation law form was used in preliminary calculations of three dimensional wing vortex interaction. Both steady and unsteady transonic flow wing vortex interactions are computed. The computations themselves are meant to guide upcoming wind tunnel experiments of the same flow field. Various modifications to the numerical method that are intended to improve computational efficiency are also described and tested in both two and three dimensions. Combination of these methods can reduce the overall computational time by a factor of 4.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166251 , NAS 1.26:166251
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  • 119
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    The aerodynamics of circulation control (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two dimensional subsonic wind tunnel tests were conducted on a 20% thickness: chord ratio circulation controlled elliptic aerofoil section equipped with forward and reverse blowing slots. Overall performance measurements were made over a range of trailing edge blowing momentum coefficients from 0 to 0.04; some included the effect of leading edge blowing. A detailed investigation of the trailing edge wall jet, using split film probes, hot wire probes and total head tubes, provided measurements of mean velocity components, Reynolds normal and shear stresses, and radial static pressure. The closure of the two dimensional angular momentum and continuity equations was examined using the measured data, with and without correction, and the difficulty of obtaining a satisfactory solution illustrated. Suggestions regarding the nature of the flow field which should aid the understanding of Coanda effect and the theoretical solution of highly curved wall jet flows are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-168662 , NAS 1.26:168662 , SU-JIAA-TR-41
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  • 120
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    An experimental investigation of the flow past a finite circular cylinder at a low subcritical Reynolds number (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results of hot wire measurements made in the near wake at a Reynolds number of 9955 are reported. The measurements include the mean velocity profiles, root mean square values of the velocity fluctuations, frequency spectra, and velocity cross correlations. The mean velocity profiles were used to determine the wake width, whose variation in the downstream and spanwise directions was examined. It is observed that close to the cylinder, the wake is narrower toward the free end than it is away from it, while further downstream the wake is wider toward the tip than it is away from it. It is found that the flow over the span can be characterized by four regions: a tip region where vortex shedding occurs at a lower frequency than that prevalent for away from the tip; an intermediate region adjacent to the first one where a frequency component of a nonshedding character is present; a third region characterized by a gradually increasing shedding frequency with increasing distance from the tip; and a two dimensional region where the shedding frequency is constant.
    Keywords: AERODYNAMICS
    Type: NASA-CR-168661 , NAS 1.26:168661 , SU-JIAA-TR-40
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  • 121
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    Low-speed aerodynamic performance of an aspect-ratio-10 supercritical-wing transport model equipped with a full-span slat and part-span and full-span double-slotted flaps (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 4 by 7 Meter Tunnel to determine the static longitudinal and lateral directional aerodynamic characteristics of an advanced aspect ratio 10 supercritical wing transport model equipped with a full span leading edge slat as well as part span and full span trailing edge flaps. This wide body transport model was also equipped with spoiler and aileron roll control surfaces, flow through nacelles, landing gear, and movable horizontal tails. Six basic wing configurations were tested: (1) cruise (slats and flaps nested), (2) climb (slats deflected and flaps nested), (3) part span flap, (4) full span flap, (5) full span flap with low speed ailerons, and (6) full span flap with high speed ailerons. Each of the four flapped wing configurations was tested with leading edge slat and trailing edge flaps deflected to settings representative of both take off and landing conditions. Tests were conducted at free stream conditions corresponding to Reynolds number of 0.97 to 1.63 x 10 to the 6th power and corresponding Mach numbers of 0.12 to 0.20, through an angle of attack range of 4 to 24, and a sideslip angle range of -10 deg to 5 deg. The part and full span wing configurations were also tested in ground proximity.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1805 , L-13825 , NAS 1.60:1805
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  • 122
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    Calculation of lateral-directional stability derivatives of wings by a nonplanar quasi-vortex-lattice method (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The nonplanar quasi-vortex-lattice method is applied to the calculation of lateral-directional stability derivatives of wings with and without vortex-lift effect. Results for conventional configurations and those with winglets, V-tail, etc. are compared with available data. All rolling moment derivatives are found to be accurately predicted. The prediction of side force and yawing moment derivatives for some configurations is not as accurate. Causes of the discrepancy are discussed. A user's manual for the program and the program listing are also included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165659 , NAS 1.26:165659 , REPT-80-001
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  • 123
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    Experimental investigation of a 10-percent-thick helicopter rotor airfoil section designed with a viscous transonic analysis code (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 6- by 28-Inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics of a 10-percent-thick helicopter rotor airfoil at Mach numbers from 0.33 to 0.87 and respective Reynolds numbers from 4.9 x 10 to the 6th to 9.8 x 10 to the 6th. This airfoil, designated the RC-10(N)-1, was also investigated at Reynolds numbers from 3.0 x 10 to the 6th to 7.3 x 10 to the 6th at respective Mach numbers of 0.33 to 0.83 for comparison wit the SC 1095 (with tab) airfoil. The RC-10(N)-1 airfoil was designed by the use of a viscous transonic analysis code. The results of the investigation indicate that the RC-10(N)-1 airfoil met all the design goals. At a Reynolds number of about 9.4 x 10 to the 6th the drag divergence Mach number at zero normal-force coefficient was 0.815 with a corresponding pitching-moment coefficient of zero. The drag divergence Mach number at a normal-force coefficient of 0.9 and a Reynolds number of about 8.0 x 10 to the 6th was 0.61. The drag divergence Mach number of this new airfoil was higher than that of the SC 1095 airfoil at normal-force coefficients above 0.3. Measurements in the same wind tunnel at comparable Reynolds numbers indicated that the maximum normal-force coefficient of the RC-10(N)-1 airfoil was higher than that of the NACA 0012 airfoil for Mach numbers above about 0.35 and was about the same as that of the SC 1095 airfoil for Mach numbers up to 0.5.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1864 , L-14182 , NAS 1.60:1864 , AVRADCOM-TR-81-B-3
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  • 124
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    Design and experimental results for a flapped natural-laminar-flow airfoil for general aviation applications (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A flapped natural laminar flow airfoil for general aviation applications, the NLF(1)-0215F, has been designed and analyzed theoretically and verified experimentally in the Langley Low Turbulence Pressure Tunnel. The basic objective of combining the high maximum lift of the NASA low speed airfoils with the low cruise drag of the NACA 6 series airfoils has been achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge has also been met. Comparisons of the theoretical and experimental results show generally good agreement.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1865 , L-14409 , NAS 1.60:1865
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  • 125
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    VORCAM: A computer program for calculating vortex lift effect of cambered wings by the suction analogy (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A user's guide to an improved version of Woodward's chord plane aerodynamic panel computer code is presumed. The guide can be applied to cambered wings exhibiting edge separated flow, including those with leading edge vortex flow at subsonic and supersonic speeds. New orientations for the rotated suction force are employed based on the momentum principal. The supersonic suction analogy method is improved by using an effective angle of attack defined through a semiempirical method.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165800 , NAS 1.26:165800 , CRINC-FRL-426-2
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  • 126
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    PAN AIR: A computer program for predicting subsonic or supersonic linear potential flows about arbitrary configurations using a higher order panel method. Volume 1: Theory document (version 1.1) (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Panel aerodynamics (PAN AIR) is a system of computer programs designed to analyze subsonic and supersonic inviscid flows about arbitrary configurations. A panel method is a program which solves a linear partial differential equation by approximating the configuration surface by a set of panels. An overview of the theory of potential flow in general and PAN AIR in particular is given along with detailed mathematical formulations. Fluid dynamics, the Navier-Stokes equation, and the theory of panel methods were also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3251 , NAS 1.26:3251
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  • 127
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    A unified thermal and vertical trajectory model for the prediction of high altitude balloon performance (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer model for the prediction of the trajectory and thermal behavior of zero-pressure high altitude balloon was developed. In accord with flight data, the model permits radiative emission and absorption of the lifting gas and daytime gas temperatures above that of the balloon film. It also includes ballasting, venting, and valving. Predictions obtained with the model are compared with flight data from several flights and newly discovered features are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-156884 , TAMRF-4217-81-01
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  • 128
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    Calculation of three-dimensional, inviscid, supersonic, steady flows (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A detailed description of a computational program for the evaluation of three dimensional supersonic, inviscid, steady flow past airplanes is presented. Emphasis was put on how a powerful, automatic mapping technique is coupled to the fluid mechanical analysis. Each of the three constituents of the analysis (body geometry, mapping technique, and gas dynamical effects) was carefully coded and described. Results of computations based on sample geometrics and discussions are also presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165110 , POLY-M/AE-81-25
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  • 129
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    Analytic investigation of effect of end-wall contouring on stator performance (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Calculations show improved stator performance when the tip end wall was contoured so that the inlet area was greater than the exit area. Comparisons are made with previously published experimental data. The results of a parametric analysis of the effect contour geometry on the efficiency of a highly loaded axial stator are given. The maximum stator efficiency gain is about 0.8 percentage point, and this represents a 22 percent reduction in stator losses. The degree to which endwall contouring reduces the forces driving secondary flows was also examined. The driving forces for both cross channel and radial secondary flow were reduced.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1943 , E-719
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  • 130
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    Helicopter rotor loads using a matched asymptotic expansion technique (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The theoretical basis and computational feasibility of the Van Holten method, and its performance and range of validity by comparison with experiment and other approximate methods was examined. It is found that within the restrictions of incompressible, potential flow and the assumption of small disturbances, the method does lead to a valid description of the flow. However, the method begins to break down under conditions favoring nonlinear effects such as wake distortion and blade/rotor interaction.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165742 , NAS 1.26:165742
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  • 131
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    Aeroelasticity matters: Some reflections on two decades of testing in the NASA Langley transonic dynamics tunnel (1981)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: Testing of wind-tunnel aeroelastic models is a well established, widely used means of studying flutter trends, validating theory and investigating flutter margins of safety of new vehicle designs. The Langley Transonic Dynamics Tunnel was designed specifically for work on dynamics and aeroelastic problems of aircraft and space vehicles. A cross section of aeroelastic research and testing in the facility since it became operational more than two decades ago is presented. Examples selected from a large store of experience illustrate the nature and purpose of some major areas of work performed in the tunnel. These areas include: specialized experimental techniques; development testing of new aircraft and launch vehicle designs; evaluation of proposed "fixes" to solve aeroelastic problems uncovered during development testing; study of unexpected aeroelastic phenomena (i.e., "surprises"); control of aeroelastic effects by active and passive means; and, finally, fundamental research involving measurement of unsteady pressures on oscillating wings and control surface.
    Keywords: AERODYNAMICS
    Type: NASA-TM-83210 , Intern. Symp. on Aeroelasticity; 5-7 Oct. 1980 - 1 Oct. 1981; Nuremberg; Germany
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  • 132
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    Freestream turbulence effects on the development of a rotor wake (1981)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effect of freestream turbulence on the development of a three-dimensional wake of a compressor rotor blade was studied experimentally. The turbulence level at the inlet of a rotor was varied systematically using grids upstream of the rotor. The rotor wake was measured with inlet turbulence intensities of 0.5, 3, and 5%. The experimental results indicate that the maxium change in the mean velocity defect is 4% over the range of inlet turbulence levels employed, while the turbulence structure in the wake is altered more substantially. The freestream turbulence effect was also analyzed, numerically, using the modified Reynolds stress closure model. The comparison between numerical prediction and experimental data shows that the freestream turbulence effect can be represented successfully with the turbulence closure model employed in this paper.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1431
    Format: text
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  • 133
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    The challenge of unsteady separating flows (1981)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: It is noted that the general fluid dynamic problem of unsteady separation at most practical Reynolds numbers remains an unsolved one and that no completely reliable prediction techniques exist at the present time. The modern design engineer must therefore draw from a combination of approximate theories, empirical correlations of data, and finite difference programs based on uncertain physical modeling of turbulence. An attempt is made to describe the basic features of several representative classes of problems for which unsteady effects produce strong or unusual changes in the separation characteristics of the flow. The analysis concerns itself largely with external flow, and emphasis is placed on the physical phenomena involved.
    Keywords: AERODYNAMICS
    Type: AD-A102378 , American Society of Civil Engineers; Journal
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  • 134
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    Boundary conditions for subsonic compressible Navier-Stokes calculations (1981)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A systematic study of inflow and outflow boundary conditions for the numerical solution of the compressible Navier-Stokes equations is presented. Combinations of several representative inflow and outflow boundary conditions are applied in the solution of subsonic flow over a flat plate in a finite computational domain. These boundary conditions are evaluated in terms of their effect on the accuracy of the solution and the rate of convergence to a steady state. It is shown that errors in the data specified at the inflow boundary can produce significant errors in the computed flow field. It is also shown that a non-reflecting outflow boundary condition can significantly reduce the total computational time required.
    Keywords: AERODYNAMICS
    Type: Computers and Fluids; 9; Sept
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  • 135
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    Prediction of blade-vortex interaction noise from measured blade pressure (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The impulsive nature of noise due to the interaction of a rotor blade with a tip vortex is studied. The time signature of this noise is calculated theoretically based on the measured blade surface pressure fluctuation of an operational load survey rotor in slow descending flight and is compared with the simultaneous microphone measurement. Particularly, the physical understanding of the characteristic features of a waveform is extensively studied in order to understand the generating mechanism and to identify the important parameters. The interaction trajectory of a tip vortex on an acoustic planform is shown to be a very important parameter for the impulsive shape of the noise. The unsteady nature of the pressure distribution at the very leading edge is also important to the pulse shape. The theoretical model using noncompact liner acoustics predicts the general shape of interaction impulse pretty well except for peak amplitude which requires more continuous information along the span at the leading edge.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81320 , A-8692 , USAAVRADCOM-TR-81-A-24 , PAPER-32 , European Rotorcraft and Powered Lift Aircraft Forum; Moffett Field, CA; United States
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  • 136
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    Experimental and theoretical study of three interacting, closely-spaced, sharp-edged 60 deg delta wings at low speeds (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wind tunnel tests were conducted to determine the subsonic longitudinal aerodynamic characteristics of lifting configuration consisting of a 60 deg delta main wing with two smaller 60 deg delta wings (called sub-wings) attached underneath. The test was designed to determine the effects on lift, drag, and pitching moment due to various placement of the subwings in relation to the main wing. Test results indicate the increasing vertical separation between the main wing and the sub-wings produced the most significant results; a 23.1% increase in maximum lift coefficient, a reduction in drag coefficient at high lift coefficients, and an increase in longitudinal stability. Lateral separation of the sub-wings produced no significant changes. Placement of the sub-wings rearward increases the initial lift curve slope and maximum lift coefficient and also increase the longitudinal stability. Results of a computer study using a vortex lattice code supported the experimental conclusions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3460
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  • 137
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    Analysis of the performance of the drive system and diffuser of the Langley unitary plan wind tunnel (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A broad program was initiated at the Langley Research Center in 1973 to reduce the energy consumption of the laboratory. As a part of this program, the performance characteristics of the Unitary Plan Wind Tunnel were reexamined to determine if potential methods for incresing the operating efficiencies of the tunnel could be formulated. The results of that study are summarized. The performance characteristics of the drive system components and the variable-geometry diffuser system of the tunnel are documented and analyzed. Several potential methods for reducing the energy requirements of the facility are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-83168 , L-14543
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  • 138
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    Static test of a fan-powered chin nozzle for V/STOl applications (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The performance of a "chin" nozzle which diverts flow in a downward direction immediately downstream of a fan typical of designs suitable for V/STOL A applications was evaluated. Back pressure distortion to the fan and fan discharge pressure distortion were also measured. Results show that the distortion is significant at the closest spacing between the fan exit and cascade entrance tested, and that the chin nozzle performance deteriorates with increased flow diversion to the chin nozzle. Color oil flow visualization on video tape and still photos were also obtained. Tests were conducted behind a 12" model fan in the NASA-Lewis fan calibration facility.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165361 , D180-26446-1
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  • 139
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    A comprehensive model to determine the effects of temperature and species fluctuations on reactions in turbulent reacting flows (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The joint probability distribution function (pdf), which is a modification of the bivariate Gaussian pdf, is discussed and results are presented for a global reaction model using the joint pdf. An alternative joint pdf is discussed. A criterion which permits the selection of temperature pdf's in different regions of turbulent, reacting flow fields is developed. Two principal approaches to the determination of reaction rates in computer programs containing detailed chemical kinetics are outlined. These models represent a practical solution to the modeling of species reaction rates in turbulent, reacting flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-164552
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  • 140
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    Prediction of aerodynamic characteristics for wings with extensive separations (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The development of a simple yet effective technique for modelling the effects of trailing edge separation is discussed. The model encloses the low energy region with free vortex sheets coupled with a potential flow panel method. The technique includes an interation cycle between viscous and potential flow routines and its development from the two dimensional case to the three dimensional case is discussed. A description of the potential flow panel method, which is based on an internal Dirichlet boundary condition, is included.
    Keywords: AERODYNAMICS
    Type: AGARD Computation of Viscous-Inviscid Interactions; 15 p
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  • 141
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    Solution of complex nonlinear problems by a generalized application of the method of base and comparison solutions with applications to aerodynamics problems (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A theory for obtaining approximate solutions to nonlinear problems whose exact solutions require the use of large computational procedures is described. The technique represents in some respects a generalization of the method of base and comparison solutions for flows depending on a parameter. For the generalized problem, the input variable is no longer a parameter but a function that is incremented over its entire domain. After performing calculations for a base configuration and a small number of variations of it, solutions for a large class of configurations can be obtained by forming linear combinations of the solution increments. For a restricted class of problems, approximate solutions can be obtained for general variations of a base configuration by using a function-space derivative estimate obtained from a base solution and a single variation.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1857 , L-14197
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  • 142
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    Numerical simulation of steady supersonic flow (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A noniterative, implicit, space-marching, finite-difference algorithm was developed for the steady thin-layer Navier-Stokes equations in conservation-law form. The numerical algorithm is applicable to steady supersonic viscous flow over bodies of arbitrary shape. In addition, the same code can be used to compute supersonic inviscid flow or three-dimensional boundary layers. Computed results from two-dimensional and three-dimensional versions of the numerical algorithm are in good agreement with those obtained from more costly time-marching techniques.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1749 , A-7923
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  • 143
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    A study of a VTOL thrusting ejector in low speed flight, volume 1 (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Low speed aerodynamic characteristics of a thrust augmenter wing suitable for vertical operation were investigated. Wind tunnel test results on the ejector and a similar configuration with a blown flap are analyzed. The configurations represented a VTOL concept at conditions of thrust deflections required for low forward speed flight. The model tested had an unswept untapered wing. Specific data included normal longitudinal forces and monents, surface pressures, ejector exit surveys, and flow field surveys behind the wing.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166137-VOL-1 , NR80H-102-VOL-1
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  • 144
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    Charts for determining potential minimum sonic-boom overpressures for supersonic cruise aircraft (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Charts which give an estimation of minimum achievable sonic-boom levels for supersonic cruise aircraft are presented. A minimization method based on modified linear theory was analyzed. Results show several combinations of Mach number, altitude, and aircraft length and weight. Overpressure and impulse values are given for two types of sonic boom signatures for each of these conditions: (1) a flat top or minimum overpressure signature which has a pressure plateau behind the initial shock, and (2) a minimum shock signature which allows a pressure rise after the initial shock. Results are given for the effects of nose shape.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1820 , L-14190
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  • 145
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    An extension of the local momentum theory to a distorted wake model of a hovering rotor (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The local momentum theory is based on the instantaneous balance between the fluid momentum and the blade elemental lift at a local station in the rotor rotational plane. Therefore, the theory has the capability of evaluating time wise variations of air loading and induced velocity distributions along a helicopter blade span. Unlike a complex vortex theory, this theory was developed to analyze the instantaneous induced velocity distribution effectively. The boundaries of this theory and a computer program using this theory are discussed. A concept introduced into the theory is the effect of the rotor wake contraction in hovering flight. A comparison of this extended local momentum theory with a prescribed wake vortex theory is also presented. The results indicate that the extended local momentum theory has the capability of achieving a level of accuracy similar to that of the prescribed wake vortex theory over wide range variations of rotor geometrical parameters. It is also shown that the analytical results obtained using either theory are in reasonable agreement with experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81258 , A-8436
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  • 146
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    Total main rotor isolation system analysis (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Requirements, preliminary design, and verification procedures for a total main rotor isolation system at n/rev are presented. The fuselage is isolated from the vibration inducing main rotor at one frequency in all degrees of freedom by four antiresonant isolation units. Effects of parametric variations on isolation system performance are evaluated.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165666 , D-210-11788-1
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  • 147
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    Experimental investigations on the V/STOL tunnel at NASA/Langley Research Center (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The flow characteristics of the V/STOL tunnel were investigated. The results show an interaction between tunnel components. The flow around the tunnel circuit gradually deteriorated with increasing distance from the testing area. The flow in the first diffuser was still satisfactory at the beginning of the circuit, while at the end of the circuit, the flow approaching the contraction became entirely unsatisfactory. Deterioration of flow was due largely to turning the stream around the corners, with the resulting flow distortion affecting the diffusers downstream. The large end of the last diffuser stalled on one side and nearly stalled the flow at the tip of the fan. It was found that these adverse flow characteristics reduce the flow quality and the efficiency of the tunnel.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165655
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  • 148
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    Exploratory subsonic investigation of vortex-flap concept on arrow wing configuration (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The drag reduction potential of a vortex flap concept, utilizing the thrust contribution of separation vortices maintained over leading edge flap surfaces, was explored in subsonic wind tunnel tests on a highly swept arrow wing configuration. Several flap geometries were tested in comparison with a previous study on the same model with leading edges drooped for attached flow. The most promising vortex flap arrangements produced drag reductions comparable with leading edge droop over a range of lift coefficients from 0.3 to 0.6 (untrimmed), and also indicated beneficial effects in the longitudinal and lateral static stability characteristics.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Supersonic Cruise Res. 1979, Pt. 1; p 117-129
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  • 149
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    Some aspects of calculating flows about three-dimensional subsonic inlets (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Various three dimensional inlet models were calculated based on the potential flow model. Results are presented in the forms of surface static pressure, flow angularity, surface flow pattern, and inlet flow field. It is indicated that the extension of the lower lip can reduce the adverse pressure gradient and increase the flow separation bound.
    Keywords: AERODYNAMICS
    Type: NASA-TM-82789 , E-941 , NAS 1.15:82789
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  • 150
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    Some observations of surface pressures and the near wake of a blunt trailing edge airfoil (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experiments with a truncated and untruncated airfoils of profiles NACA 640A10, were carried out in subsonic wind tunnels in a velocity range of 19m/s to 54m/s corresponding to Reynolds numbers of 200,000 to 468,000 based on the chord. Airfoil spanned the test section to achieve two dimensionality of the model. Velocity measurements, pressure measurements, and vortex shedding in the wake were measured using a hotwire and pressure transducers. The measured chordwise static pressure distribution on the smooth trailing edge airfoil along the midspan plane, agreed with the theoretical results calculated on the basis of the potential flow for that airfoil. Boundary layer profiles measured in the midspan plane, behind the maximum thickness of the airfoil show no separation of the flow. Spanwise distribution of the measured static pressure on the upper surface of the airfoil shows uniformity for both configurations with and without the boundary layer trip. This uniformity of pressure distribution and separation indicates that the flow on the airfoil was uniform and two dimensional in character.
    Keywords: AERODYNAMICS
    Type: NASA-CR-168563 , SU-JIAA-TR-39
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  • 151
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    An experimental investigation of gapwise periodicity and unsteady aerodynamic response in an oscillating cascade. Volume 2: Data report. Part 2: Mode 2 data (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Computer data are provided for tests conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165457-VOL-2-PT-2
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  • 152
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    An experimental investigation of gapwise periodicity and unsteady aerodynamic response in an oscillating cascade. Volume 2: Data report. Part 1: Text and mode 1 data (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests were conducted a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blade along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge. The tests were conducted for all 96 combinations 2 mean camberline incidence angles 2 pitching amplitudes 3 reduced frequencies and 8 interblade phase angles. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and particularly, the aerodynamic damping coefficient. Data obtained during the test program, reproduced from the printout of the data reduction program are complied. A further description of the contents of this report is found in the text that follows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165457-VOL-2-PT-1 , R81-914618-28-VOL-2-PT-1
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  • 153
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    Experimental trim drag values for conventional and supercritical wings (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Supercritical wings were studied to determine whether they incur higher trim drag values at cruise conditions than wide body technology wings. Relative trim drag increments were measured in an experimental wind tunnel investigation. The tests utilized high aspect ratio supercritical wing and a wide body wing in conjunction with five different horizontal tail configurations, mounted on a representative wide body fuselage. The three low tail configurations and two T tail configurations were chosen to measure the effects on horizontal tail size, location, and camber on the trim drag increments for the two wings. The increase in performance (lift to drag ratio) for supercritical wing over the wide body wing was 11 percent for both the optimum low tail and T tail configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85345
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  • 154
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    Flexiwall 3 SO: A second order predictive strategy for rapid wall adjustment in two-dimensional compressible flow (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An improvement is presented for the 2-D strategies for adjustment of the flexible top and bottom walls of an Adaptive (Wind Tunnel) Wall Test Section (AWTS). This adjustment is part of the wall adaptation process to eliminate top and bottom wall interference at the source. The improvements to account for second order effects are described in mathematical detail. It is intended that these improvements should further minimize the necessary iterations in the wall adaptation process. An associated computer program written in BASIC is presented and several test cases run with this program are discussed. The strategy performs well for a theoretical test case but when applied to experimental AWTS data some discrepancies in the adapted wall shapes are found.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181662 , NAS 1.26:181662
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  • 155
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    Effects of free-stream turbulence on diffuser performance (1981)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental evaluation of the effects of free-stream turbulence on the performance of a subsonic two-dimensional diffuser has been made. Increases of the diffuser's static pressure recovery coefficient of 11.3 and 23.9 percent at total included 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 7.2 times larger than the inlet boundary layer displacement thickness, when the inlet total free-stream turbulence intensity was at least 3.5 percent, and when the axes of upstream rods used to generate turbulence were perpendicular to the flow and parallel to the diverging walls of the diffuser. It is hypothesized that a larger scale of turbulence with the specified eddy axis orientation transmits the free-stream energy to the walls more effectively and, when coupled with large turbulence intensities, are mechanisms which act to decrease the distortion and delay separation within the diffuser.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 81-FE-4
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  • 156
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    Theoretical and experimental supersonic lateral-directional stability characteristics (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: For abstract, see A81-37375.
    Keywords: AERODYNAMICS
    Type: AD-A111768 , NASA-TM-85236 , NAS 1.15:85236
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  • 157
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    NASA research related to sailplane airfoils (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The theoretical methods and experimental facilities at the NASA Langley Research Center have been employed to conduct investigations of sailplane airfoils. The unique and powerful capabilities of the Eppler Program have been used to design and analyze many airfoils and to smooth several Wortmann airfoils. Wind-tunnel investigations of two sailplane airfoils have been conducted in the Langley low-turbulence pressure tunnel. A procedure for sailplane performance improvement has been outlined.
    Keywords: AERODYNAMICS
    Type: National Convention; Jan 14, 1981 - Jan 18, 1981; Phoenix, AZ
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  • 158
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    Treatment of conical and nonconical supersonic flows by an implicit marching scheme applied to the full potential equation (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: An aerodynamic prediction technique based on the full potential equation in conservation form is developed for the treatment of supersonic flows. This technique bridges the gap between simplistic linear theory methods and complex Euler solvers. A novel local density linearization concept and a second order accurate retarded density scheme, both producing the correct artificial viscosity, are introduced in developing an implicit marching scheme for solving the scalar potential. Results for conical flows over delta wings and a wing-body combination and for non-conical flows over bodies of revolution at angles of attack are compared with Euler and nonconservative full potential calculations and experimental data. The present formulation requires an order of magnitude less computer time and significantly less computer memory over Euler codes and exhibits a considerable improvement in computational efficiency and generality over an existing nonconservative full potential code.
    Keywords: AERODYNAMICS
    Type: Computers in flow predictions and fluid dynamics experiments; Nov 15, 1981 - Nov 20, 1981; Washington, DC
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  • 159
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    Experimental and numerical studies of the incompressible viscous flow over a two-dimensional airfoil (1981)
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    Publication Date: 2019-07-13
    Description: Physical and numerical experiments for the low Reynolds number flow over a two-dimensional NACA 66(3)-018 airfoil have been performed. Pressure distributions and smoke flow photographs have been obtained for a Reynolds number based on airfoil chord and free-stream conditions of approximately 40,000 at angles of attack of 0 and 6 deg, and for a Reynolds number based on airfoil chord and free-stream conditions of approximately 400,000 at angles of attack of 0 and 12 deg in a low turbulence wind tunnel. Finite difference numerical experiments, using an approximate factorization method, have been obtained for a Reynolds number of 40,000 at angles of attack of 0 and 6 deg. Although the comparison of the wind tunnel and computer results is encouraging, further studies of this type are clearly necessary.
    Keywords: AERODYNAMICS
    Type: Computers in flow predictions and fluid dynamics experiments; Nov 15, 1981 - Nov 20, 1981; Washington, DC
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  • 160
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    Numerical simulation of reentry flow around the Shuttle Orbiter including real gas effects (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The flight environment of a reentry vehicle is predicted from the numerical solution of fluid dynamics equations for the region between the blunt nose and the swept wings. The inviscid portion of the shock layer is modeled by the Euler equations, but the laminar viscous flow adjacent to the wall is modeled by the approximate parabolic Navier-Stokes equations. The approximations made to the axial gradients of pressure and diffusive fluxes enable the coupled inviscid and viscous equations to be solved efficiently along the body axis. The equilibrium air aftbody code contains significant improvements over its predecessor which only considers a perfect gas model and noncircular configurations. The inclusion of numerical damping either explicitly or implicitly has extended its capabilities for predicting flow field around a winged configuration at higher Machs and greater angles of attack. The results are obtained on the cylindrical coordinates and satisfactory for the Shuttle Orbiter at a free-stream Mach number of 22 and an angle of attack of 40 deg. Also discussed are the inviscid formulation and its application for wind-tunnel conditions.
    Keywords: AERODYNAMICS
    Type: Computers in flow predictions and fluid dynamics experiments; Nov 15, 1981 - Nov 20, 1981; Washington, DC
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  • 161
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    Analysis and computation of three-dimensional flow in strongly curved ducts (1981)
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    Publication Date: 2019-07-13
    Description: Numerical solutions are presented for three-dimensional laminar and turbulent flow in curved ducts of rectangular cross section and significant curvature. The analysis is based on a primary-secondary velocity decomposition in a given coordinate system, and leads to approximate governing equations which correct an a priori inviscid solution for viscous effects, secondary flows, total pressure distortion, heat transfer, and internal flow blockage and losses. Solution of the correction equations is accomplished as an initial-value problem in space using an implicit forward-marching technique. The overall solution procedure requires significantly less computational effort than Navier-Stokes algorithms. The present solution procedure is effective even with the extreme local mesh resolution which is necessary to resolve near-wall sublayer regions in turbulent flow calculations. Computed solutions for both laminar and turbulent flow compare very favorably with available analytical and experimental results.
    Keywords: AERODYNAMICS
    Type: Computers in flow predictions and fluid dynamics experiments; Nov 15, 1981 - Nov 20, 1981; Washington, DC
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  • 162
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    Multigrid simulation of asymptotic curved-duct flows using a semi-implicit numerical technique (1981)
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    Publication Date: 2019-07-13
    Description: Asymptotic flows inside curved ducts of rectangular as well as polar cross section are analyzed using the Navier-Stokes equations in terms of the axial velocity and vorticity and the cross-flow stream function. Numerical solutions of the three second-order coupled elliptic partial differential equations governing this flow are obtained efficiently using the coupled alternating-direction implicit (ADI) method as well as the multigrid strongly-implicit (SI) scheme. For the flow configuration studied, the ADI method is found to be more sensitive to the time steps used than is the SI scheme. Use of the multigrid-coupled-strongly-implicit (MG-SI) scheme makes it possible to efficiently obtain fine-grid solutions for configurations having strong secondary flow. It is shown that, for this asymptotic curved-duct flow, the similarity parameter of significance is the Dean's number K rather than the Reynolds number Re. Results are obtained for curved ducts with square cross sections for K up to 900, which here corresponds to Re = 9,000 for this internal flow configuration.
    Keywords: AERODYNAMICS
    Type: Computers in flow predictions and fluid dynamics experiments; Nov 15, 1981 - Nov 20, 1981; Washington, DC
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  • 163
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    Longitudinal afterbody grooves and shoulder radiusing for low-speed bluff body drag reduction (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A new low-speed drag reduction approach is proposed which employs longitudinal surface V-shaped grooves cutting through the afterbody shoulder region. The test Reynolds number range was from 20,000 to 200,000 based on undisturbed free-stream flow and a body diameter of 6.08 cm. The V-grooves are shown to be most effective in reducing drag when the afterbody shoulder radius is zero. Reductions in drag of up to 33% have been measured for this condition. For large shoulder radius, the grooves are only effective at the lower Reynolds numbers of the test.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 81-WA/FE-5 , Winter Annual Meeting; Nov 15, 1981 - Nov 20, 1981; Washington, DC
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  • 164
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    Mixed spectral/finite difference approximations for slightly viscous flows (1981)
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    Publication Date: 2019-07-13
    Description: A numerical scheme is presented, which employs a standard finite difference approximation for the viscous terms in high Reynolds number flows, and resorts to pseudo-spectral methods for the advection terms to greatly reduce the major source of numerical error without dramatically increasing computational cost. The spectral/finite difference (SFD) method evaluates the advection term and second-order differences to evaluate the diffusion term. The fully finite difference (FFD) method with second-order central differences on both terms is also used. The SFD method can handle strong shocks and can outperform the FFD method at moderate viscosity.
    Keywords: AERODYNAMICS
    Type: In: International Conference on Numerical Methods in Fluid Dynamics; Jun 23, 1980 - Jun 27, 1980; tanford; US
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  • 165
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    Direct numerical simulations of turbulent shear flows (1981)
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    Publication Date: 2019-07-13
    Description: Numerical simulations of wakes of axisymmetric bodies and of turbulent mixing layers are reported. The flows were assumed to be statistically homogeneous in the mean flow direction, in concert with experimental data and the self-similarity theorem. The nonlinear Navier-Stokes equations were solved by a pseudo-spectral numerical method using a 32 x 32 x 33 point grid and an algorithm for fast Fourier transforms and inverse transforms. Leapfrog time differencing was employed on nonlinear terms and time differencing on viscous terms. Towed wakes and wakes behind a self-propelled body were simulated, showing that the towed wakes exhibited a proper temporal behavior after an initial period of adjustment, including the development of a kurtosis near the wake edge, which is experimentally verifiable. The mixing-layer simulation displayed the laboratory demonstrated presence of large scale features such as vortex cores, while the lateral coherence was weak.
    Keywords: AERODYNAMICS
    Type: In: International Conference on Numerical Methods in Fluid Dynamics; Jun 23, 1980 - Jun 27, 1980; tanford; US
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  • 166
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    Trends and pacing items in computational aerodynamics (1981)
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    Publication Date: 2019-07-13
    Description: A perspective is presented of trends in computational aerodynamics, and of important technology development items that pace future advanced applications. From a survey of AIAA Journal papers published during the past two decades, the growth trends and the progressively increasing emphasis on code development for viscous, compressible, turbulent flow are illustrated. These trends are reflected in the chronology of introduction by the aerospace industry of new computational methods in aircraft design. Key pacing items outlined are: automatic grid generation for nonlinear inviscid computations; advanced computers, improved efficiency of numerical methods, and improved turbulence models for Reynolds-averaged Navier-Stokes computations; advanced computers, time-dependent three-dimensional law-of-the-wall, code development, improved efficiency of numerical methods, and improved subgrid-scale turbulence modeling for large eddy simulations.
    Keywords: AERODYNAMICS
    Type: In: International Conference on Numerical Methods in Fluid Dynamics; Jun 23, 1980 - Jun 27, 1980; tanford; US
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  • 167
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    International Conference on Numerical Methods in Fluid Dynamics, 7th, Stanford University, Stanford and Moffett Field, CA, June 23-27, 1980, Proceedings (1981)
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    Publication Date: 2019-07-13
    Description: Topics discussed include polygon transformations in fluid mechanics, computation of three-dimensional horseshoe vortex flow using the Navier-Stokes equations, an improved surface velocity method for transonic finite-volume solutions, transonic flow calculations with higher order finite elements, the numerical calculation of transonic axial turbomachinery flows, and the simultaneous solutions of inviscid flow and boundary layer at transonic speeds. Also considered are analytical solutions for the reflection of unsteady shock waves and relevant numerical tests, reformulation of the method of characteristics for multidimensional flows, direct numerical simulations of turbulent shear flows, the stability and separation of freely interacting boundary layers, computational models of convective motions at fluid interfaces, viscous transonic flow over airfoils, and mixed spectral/finite difference approximations for slightly viscous flows.
    Keywords: AERODYNAMICS
    Type: International Conference on Numerical Methods in Fluid Dynamics; Jun 23, 1980 - Jun 27, 1980; tanford; US
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  • 168
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    Flow visualization using a computerized data acquisition system (1981)
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    Publication Date: 2019-07-13
    Description: A computer-driven traversing mechanism combined with mass data storage, data reduction programs, and general-purpose graphics programs permits a visualization of complex flows. A unique seven-hole probe is used which permits reasonably accurate measurements of all average flow properties if the local flow angle does not exceed 80 degrees. A description is given of the wake of a lifting canard surface as this wake passes over a wing. The flow includes concentrated and dissipating vortices, large regions of reduced total pressure, and local flow angles up to 60 deg. All these features can be clearly seen and accurately located in the graphical output.
    Keywords: AERODYNAMICS
    Type: In: International Symposium on Flow Visualization; Sep 09, 1980 - Sep 12, 1980; Bochum
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  • 169
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    An experimental study of separated flow on a finite wing (1981)
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    Publication Date: 2019-07-13
    Description: The flow field associated with the formation of a mushroom shaped trailing edge stall cell on a low-aspect-ratio (AR = 4.0) wing was investigated in a series of low speed wind tunnel tests (Reynolds number based on 15.2 cm chord = 480,000). Flow field surveys of the separation bubble and wake of a partially stalled and fully stalled wing were completed using a hot-wire probe, a split-film probe, and a directional sensitive pressure probe. A new color video display technique was developed to display the flow field survey data. Photographs were obtained of surface oil flow patterns and smoke flow visualization
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-1882 , Atmospheric Flight Mechanics Conference; Aug 19, 1981 - Aug 21, 1981; Albuquerque, NM
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  • 170
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    Interaction of compressor rotor blade wake with wall boundary layer/vortex in the end-wall region (1981)
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    Publication Date: 2019-07-13
    Description: This paper reports the experimental study of the three-dimensional characteristics of the mean velocity of the rotor wake inside the annulus- and hub-wall boundary layers. The measurements were taken with a rotating three-sensor hot wire behind the rotor. This set of measurements probably represents the first set of comprehensive measurements taken inside the annulus- and hub-wall boundary layers. The wake was surveyed at several radial locations inside the boundary layer region and at several axial locations. Interaction of the wake with the annulus-wall boundary layer, secondary flow, tip-leakage flow, and the trailing vortex system results in slower decay and larger width of the wake. The presence of a strong vortex and its merger with the wake is also observed. The end-wall boundary layers and the secondary flow were found to have a substantial effect on both the decay characteristics and the profile of the wake. These and other measurements are reported and interpreted in this paper.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 81-GR/GT-1 , International Symposium on Applications of Fluids Mechanics and Heat Transfer to Energy and Environmental Problems, University of Patras; Jun 29, 1981 - Jul 03, 1981; Patras; Greece
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  • 171
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    Airframe effects on top-mounted inlet systems for VSTOL fighter aircraft (1981)
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    Publication Date: 2019-07-13
    Description: Inlet flow-field and compressor-face performance data were obtained for a 0.095-scale model of a VSTOL fighter-attack aircraft configuration with twin top-mounted inlets. Tests were conducted at Mach numbers from 0.6 to 2.0 and at angles of attack and sideslip up to 27 deg and 12 deg, respectively. The effects of inlet location, wing leading-edge extension planform area, canopy-dorsal integration, and variable incidence canards were determined. The results show that distortion at the compressor face when maneuvering is relatively low (20% or less) at Mach numbers up to 0.9. However, at Mach numbers of 1.2 and above, maneuverability may be restricted because of high distortion or low pressure recovery (80% or less) or both.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-2631 , V/STOL Conference; Dec 07, 1981 - Dec 09, 1981; Palo Alto, CA
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  • 172
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    Low speed testing of the inlets designed for a tandem-fan V/STOL nacelle (1981)
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    Publication Date: 2019-07-13
    Description: An approximately 0.25 scale model of a tandem fan nacelle, designed for a subsonic V/STOL aircraft, was tested in a Lewis wind tunnel. Model variables included long and short aft inlet cowls and the addition of exterior strakes to the short inlet cowl. Inlet pressure recoveries and distortion were measured at pitch angles to 40 deg and at combinations of pitch and yaw to 30 deg. Airspeeds covered a range to 135 knots (69 m/sec). The short aft inlet with added strakes had the best aerodynamic performance and is considered suitable for the intended V/STOL application.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-2627 , V/STOL Conference; Dec 07, 1981 - Dec 09, 1981; Palo Alto, CA
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  • 173
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    Analysis of data from a wind tunnel investigation of a large-scale model of a highly maneuverable supersonic V/STOL fighter - STOL configuration (1981)
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    Publication Date: 2019-07-13
    Description: The surface pressure and thermal characteristics of a large-scale model of a highly maneuverable supersonic fighter with STOL capability are described. The 7.28 m span model is powered by two J-97 turbojets, operated at 9340 N thrust. It combines upper-surface and spanwise blowing to augment the lift characteristics over a wide angle-of-attack range. The most significant feature of the fighter's flow field is the leading edge vortex that forms at low alphas, grows stronger, and moves inboard as alpha is increased. Upper surface blowing enhanced the lift on the wing in both stalled and unstalled areas significantly, while generating only a modest aft shift in the center of pressure. Lift gains were greatest at high alphas and with the flap deflected. Spanwise blowing was most significant at angles-of-attack greater than 8 deg, when the jet strengthened the vortex. The 1100 F spanwise blowing jet mixed very rapidly with the wing flow field, creating a maximum temperature rise of only 300-350 F. A comparison of small-scale and large-scale model wing pressure characteristics showed similar trends created by upper surface blowing, while spanwise blowing characteristics differed considerably. Force data correlated well with semi-empirical predictions for gross thrust coefficients less than 1.0.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-2620 , V/STOL Conference; Dec 07, 1981 - Dec 09, 1981; Palo Alto, CA
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  • 174
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    Finite element analyses of plane thermal entry-length flows (1981)
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    Publication Date: 2019-07-13
    Description: It is pointed out that the design of actively cooled engine and airframe structures for hypersonic flight requires strong interaction between fluid, thermal, and structural analyses. Research programs are being conducted to develop an integrated thermal-structural analysis capability based upon the finite element method. In connection with these programs, a finite element engineering approach has been developed based upon a number of assumptions customarily used in practical heat transfer analysis. An evaluation is conducted of the finite element engineering approach for entry-length flows. The evaluation makes use of an analysis of plane thermal entry-length flows by a finite element approach utilizing mean fluid temperatures with a convection coefficient. Another analysis is based on a finite element continuum approach, taking into account the rigorous momentum, mass, and energy equations of viscous, incompressible flow. The engineering approach was found to give excellent agreement with the rigorous continuum approach for Peclet number values not less than 100.
    Keywords: AERODYNAMICS
    Type: In: International Conference on Finite Elements in Flow Problems; Jun 10, 1980 - Jun 13, 1980; Alberta; Canada
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  • 175
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    An experimental study of the flow in a wing with a partial span drooped leading edge (1981)
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    Publication Date: 2019-07-13
    Description: The flow field produced by a low aspect ratio wing (AR = 3.0) with a partial span leading edge droop was investigated in a series of low speed wind tunnel tests (Reynolds number based on 17.8 cm chord = 560,000). Photographs were obtained of surface oil flow patterns over an angle of attack range of alpha = 0 to 29 deg. Flow field surveys of the partially stalled wing at alpha = 25 deg were completed using a hot-wire probe, a split-film probe and a Conrad probe. The flow field survey data was presented using a color video display. The data indicated regions of apparent reversed flow in the separation region behind the wing and indicated the general cross-sectional shape of the separated wake flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-1665 , American Institute of Aeronautics and Astronautics, Aircraft Systems and Technology Conference; Aug 11, 1981 - Aug 13, 1981; Dayton, OH
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  • 176
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    Development and validation of the V/STOL aerodynamics and stability and control manual (1981)
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    Publication Date: 2019-07-13
    Description: A V/STOL Aerodynamics and Stability and Control Manual was developed to provide prediction methods which are applicable to a wide range of V/STOL configurations in hover and transition flight, in and out of ground effect. Propulsion-induced effects have been combined with unpowered aerodynamics in a buildup of total forces and moments for the jet-lift concept, so that total aerodynamics can be used to predict aircraft stability, control, and flying qualities characteristics. Results of longitudinal aerodynamic predictions have been compared with test data, and indicate that the methods are fast, inexpensive, and within the desired accuracy for the objective preliminary design stage.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-2611 , V/STOL Conference; Dec 07, 1981 - Dec 09, 1981; Palo Alto, CA
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  • 177
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    Preflight reentry heat transfer analysis of Space Shuttle (1981)
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    Publication Date: 2019-07-13
    Description: Preflight predictions of the structural temperature distributions during entry are compared with data from the initial Shuttle flight. Finite element thermal analysis programming was used to model the heat flow on Shuttle structures and actual gas properties of air were employed in the analyses of aerodynamic heating. Laminar, separated, and turbulent heat fluxes were calculated for varying locations on the craft using velocity-attitude and angle-of-attack projections taken from the nominal STS-1 trajectory. Temperature time histories of the first flight are compared with laminar and turbulent flow assumptions and an unpredicted rapid cooling 1800 sec into entry is credited to inaccurate assumptions of structural heat dissipative properties or flow conditions in that time phase of the flight; additional discrepancies in descriptions of heating of the upper fuselage are attributed to a lack of knowledge of the complex flow patterns existing over that area of the Shuttle body.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-2382 , Flight Testing Conference; Nov 11, 1981 - Nov 13, 1981; Las Vegas, NV
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  • 178
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    Propfan installation aerodynamics of a supercritical swept wing transport configuration (1981)
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    Publication Date: 2019-07-13
    Description: A semispan model with a powered propeller has been tested to provide data on the installation drag penalty of advanced propfan-powered transports designed to cruise at a Mach number of 0.8. These tests, conducted in 14-foot and 11-foot transonic wind tunnels, are a part of a NASA program to develop efficient, high-speed propellers for more fuel-efficient commercial transports for the 1990s and beyond. The model is instrumented for measuring propeller forces, wing/nacelle forces and moments, and pressure distributions over the wing and nacelle. The body in these tests was nonmetric, being connected to the wing by an RTV seal at the wing/body juncture. Tests were run at angles of attack from -3 to +5 deg over the Mach number range 0.6 to 0.85 at a Reynolds number of about 9,000,000. Results of these tests indicate that the nacelle interference drag can be quite large relative to an uninstalled nacelle. However, the losses due to the nacelle were reduced to acceptable levels by changes to the wing leading edge and nacelle intersection. The propeller slipstream causes substantial changes in the wing span load distribution indicating that twist modifications are needed to recover a more favorable span load distribution.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-1563 , Joint Propulsion Conference; Jul 27, 1981 - Jul 29, 1981; Colorado Springs, CO
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  • 179
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    Development of the ARIES parachute system (1981)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The design and testing of a two-stage parachute system to recover a space telescope weighing up to 2000 pounds is described. The system consists of a 15-ft dia ribbon parachute reefed to 50% for 10 seconds and a 73-ft dia paraform or cross second stage reefed to 10% for 10 seconds. The results of eight drop tests and one operational rocket launched flight and recovery are presented. A successful operational recovery of a 1600-lb NASA space telescope was conducted. The payload was launched by a second stage Minuteman rocket to an altitude of about 300 miles above sea level.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84082 , DE81-012161 , SAND-81-0208C , CONF-811002-1 , AIAA Aerodyn. Decelerator and Balloon Technol. Conf.; Oct 21, 1981 - Oct 23, 1981; San Diego, CA; United States
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  • 180
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    Correlation of transonic-cone Preston-tube data and skin friction (1981)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The distribution of Preston tube pressures within turbulent boundary layers along the surface of a sharp-nosed, ten degree cone was correlated with theoretical values of turbulent skin friction for freestream Mach numbers less than one. The mini-basic computer code, the Wu and Lock computer code, and the STAN-5 computer code were used to analyze the data and to solve the boundary layer conservation equations. The skin friction which results from using Preston tube pressures in the correlation equation, has a rms error of 1.125 percent. It was found that the effective center of the probe is not a constant but increases as the surface distance increases. For a specified unit Reynolds number, the effective center of the probe decreases as the Mach number increases. The variation of the fluid (air) properties across the face of the probe may be neglected for subsonic flows. The possible transverse errors caused by the use of the concept of a virtual origin for the turbulent boundary layer were investigated and found to be negligible.
    Keywords: AERODYNAMICS
    Type: NASA-CR-165065
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  • 181
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    Computational methods for internal flows with emphasis on turbomachinery (1981)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Current computational methods for analyzing flows in turbomachinery and other related internal propulsion components are presented. The methods are divided into two classes. The inviscid methods deal specifically with turbomachinery applications. Viscous methods, deal with generalized duct flows as well as flows in turbomachinery passages. Inviscid methods are categorized into the potential, stream function, and Euler aproaches. Viscous methods are treated in terms of parabolic, partially parabolic, and elliptic procedures. Various grids used in association with these procedures are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-82764 , E-1085 , The ASME Winter Ann. Meeting; Nov 15, 1981 - Nov 20, 1981; Washington, DC; United States
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  • 182
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    Three-dimensional periodic disturbances acting upon airfoils in cascade (1981)
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    Publication Date: 2019-07-13
    Description: A theory is developed for three-dimensional periodic gusts interacting with loaded airfoils. The unsteady disturbances are linearized with respect to the mean potential flow of the airfoils. The vorticity transport equation is then integrated analytically in a Lagrangian form. The vorticity vector shows strong variations in its magnitude and wavelength as it interacts with the flowfield of a lifting airfoil. The streamwise component of the vorticity increases significantly with flow acceleration and flow turning. For simplicity the rectangular fan approximation is used and a single Helmholtz-like equation is derived to characterize the unsteady 3D flowfield. This is a new and significant result. The present theory is first applied to symmetric airfoils.
    Keywords: AERODYNAMICS
    Type: International Symposium on Aeroelasticity in turbomachines; Sep 08, 1980 - Sep 12, 1980; Lausanne; Switzerland
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  • 183
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    Boundary layer development on turbine airfoil suction surfaces (1981)
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    Publication Date: 2019-07-13
    Description: The results of a study supported by NASA under the Energy Efficient Engine Program, conducted to investigate the development of boundary layers under the influence of velocity distributions that simulate the suction sides of two state-of-the-art turbine airfoils, are presented. One velocity distribution represented a forward loaded airfoil ('squared-off' design), while the other represented an aft loaded airfoil ('aft loaded' design). These velocity distributions were simulated in a low-speed, high-aspect-ratio wind tunnel specifically designed for boundary layer investigations. It is intended that the detailed data presented in this paper be used to develop improved turbulence model suitable for application to turbine airfoil design.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 81-GT-204 , Gas Turbine Conference and Products Show; Mar 09, 1981 - Mar 12, 1981; Houston, TX
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  • 184
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    Solution of plane cascade flow using improved surface singularity methods (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A solution method has been developed for calculating compressible inviscid flow through a linear cascade of arbitrary blade shapes. The method uses advanced surface singularity formulations which were adapted from those found in current external flow analyses. The resulting solution technique provides a fast flexible calculation for flows through turbomachinery blade rows. The solution method and some examples of the method's capabilities are presented.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 81-GT-169 , Gas Turbine Conference and Products Show; Mar 09, 1981 - Mar 12, 1981; Houston, TX
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  • 185
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    A quasi-three-dimensional blade surface boundary layer analysis for rotating blade rows (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A quasi-three-dimensional approximation has been developed for a blade boundary layer which involves the calculation of the effect of nonzero pressure gradients, turbulent flow, and blade twist, but includes only a simple coupling between streamlines. The resulting set of equations is solved using Keller's box scheme. The solution scheme is checked against available incompressible flow solutions and then applied to a NASA low aspect ratio transonic compressor stage for which extensive experimental and computational data are available. It is found that the three-dimensional boundary layer separates significantly sooner and has a much larger influence on rotor performance than would be expected from a two-dimensional analysis.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 81-GT-126 , Gas Turbine Conference and Products Show; Mar 09, 1981 - Mar 12, 1981; Houston, TX
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  • 186
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    On the convergence of unsteady generalized aerodynamic forces (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Variations in generalized forces calculated by different computer programs are traced to improper mathematical modeling techniques. Comparisons of generalized forces calculated by three theoretical methods are presented to illustrate difficulties involved in obtaining prediction convergence for increasing wave number. Use of a sufficiently dense chordwise paneling arrangement, in finite panel methods, results in predictions that are essentially identical to predictions of converged solutions. Procedural modifications are suggested for application in finite panel methods to increase prediction accuracy and reduce computer usage costs.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 81-0647 , Conference on Structures, Structural Dynamics and Materials; Apr 06, 1981 - Apr 08, 1981; Atlanta, GA
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  • 187
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    Analysis of turbulent boundary layer on cascade and rotor blades of turbomachinery (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The momentum integral technique for predicting the boundary layer growth in three-dimensional flow has been extended to include the entrainment equation as the closure model. The numerical solution is compared with the cascade, inducer, compressor, and fan rotor blade data from various sources. The agreement is found to be excellent in all cases, with the exception of the separated flow. Both the momentum thickness and the limiting streamline angle predicted from this analysis compare well with the measured data for a rotor blade. The technique is extremely useful in engineering design, analysis, and performance prediction.
    Keywords: AERODYNAMICS
    Type: International Symposium on Air Breathing Engines; Feb 16, 1981 - Feb 22, 1981; Bangalore; India
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  • 188
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    Flow separation in inlets at incidence angles (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Wind-tunnel pressure data and flow pictures obtained for two two-dimensional inlet models have been examined to study the internal flow structure and separation at large incidence angles. The inlet models were 12-in. high (diffuser exit height) and had internal contraction ratio of 1.21 and 1.17. They were tested at low forward speeds over a wide range of throat Mach numbers (inlet mass flow rates) and angles of incidence. Characteristic features of the internal flow such as a drastic change of pressure gradient near the highlight, local separation bubbles and shock/boundary-layer interactions have been indicated and discussed. For a few specific cases, the experimental surface pressure distributions have been compared with theoretical predictions.
    Keywords: AERODYNAMICS
    Type: International Symposium on Air Breathing Engines; Feb 16, 1981 - Feb 22, 1981; Bangalore; India
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  • 189
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    Perturbation of transonic flow with shocks (1981)
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    Publication Date: 2019-07-13
    Description: A general formulation of the perturbation problem is studied, and a new approach, perturbation sequence expansion, is introduced for handling shock disturbances. The method is applied to unsteady effects, three-dimensional corrections to axisymmetric and two-dimensional flows, and wind tunnel corrections. The perturbation equations are nonlinear and can be solved by shock capturing methods.
    Keywords: AERODYNAMICS
    Type: Symposium on Numerical and Physical Aspects of Aerodynamic Flows; Jan 19, 1981 - Jan 21, 1981; Long Beach, CA
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  • 190
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    Basic advances in the finite-volume method for transonic potential flow calculations (1981)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The finite-volume method of Jameson and Caughey provides a framework within which it is possible to calculate transonic potential flows about essentially arbitrary geometrical configurations. Improvements designed to increase the accuracy of the basic scheme and its consistency in the far field will be described. These include the incorporation of an artificial viscosity which maintains the formal second-order accuracy of the scheme in supersonic zones, and a modification of the flux balances to allow the free-stream conditions to satisfy the difference equations identically. Results of calculations illustrating the importance of these effects will be presented.
    Keywords: AERODYNAMICS
    Type: Symposium on Numerical and Physical Aspects of Aerodynamic Flows; Jan 19, 1981 - Jan 21, 1981; Long Beach, CA
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  • 191
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    Progress on interacting boundary-layer computations at high Reynolds number (1981)
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    Publication Date: 2019-07-13
    Description: The purpose of this paper is to review progress made in the solution of the interacting boundary-layer equations for subsonic flow. The interrelationship of triple deck theory and the interacting boundary-layer approach is discussed with emphasis placed on the development of efficient and reliable algorithms for the solution of the interacting boundary-layer equations. Example studies are presented for laminar and turbulent finite flat plate flow, laminar flow past a flat plate with a separation causing depression, and laminar and turbulent flow past a blunt based trailing edge.
    Keywords: AERODYNAMICS
    Type: Symposium on Numerical and Physical Aspects of Aerodynamic Flows; Jan 19, 1981 - Jan 21, 1981; Long Beach, CA
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  • 192
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    Factors Influencing the Predicted Performance of Advanced Propeller Designs (1981)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The assumptions on which conventional propeller aerodynamic performance analyses are based can be seriously violated when advanced high speed propellers are analyzed. Studies were performed using a lifting line representation for the propeller to determine the sensitivity of predicted propeller performance to various assumptions in the analysis. Items studied include the method of determining blade section lift and the effects of blade section drag, camber and blade sweep. The effects of nonuniform flow into the propeller and compressibility were also studied. Comparisons of analytical and experimental results are presented to demonstrate the overall validity of the results.
    Keywords: AERODYNAMICS
    Type: NASA-TM-82676 , E-942 , AIAA PAPER 81-1564 , Joint Propulsion Conf.; Jul 27, 1981 - Jul 29, 1981; Colorado Springs, CO; United States
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  • 193
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    Development of a nonlinear vortex method (1981)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Progress is reported in the development of reliable nonlinear vortex methods for predicting the steady and unsteady aerodynamic loads of highly sweptback wings at large angles of attack. Abstracts of the papers, talks, and theses produced through this research are included. The modified nonlinear discrete vortex method and the nonlinear hybrid vortex method are highlighted.
    Keywords: AERODYNAMICS
    Type: NASA-CR-164351
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  • 194
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    Numerical optimization techniques for bound circulation distribution for minimum induced drag of Nonplanar wings: Computer program documentation (1981)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A two dimensional advanced panel far-field potential flow model of the undistorted, interacting wakes of multiple lifting surfaces was developed which allows the determination of the spanwise bound circulation distribution required for minimum induced drag. This model was implemented in a FORTRAN computer program, the use of which is documented in this report. The nonplanar wakes are broken up into variable sized, flat panels, as chosen by the user. The wake vortex sheet strength is assumed to vary linearly over each of these panels, resulting in a quadratic variation of bound circulation. Panels are infinite in the streamwise direction. The theory is briefly summarized herein; sample results are given for multiple, nonplanar, lifting surfaces, and the use of the computer program is detailed in the appendixes.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3458 , NAS 1.26:3458
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  • 195
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    A flight investigation of blade-section aerodynamics for a helicopter main rotor having 10-64C airfoil sections (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Pressure data at 90 percent blade radius were obtained for a helicopter main rotor with 10-64C blade sections during flight. Concurrent measurements ere made of vehicle flight state, performance and some rotor loads. The test envelope included hover, level flight from about 65 to 162 knots, climb and descent, and collective fixed maneuvers. Good agreement is shown between some sets of airfoil pressure distributions obtained in flight and those from two-dimensional wind-tunnel tests or theoretical calculations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-83226
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  • 196
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    Experimental and analytical investigation of axisymmetric supersonic cruise nozzle geometry at Mach numbers from 0.60 to 1.30 (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Quantitative pressure and force data for five axisymmetric boattail nozzle configurations were examined. These configurations simulate the variable-geometry feature of a single nozzle design operating over a range of engine operating conditions. Five nozzles were tested in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.60 to 1.30. The experimental data were also compared with theoretical predictions.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1953 , L-14661
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  • 197
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    Wind-tunnel investigation of the effects of blade tip geometry on the interaction of torsional loads and performance for an articulated helicopter rotor (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Langley transonic dynamics tunnel was used to determine the degree of correlation between rotor performance and the dynamic twist generated by changes in blade tip geometry using an articulated rotor with four different tip geometries at advance ratios of 0.20, 0.30 and 0.35. Based on the data obtained, it is concluded that: (1) there appears to be no strong correlation between blade torsion loads and rotor performance prediction; (2) for a given rotor task at each advance ratio investigated, both the azimuthal variation of torsional moment and the mean torsional moment at 81% radius are configuration dependent; (3) reducing the nose down twist on the advancing blade appears to be more important to forward flight performance than increasing the nose down twist on the retreating blade; (4) the rotor inflow model used was important in predicting the performance of the adaptive rotor; and (5) neither rigid blade solidity effects, inflow environment, nor blade torsion loads can be used alone to accurately predict active rotor performance.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1926 , AVRADCOM-TR-81-B-5 , L-14674
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  • 198
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    Prediction of subsonic aircraft flows with jet exhaust interactions (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical procedure to calculate the flow fields resulting from the viscous inviscid interactions that occur when a strong jet exhaust and aircraft flow field coupling exists was developed. The approach divides the interaction region into zones which are either predominantly viscous or inviscid. The flow in the inviscid zone, which surrounds most of the aircraft, is calculated using an existing potential flow code. The viscous flow zone, which encompasses the jet plume, is modeled using a parabolized Navier-Stokes code. The procedure features the coupling of the zonal solutions such that sufficient information is transferred between the zones to preserve the effects of the interactions. The zonal boundaries overlap and the boundary conditions are the information link between zones. An iteration scheme iterates the coupled analysis until convergence has been obtained.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. of Power Plant Installation; 12 p
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  • 199
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    Crossflow effects on steady and fluctuating pressures on an ogive-cylinder cone-frustum model in supersonic separated flow (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wind-tunnel tests were conducted on an ogive-cylinder model with two axisymmetric protuberances having cone frustum angles of cone = 23 deg and 45 deg that were used to generate detached shock waves and the resulting separated flow areas downstream of the shock. The tests were conducted in a 9 by 7 foot supersonic wind tunnel at a free-stream Mach number of 2.0 and at Reynolds numbers of 1.5 x 1 million and 3.9 x 1 million, based on body diameter. The model had an afterbody fineness ratio of 8.3, and the ogive nose had a fineness ratio of 3.0. Two characteristics of the fluctuating pressures in surface vortex flows that result from the crossflow component, (velocity along the tunnel longitudinal axis free stream angle of attack), in combination with changes in the longitudinal pressure gradient were measured: (1) the broadband, rms-pressure coefficients and (2) the power spectral densities. Measurements are presented for various flow regions on the model such as the attached turbulent boundary layer, the detached frustum shock wave, and separated flow areas. The results indicate that the pressure fluctuations around or in the neighborhood of the foci of the vortex flows had broadband intensities and power spectral densities nearly identical to the levels previously measured in separated-flow regions at angles of attack of 0 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1951 , A-8563
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  • 200
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    Airframe-propulsion system aerodynamic interference predictions at high transonic Mach numbers including off-design engine airflow effects (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The transonic speed regime for airplanes at conditions where inlet spillage takes place is discussed. A wind tunnel test program to evaluate aerodynamic performance penalties associated with propulsion system installation and operation at subsonic through low supersonic speeds was conducted. The accuracy of analytic methods for predicting transonic engine airframe interference effects was assessed. Study variables included Mach number, angle of attack, relative nacelle location, and nacelle mass flow ratio. Results include test theory comparisons of forces as well as induced pressure fields. Prediction capability of induced shock wave strength and locations is assessed. It was found that large interference forces due to engine location and flow spillage occur at transonic speeds, that theory explains these effects; and that theory can predict quantitatively these effects.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. of Power Plant Installation; 23 p
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