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  • AERODYNAMICS  (956)
  • 1980-1984  (485)
  • 1975-1979  (471)
  • 1935-1939
  • 1984  (485)
  • 1979  (471)
  • 1935
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  • 1980-1984  (485)
  • 1975-1979  (471)
  • 1935-1939
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  • 1
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    Status of advanced airfoil tests in the Langley 0.3-meter transonic cryogenic tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A joint NASA/U.S. industry program to test advanced technology airfoils in the Langley 0.3-meter Transonic Tunnel (TCT) was formulated under the Langley ACEE Project Office. The objectives include providing U.S. industry an opportunity to compare their most advanced airfoils to the latest NASA designs by means of high Reynolds number tests in the same facility. At the same time, industry would again experience in the design and construction of cryogenic test techniques. The status and details of the test program are presented. Typical aerodynamic results obtained, to date, are presented at chord Reynolds number up to 45 x 10(6) and are compared to results from other facilities and theory. Details of a joint agreement between NASA and the Deutsche Forschungs- und Versuchsantalt fur Luft- and Raumfahrt e.V. (DFVLR) for tests of two airfoils are also included. Results of these tests will be made available as soon as practical.
    Keywords: AERODYNAMICS
    Type: Advan. Aerodyn.: Selected NASA Res.; p 37-53
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  • 2
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    Component testing for dynamic model verification (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Dynamic model verification is the process whereby an analytical model of a dynamic system is compared with experimental data, adjusted if necessary to bring it into agreement with the data, and then qualified for future use in predicting system response in a different dynamic environment. These are various ways to conduct model verification. The approach taken here employs Bayesian statistical parameter estimation. Unlike curve fitting, whose objective is to minimize the difference between some analytical function and a given quantity of test data (or curve), Bayesian estimation attempts also to minimize the difference between the parameter values of that funciton (the model) and their initial estimates, in a least squares sense. The objectives of dynamic model verification, therefore, are to produce a model which: (1) is in agreement with test data; (2) will assist in the interpretation of test data; (3) can be used to help verify a design; (4) will reliably predict performance; and (5) in the case of space structures, will facilitate dynamic control.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Recent Experiences in Multidisciplinary Analysis and Optimization, Part 2; 15 p
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  • 3
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    Post and a random-walk search mode (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Multidisciplinary analysis often requires optimization of nonlinear systems that are subject to constraints. Trajectory optimization is one example of this situation. The Program to Optimize Simulated Trajectories (POST) was used successfully for a number of problems. The purpose is to describe POST and a new optimization approach that has been incorporated into it. Typical uses of POST will also be illustrated. The projected-gradient approach to optimization is the preferred option in POST and is discussed. A new approach to optimization, the random-walk approach, is described, and results with the random-walk approach are presented.
    Keywords: AERODYNAMICS
    Type: Recent Experiences in Multidisciplinary Analysis and Optimization, Part 2; 23 p
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  • 4
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    Wall interference measurements for three-dimensional models in transonic wind tunnels: Experimental difficulties (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: The purpose is not to provide a detailed discussion of several wall interference experiments, but rather to use these experiments (recently accomplished in the Boeing Transonic Wind Tunnel (BTWT) to illustrate the problems associated with many of the measurements required by current wall interference assessment/correction (WIAC) procedures. The wall correction to lift is emphasized. It is shown that, because conventional tunnels and relatively small models continue to be used, the flow field or flow boundary measurements to be made impose severe requirements on the experiment itself. In some cases, existing instrumentation and test techniques may not be adequate to obtain the data accuracies needed.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 21-42
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  • 5
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    Adaptation of a four-wall interference assessment/correction procedure for airfoil tests in the 0.3-m TCT (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Based upon limited, initial observations of wall interference corrections obtained for one airfoil test, there is a need for assessing the upstream flow direction. If there is no direct measurement then a two-pass correction procedure similar to the one described here is required. Questions have arisen pertaining to the correct interpretation of the pressure coefficients measured on the slats of a slotted tunnel wall, the interpretation of just what the calculated equivalent body encompasses or should include, and what can or should be considered as quantitative criteria for data correctability. Further studies using this modified procedure will address these questions. Hopefully, a meaningful WIAC procedure can be validated for the airfoil tests in the 0.3-m TCT.
    Keywords: AERODYNAMICS
    Type: Wind Tunnel Wall Interference Assessment and Correction, 1983; p 393-414
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  • 6
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    Some experience with Barnwell-Sewall type correction to two-dimensional airfoil data (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A series of airfoils were tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) at Reynolds numbers from 2 to 50 million. The 0.3-m TCT is equipped with Barnwell slots designed to minimize blockage due to the tunnel flow and ceiling. This design suggests that sidewall corrections for blockage is needed, and that a lifting airfoil produces a change in angle of attack. Sidewall correction methods were developed for subsonic and subsonic-transonic flow. Comparisons of theory with experimental data obtained in the 0.3-m TCT for two airfoils, the British NPL 9510 and the German R-4 are presented. The NPL 9510 was tested as part of the NASA/United Kingdom Joint Aeronautical Program and R-4 was tested as part f the DFVLR/NASA Advanced Airfoil Research Program. For the NPL 9510 airfoil, only those test points that one would anticipate being difficult to predict theoretically are presented.
    Keywords: AERODYNAMICS
    Type: Wind Tunnel Wall Interference Assessment and Correction, 1983; p 375-392
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  • 7
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    Experiences with a high-blockage model tested in the NASA Ames 12-foot pressure wind tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Representation of the flow around full-scale ships was sought in the subsonic wind tunnels in order to a Hain Reynolds numbers as high as possible. As part of the quest to attain the largest possible Reynolds number, large models with high blockage are used which result in significant wall interference effects. Some experiences with such a high blockage model tested in the NASA Ames 12-foot pressure wind tunnel are summarized. The main results of the experiment relating to wind tunnel wall interference effects are also presented.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 345-360
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  • 8
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    An inerference assessment approach for a three-dimensional slotted tunnel with sparse wall pressure data (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: The various procedures referred to as wall interference assessment and correction procedures presume the existence of a surface distribution of data (usually static pressure) measured over a surface on or near the tunnel walls for each test point to be assessed. An alternative approach in which a reasonably sophisticated computer model of the test section flow would be fitted parametrically to a sparse set of measured data is presented. The measurements provides line distributions of static pressure near the center lines of the top, side and bottom walls. The development of a test section model incorporating explicit recognition of discrete slots of finite length with controlled flow reentry into the solid wall downstream portion of the tunnel is shown.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center. Wind Tunnel Wall Interference Assessment and Correction, 1983; p 323-334
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  • 9
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    Executive summary of Aircraft Icing Specialists Workshop (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: In a period of escalating development costs for new aircraft, there is growing interest in a renewed and coordinated icing research effort to achieve an updating or modernization of each aspect of the technological issues that are involved. This includes the data base, analysis methods, test techniques, and test facilities.
    Keywords: AERODYNAMICS
    Type: Aircraft Icing; p 1-16
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  • 10
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    Safety hazard of aircraft icing (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The problem of aircraft icing is reported as well as the type of aircraft affected, the pilots involved, and an identification of the areas where reduction in icing accidents are readily accomplished.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Res. Center Aircraft Icing; p 21-27
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  • 11
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    Civil Helicopter icing problems (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The ice capabilities of rotary wing aircraft are examined. Recommendations are given to improve the inadequacies of the weather forecasts pertaining to ice, and to adopt a low maintenance anti-ice system.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Res. Center Aircraft Icing; p 29-30
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  • 12
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    Overview of helicopter ice protection system developments (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Helicopter ice protection design criteria was developed and technological shortcoming in meeting helicopter mission requirements is that of helicopter rotor blade ice protection. Airframe components are protected using existing technology while the rotor blade protected using the cyclic electrothermal deicing concept.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Res. Center Aircraft Icing; p 39-65
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  • 13
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    A review of the icing situation from the standpoint of general aviation (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: An overview of the present situation in the field of aircraft icing with respect to certification and operation of nontransport category airplanes is given. Problems of definition and inconsistencies are pointed out. Problems in the forecasting and measurement of icing intensities are discussed. The present regulatory environment is examined with respect to its applicability and appropriateness to nontransport airplanes.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Res. Center Aircraft Icing; p 31-38
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  • 14
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    Icing of aircraft Some remarks with an historical slant from a cloud physicist (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Three areas of interest are commented on: cloud physics, nowcasting, and instrumentation. A comparison is made of what was done 30 years ago to what might be done in light of developments in related areas of cloud physics, weather modification and instrumentation.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Res. Center Aircraft Icing; p 17-19
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  • 15
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    Interference from slotted walls (1984)
    In:  CASI
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    Publication Date: 2006-02-14
    Description: Wall interference is made predominant in tunnel models and by wall geometries to facilitate the study of slot flow. The viscous effects in slots are studied by two dimensional measurements of flow. Wall interference is assessed by measuring pressure distributions at two levels near the walls. Interference on lifting delta wings is calculated. Pressure distributions at inner boundaries show basis axisymetries between the pressure side and the suction side, pointing to the necessity of having wider slots on the pressure side.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 293-300
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  • 16
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    Development in UK of a method for calculating tunnel wall corrections from flow measurements (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Classical methods for calculation of wall corrections which are not satisfactory for a number of flows of interest are discussed. To meet these objections, a number of methods were developed which use measurements of the low at or close to the tunnel walls as an outer boundary condition to define wall interference. The development, assessment and application of one such method is summarized.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 259-271
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  • 17
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    NLR activities in the field of wind tunnel wall interference (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Measured field data as a boundary condition for calculating the interference flow field were applied. They are divided into two categories. In the first category, the field data must consist of distributions of a single velocity component, and an accurate estimate of the hypothetical free air contribution of the model to this component is required. The differences between measured values and estimated model contributions are attributed to wall interference and they establish the boundary condition. The associated field data measurements are simple, yet the necessary model representation generally is a serious drawback. The second category requires field data which consist of velocity vector distributions at the price of multicomponent measurements, but at the profit that no information at all is required about the model. In solid wall test sections, the price is reduced to virtually zero but the profit remains.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 221-229
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  • 18
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    Performance of two transonic airfoil wind tunnels utilizing limited ventilation (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A limited-zone ventilated wall panel was developed for a closed-wall icing tunnel which permitted correct simulation of transonic flow over model rotor airfoil sections with and without ice accretions. Candidate porous panels were tested in the Ohio State University 6- x 12-inch transonic airfoil tunnel and result in essentially interference-free flow, as evidenced by pressure distributions over a NACA 0012 airfoil for Mach numbers up to 0.75. Application to the NRC 12- x 12-inch icing tunnel showed a similar result, which allowed proper transonic flow simulation in that tunnel over its full speed range.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 165-170
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  • 19
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    Investigations of flow field perturbations induced on slotted transonic-tunnel walls (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: The free-stream interference caused by the flow through the slotted walls of the test sections of transonic wind tunnels has continuously a problem in transonic tunnel testing. The adaptive-wall transonic tunnel is designed to actively control the near-wall boundary conditions by sucking or blowing through the wall. In order to make the adaptive-wall concept work, parameters for computational boundary conditions must be known. These parameters must be measured with sufficient accuracy to allow numerical convergence of the flow field computations and must be measured in an inviscid region away from the model that is placed inside the wind tunnel. The near-wall flow field was mapped in detail using a five-port cone probe that was traversed in a plane transverse to the free-stream flow. The initial experiments were made using a single slot and recent measurements used multiple slots, all with the tunnel empty. The projection of the flow field velocity vectors on the transverse plane revealed the presence of a vortex-like flow with vorticity in the free stream. The current research involves the measurement of the flow field above a multislotted system with segmented plenums behind it, in which the flow is controlled through several plenums simultaneously. This system would be used to control a three-dimensional flow field.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 119-142
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  • 20
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    Assessment of lift- and blockage-induced wall interference in a three-dimensional adaptive-wall tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A three-dimensional adaptive-wall wind tunnel experiment was conducted at Ames Research Center. This experiment demonstrated the effects of wall interference on the upwash distribution on an imaginary surface surrounding a lifting wing. This presentation demonstrates how the interference assessment procedure used in the adaptive-wall experiments to determine the wall adjustments can be used to separately assess lift- and blockage-induced wall interference in a passive-wall wind tunnel. The effects of lift interference on the upwash distribution and on the model lift coefficient are interpreted by a simple horseshoe vortex analysis.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 89-100
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  • 21
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    Wind tunnel wall interference in closed, ventilated and adaptive test sections (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A wall interference correction method for closed rectangular test sections was developed which uses measured wall pressures. Measurements with circular discs for blockage and a rectangular wing as a lift generator in a square closed test section validate this method. These measurements are intended to be a basis of comparison for measurements in the same tunnel using ventilated (in these case, slotted) walls. Using the vortex lattice method and homogeneous boundary conditions, calculations were performed which show sufficiently high pressure levels at the walls for correction purposes in test sections with porous walls. In Gottingen, an adaptive test section (which is a deformable rubber tube of 800 mm diameter) was built and a computer program was developed which is able to find the necessary wall adaptation for interference-free measurements in a single step. To check the program prior to the first run, the vortex lattice method was used to calculate wall pressure distributions in the nonadapted test section as input data for the one-step method. Comparison of the pressure distribution in the adapted test section with free-flight data shows nearly perfect agreement. An extension of the computer program can be made to evaluate the remaining interference corrections.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 61-78
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  • 22
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    Review of the advanced technology airfoil test program in the 0.3-meter transonic cryogenic tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: The following areas were addressed: interchangeable test sections in the 0.3-M Transonic Cryogenic Tunnel (TCT); typical airfoil installation; airfoil capability; advanced technology airfoil test (ATAT); effects of the Reynolds number on the normal force coefficient; effects of the Reynolds number on the drag coefficient; and comparison of experimental results with theory.
    Keywords: AERODYNAMICS
    Type: Wind Tunnel Wall Interference Assessment and Correction, 1983; p 361-374
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  • 23
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    Determination of equivalent model geometry for tunnel wall interference assessment/correction (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A formula for the determination of equivalent model geometry with two variables measured at the interface is derived, based on two dimensional subsonic flow. This predicted model profile is a reasonable initial estimate for transonic flow as long as the sonic region does not reach the interface. A general formula is given in two forms. One is in terms of complex variable functions and the other is an integral equation. The complex-function formula has the advantage of using analytic expressions. The integral equation form requires a numerical solution after assuming the model geometry as a polynomial function. Examples are given to illustrate the application of the formulas.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 335-342
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  • 24
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    Wind tunnel wall interference correction for aircraft models (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Wall interference correction procedures seek to determine the required changes in certain flow or geometric parameters so that the difference between the flow properties at the model's surface in the tunnel and free air are minimized. A transonic and a linear correction procedure were developed for aircraft models. In addition to Mach number and angle of attack corrections, an estimate of the accuracy of the corrections is provided by the transonic correction procedure. Lift, pitching moment and pressure measurements near the tunnel walls are required. The efficiency and accuracy of the correction procedure are improved. Moreover, correction of both the wing and tail angles of attack is allowed. The procedure is valid for transonic as well as subcritical flows. However, for subcritical flows further approximations and simplifying assumptions are made, leading to a very simple and efficient correction procedure.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center. Wind Tunnel Wall Interference Assessment and Correction, 1983; p 301-322
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  • 25
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    Tunnel constraint for a jet in crossflow (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A facet of a unified tunnel correction scheme which uses wall pressures to determine tunnel induced blockage and upwash is described. With this method, there is usually no need to use data concerning model forces or power settings to find the interference; it follows directly from the pressures and tunnel dimensions. However, highly inclined jets do not produce good pressure signatures and are highly three dimensional, so they must be treated differently. Flow modeling is also discussed.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center. Wind Tunnel Wall Interference Assessment and Correction, 1983; p 273-290
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  • 26
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    Progress in wind tunnel wall interference assessment/correction procedures at the NAE (1984)
    In:  CASI
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    Publication Date: 2006-02-14
    Description: Wall corrections as a function of wall porosity in the transonic wall interference problem was assessed. Effective porosities primarily for the two dimensional case were established as follows: (1) comparison of experimental data for two geometrically similar models of different chord/height ratio, an overall value of wall porosity could be deduced; (2) theoretical development which allows for unequal porosity for the floor and ceiling and wall boundary pressure measurements, porosities for floor and ceiling could be deduced; (3) a scheme was developed which allowed unequal porosity of floor and ceiling and streamwise varying porosity. The boundary layer development along the perforated floor and ceiling under the influence of the model pressure field, variations in boundary layer thickness underlining the difficulties in deducing meaningful values of wall porosity were determined. Wall boundary pressure measurement, in combination with singularity modelling of the airfoil, was sufficient to yield required information on the wall interference flow without having to establish some value for wall porosity. The singularity modelling of the airfoil initially covered only lift and volume but was extended to include drag and pitching moment, and second order volume term. It is shown by asymptotic transonic small disturbance analysis, that the derived corrections to angle of attack and free stream Mach number are correct to the first order.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 231-257
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  • 27
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    Asymptotic methods for wind tunnel wall corrections at transonic speed (1984)
    In:  CASI
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    Publication Date: 2006-02-14
    Description: The effort to develop classical methods to compute wall interference at transonic speeds is outlined. The two-dimensional theory and three-dimensional development are discussed. Also, some numerical application of the two-dimensional work are indicated. The basic advantages of the asymptotic theory are noted.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 193-203
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  • 28
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    Effect of boundary layers on solid walls in three-dimensional subsonic wind tunnels (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: A solution for the tunnel wall boundary layer effects for three-dimensional subsonic tunnels is presented. The model potentials are represented with simple singularities placed on the centerline of the tunnel and Laplace's equation in cylindrical coordinates is solved for either the conventional homogeneous slotted-wall boundary condition, the solid-wall viscous boundary condition, or a combination of them. The most pronounced wall boundary layer effect is on solid blockage for completely closed wind tunnels. Boundary layers on the wall reduce the blockage from the solid-wall, no-boundary-layer case in a manner similar to opening slots in a solid wall. Additionally, for solid-wall tunnel configurations, the streamline curvature interference factor is reduced by a significant amount, whereas the lift interference factor at the model station does not depend on the boundary layer parameter. For combination wall configurations, the slot effect of the horizontal walls dominates the viscous effect of the solid sidewalls.
    Keywords: AERODYNAMICS
    Type: Wind Tunnel Wall Interference Assessment and Correction, 1983; p 205-218
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  • 29
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    Experiments suitable for wind tunnel wall interference assessment/correction (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Three experiments suitable for wall interference assessment and evaluation of proposed correction methods are presented. The experiments are: (1) a series of airfoil tests using a newly designed transonic flow facility that employs side-wall boundary layer suction and upper- and lower-wall shaping; (2) tests on a swept airfoil section spanning a solid-wall wind tunnel with fixed contouring on all four walls; and (3) tests on a swept wing of aspect ratio 3 mounted in a solid-wall wind tunnel with fixed flat walls. Each of the experiments provides data on the airfoil sections as well as on the wind tunnel walls. All the experiments were performed in solid wall wind tunnels corrected for boundary layer displacement effects. Although the experiments were performed primarily to evaluate computer code performance, it is believed that they also provide information that can be used to evaluate methods for assessing and correcting wall interference effects.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 171-190
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  • 30
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    Effect of upstream sidewall boundary layer removal on an airfoil test (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: Sidewall boundary layer effects were investigated by applying partial upstream sidewall boundary layer removal in the Langley 0.3-m transonic cryogenic tunnel. Over the range of sidewall boundary layer displacement thickness of these tests the influence on pressure distribution was found to be small for subcritical conditions; however, for supercritical conditions the shock position was affected by the sidewall boundary layer. For these tests (with and without boundary layer remove) comparisons with predictions of the GRUMFOIL computer code indicated that Mach number corrections due to the sidewall boundary layer improve the agreement for both subcritical and supercritical conditions. The results also show that sidewall boundary layer removal reduces the magnitude of the sidewall correction; however, a suitable correction must still be made.
    Keywords: AERODYNAMICS
    Type: Wind Tunnel Wall Interference Assessment and Correction, 1983; p 143-163
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    A data base for three-dimensional all-interference code evaluation (1984)
    In:  CASI
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    Publication Date: 2006-02-14
    Description: A validation of a measured boundary condition technique was carried out to demonstrate the feasibility of a wall interference assessment/correction (WIAC) system. An experimental evaluation was also carried out to compare performances of various techniques, to define the number of necessary boundary measurements for accurate assessment/corrections and to define the envelope of test conditions for which accurate assessment/corrections are achieved. The relative merits of a WIAC system and an adaptive wall tunnel are compared. The measurement surface boundary data is performed with a system of two rotating pipes. These pipes sweep out a cylindrical measurement surface near the tunnel walls, approximately one inch from the wall at the closest point. The experimental model was specially designed and fabricated for the adaptive wall experiments. The model is a wing/tail/body configuration with swept lifting surface. The boundary data taken in Tunnel 1T with the rotating pipe system has been shown to offer several attractive features for WIAC code evaluation. Good spatial resolution of measurements is achieved and measurements are made upstream and downstream of the model. Also, two velocity components are determined.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 101-118
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    Survey of ONERA activities on adaptive-wall applications and computation of residual corrections (1984)
    In:  CASI
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    Publication Date: 2006-02-14
    Description: The research undertaken concerning the computation and/or reduction of wall interference follows two main axes: improvement of wall correction determinations, and use of adaptive flexible walls. The use of wall-measured data to compute interference effects is reliable when the model representation is assessed by signatures with known boundary conditions. When the computed interferences are not easily applicable to correcting the results (especially for gradients in two-dimensional cases), the flexible adaptive walls in operation in T2 are an efficient and assessed means of reducing the boundary effects to a negligible level, if the direction and speed of the flow are accurately measured on the boundary. The extension of the use of adaptive walls to three-dimensional cases may be attempted since the residual corrections are assumed to be small and are computable.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Wind Tunnel Wall Interference Assessment and Correction, 1983; p 43-60
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  • 33
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    Viscous transonic flow computation over Space Shuttle configuration (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A thin-layer Navier-Stokes code capable of predicting steady-state viscous flows is applied to the transonic flow over a Space Shuttle configuration. The code is written in the generalized coordinate system, and the grid-generation code of Fujii (1983) is used for the discretization of the flow field. The flow-field computation is done using the CRAY 1S computer at NASA Ames. The computed result is physically reasonable, even though no experimental data is available for the comparison purpose.
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  • 34
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    Nonlinear aerodynamic effects on bodies in supersonic flow (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 809-815
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  • 35
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    New implicit boundary procedures - Theory and applications (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 5, p. 579, Accession no. A83-16536
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1094-110
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  • 36
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    Numerical simulations of unsteady transonic flow in diffusers (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 15, p. 2346, Accession no. A82-31959
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1139-114
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  • 37
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    Effects of viscosity on transonic-aerodynamic and aeroelastic characteristics of oscillating airfoils (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 700-707
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  • 38
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    Leeside flows over delta wings at supersonic speeds (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 680-686
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  • 39
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    Fresh attack on laminar flow (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Laminar flow control is a technology with great potential for aircraft drag reduction. Stabilization of laminar boundary layers became known as natural laminar flow (NLF) and research led to the development of NLF airfoils. Research was also conducted on stabilization by suction, referred to as laminar flow control (LFC). Experiments demonstrated that extensive laminar flow could be achieved in flight. However, there remained doubts regarding the practicality of producing, with the technology then available, wing surfaces sufficiently smooth and wavefree to meet laminar-flow criteria and maintaining the wing surface quality in normal service. In 1976, the Aircraft Energy Efficiency (ACEE) program was begun by NASA to develop fuel-conservative technology for commercial transports. The progress of the ACEE program is discussed. Attention is given to LFC wing structures, and LFC leading-edge systems.
    Keywords: AERODYNAMICS
    Type: Aerospace America (ISSN 0740-722X); 22; 72-76
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  • 40
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    Space Shuttle heating analysis with variation in angle of attack and catalycity (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 5, p. 586, Accession no. A83-16747
    Keywords: AERODYNAMICS
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4560); 21; 217-219
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  • 41
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    Metric-discontinuous zonal grid calculations using the Osher scheme (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Computations on zonal grids - in particular, grids with metric discontinuities resulting from the interspersion of highly clustered regions with coarse regions - are possible using a fully conservative form of the Osher upwind scheme. These zonal grids can result from an abrupt clustering of points near solution discontinuities or near other flow features that require improved resolution. The zonal approach is shown to capture shocks with almost 'shock-fitting' quality but with minimal effort. Results for inviscid flow, including quasi-one-dimensional nozzle flow, supersonic flow over a cylinder, and blast-wave diffraction by a ramp, are presented. These calculations demonstrate the powerful capabilities of the Osher scheme used in conjunction with zonal grids in simulating flow fields with complex shock patterns.
    Keywords: AERODYNAMICS
    Type: Computers and Fluids (ISSN 0045-7930); 12; 3, 19
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  • 42
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    Lifting surface theory for a helicopter rotor in forward flight (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A lifting surface theory was developed for a helicopter rotor in forward flight for compressible and incompressible flow. The method utilizes the concept of the linearized acceleration potential and makes use of the vortex lattice procedure. Calculations demonstrating the application of the method are given in terms of the lift distribution on a single rotor, a two-bladed rotor, and a rotor with swept-forward and swept-back tips. In addition, the lift on a rotor which is vibrating in a pitching mode at 4/rev is given. Compressibility effects and interference effects for a two-bladed rotor are discussed.
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  • 43
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    Analysis of surface pressure distributions on two elliptic missile bodies (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 528-533
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  • 44
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    A general perturbation approach for computational fluid dynamics (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1748-175
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  • 45
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    Supersonic axisymmetric flow over boattails containing a centered propulsive jet (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1358-136
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  • 46
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    A study of the aerodynamic interaction between a main rotor and a fuselage (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The wake of a helicopter rotor can have a significant effect on a fuselage. Results from a recent wind-tunnel investigation show that certain fuselage characteristics, normalized by rotor thrust, scale proportionally to a rotor-wake-induced velocity parameter. Effects on the body of changes in velocity, thrust, tip-path-plane angle of attack, and rotor/body position are discussed. These results show that the rotor can have a favorable or unfavorable influence on the body, depending upon the operating condition.
    Keywords: AERODYNAMICS
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  • 47
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    Applying slender wing benefits to military aircraft (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 545-559
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  • 48
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    Transonic pressure distributions on a rectangular supercritical wing oscillating in pitch (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 576-582
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  • 49
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    A new consistent spatial differencing scheme for the transonic full-potential equation (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 5, p. 585, Accession no. A83-16678
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1027-103
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  • 50
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    Flow and heat transfer measurements along a cooled supersonic diffuser (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: An experimental investigtion was conducted to ascertain the mean flowfield, including shock wave structure, separated flow regions, turbulent boundary-layer growth, static pressure variations, wall heat transfer, and shear stresses in a second-throat, axisymmetric, supersonic diffuser with wall cooling. The diffuser inlet Mach number of the heated air flow was 3.76, the stagnation pressure was 6.8 atm, the ratio of wall to total gas temperature was 0.44, and the diffuser discharged to the atmosphere. The complex flowfield involved deceleration and acceleration regions, supersonic and embedded subsonic regions, and strong viscous regions with relatively large radial and axial variations. The heat transfer and wall static pressure distributions were remarkably similar, and heat transfer rates were high locally at oblique shock/turbulent boundary-layer interactions, in the pseudoshock region, and in the separation region in the diffuser outlet section.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 777-780
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  • 51
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    On the convergence of unsteady generalized aerodynamic forces (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 12, p. 1923, Accession no. A81-29500
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 9921-8669); 21; 420-427
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  • 52
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    Heating distributions for a 0.006-scale Shuttle Orbiter over a range of hypersonic flow conditions (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 15, p. 2342, Accession no. A82-31855
    Keywords: AERODYNAMICS
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 120-122
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  • 53
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    Overall aerodynamic performance of an annular flat-plate airfoil cascade (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 05, p. 588, Accession no. A83-16824
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 250
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  • 54
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    Calculated effects of varying Reynolds Number and dynamic pressure on flexible wings at transonic speeds (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: A computational method is described that includes the effects of static aeroelastic wing deflections in steady transonic aerodynamic calculations. This method, known as the Transonic Aero-elastic Program System (TAPS), interacts a 3D transonic computer code with boundary layer and a linear finite element structural analysis codes to calculate wing pressures and deflections. The nonlinear nature of the transonic flow makes it necessary to couple the aerodynamic and structures codes in an iterative manner. TAPS has been arranged in a modular fashion so that different aerodynamic or structures programs may be used with a minimum of coding changes required. Results obtained using two different aerodynamic codes in TAPS are given, and those results are correlated with experimental data.
    Keywords: AERODYNAMICS
    Type: Recent Experiences in Multidisciplinary Analysis and Optimization, Part 1 (date]; 19 p
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  • 55
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    Sensitivity analysis in computational aerodynamics (1984)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: Information on sensitivity analysis in computational aerodynamics is given in outline, graphical, and chart form. The prediction accuracy if the MCAERO program, a perturbation analysis method, is discussed. A procedure for calculating perturbation matrix, baseline wing paneling for perturbation analysis test cases and applications of an inviscid sensitivity matrix are among the topics covered.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center Recent Experiences in Multidisciplinary Analysis and Optimization, Part 1; 10 p
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  • 56
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    Numerical solution for perpendicular sonic hydrogen injection into a ducted supersonic airstream (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: This note discusses a computer program being developed to study the flow field near opposing perpendicular fuel injectors in scramjets. The MacCormack time-split, finite difference relaxation technique was used to solve the full two-dimensional compressible Navier-Stokes equations along with energy and species equations. By using this technique, a program was developed to consider the turbulent nonreacting flow of hydrogen and air in a rectangular duct. A damping term, proportional to the second derivative of pressure and temperature, was used to produce a stable solution behind the hydrogen jet in the neighborhood of the recompression shock. A case using actual conditions encountered in current scramjet design was analyzed, with results agreeing qualitatively with experimental observations.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 17; May 1979
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  • 57
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    Generalized velocities in the outer region of hypersonic turbulent boundary layers (1979)
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    Publication Date: 2011-08-17
    Description: An attempt is made to show that the outer portion of the velocity profile of hypersonic turbulent boundary layers can be transformed so that the constants determined by a best fit to the law of the wake are in reasonable agreement with the wake constant for incompressible boundary layers at the same Reynolds number. Both y transformations (where y is distance from the surface) and velocity transformations produce velocity profiles which, with the proper choice of wall shear stress to give shear velocity, can be reduced to the incompressible law of the wall.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 17; Aug. 197
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  • 58
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    Accurate numerical evaluation of modified Struve functions occurring in unsteady aerodynamics (1979)
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    Publication Date: 2011-08-17
    Description: A method is developed for computing the modified Struve functions that occur in unsteady aerodynamics. The method uses a rational approximation supplemented by an asymptotic series for large argument. Simple recursive formulas for generating the coefficients are derived. The method is capable of generating results of arbitrary accuracy. It can also be used for complex argument and order. For greater computing speed, a method is presented that uses the rational and asymptotic approximations to generate Chebyshev coefficients.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 16; July 197
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  • 59
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    V/STOL wind-tunnel testing (1984)
    In:  CASI
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    Publication Date: 2012-05-17
    Description: Factors influencing effective program planning for V/STOL wind-tunnel testing are discussed. The planning sequence itself, which includes a short checklist of considerations that could enhance the value of the tests, is also described. Each of the considerations, choice of wind tunnel, type of model installation, model development and test operations is discussed, and examples of appropriate past and current V/STOL test programs are provided. A short survey of the moderate to large subsonic wind tunnels is followed by a review of several model installations, from dimensional to large-scale models of complete aircraft configurations. Model sizing, power simulation, and planning are treated, including three areas in test operations: data acquisition systems, acoustic measurements in wind tunnels, and flow surveying.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on V(STOL Aerodyn.; 71 p
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  • 60
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    Laminar and turbulent flows over a spherically blunted cone with massive surface blowing (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Numerical solutions are presented for the flow over a spherically blunted cone with massive surface blowing. Time-dependent viscous shock layer equations are used to describe the flowfield. The boundary conditions on the body surface include a prescribed blowing rate distribution. The governing equations are solved by a time-asymptotic finite-difference method. Results presented here are only for a perfect gas-type flow at zero angle of attack. Both laminar and turbulent flow solutions are obtained. It is found that the surface blowing smooths out the effect of the curvature discontinuity at the sphere-cone juncture point on the laminar flowfield and results in a negative pressure gradient over the body. The shock slope increases on the downstream portion of the body as the surface blowing rate is increased. The turbulent flow with surface blowing is found to redevelop a boundary-layer-like region near the surface. The effects of this boundary-layer-like region on the flowfield and heating rates are discussed.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 17; Dec. 197
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    Fast, conservative schemes for the full potential equation applied to transonic flows (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Implicit approximate factorization techniques (AF) are investigated for the solution of matrix equations resulting from finite-difference approximations to the full potential equation in conservation form. For transonic flows, an artificial viscosity, required to maintain stability in supersonic regions, is introduced by an upwind bias of the density. Two implicit AF procedures are presented, and their convergence performance is compared with that of the standard transonic solution procedure: successive line overrelaxation (SLOR). Subcritical and supercritical test cases are considered. Results indicate a substantial improvement in convergence rate for AF schemes relative to SLOR.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 17; Feb. 197
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  • 62
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    Criteria for asymptotic supersonic two-dimensional turbulent wakes (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Indicators used to determine the fully developed mean flow for two dimensional turbulent supersonic wakes are examined. The similarity variables for velocity temperature and the transverse coordinate used by Demetriades (1969) and Wagner (1972) are shown not to adequately distinguish transition from developed turbulent flows. The growth rate of the two dimensional turbulent wake as the wake develops from laminar to turbulent are shown to be low in the laminar region, increase at transition, and level off as fully developed turbulence is attained, demonstrating that wake growth rates are better indicators of a fully developed mean flow than are similarity variables.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 17; July 197
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  • 63
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    Quasi-periodic structure of a turbulent jet (1979)
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    Publication Date: 2011-08-17
    Description: The instantaneous near field pressure fluctuations of an axisymmetric subsonic jet were measured by using a longitudinal and an azimuthal microphone arrays in order to qualitatively determine the behaviors of the quasi-periodic structure within the flow. Statistical analysis is used to explain the characteristic of the pressure signals. In addition to the information obtained by forming the power spectral density, auto- and cross-correlation functions, two types of signals are extracted through a conditional probability analysis to represent the quasi-periodic and the random fine structures within the turbulent jet. The quasi-periodic structure first appears as a rolling up of the mixing layer flow within one nozzle diameter downstream of the exit, then becomes fully developed at approximately 3 nozzle diameters downstream with a preferred Strouhal number range 0.3-0.4, and finally disappears beyond the end of potential core. This behavior is also reflected in the variation of the convection velocity.
    Keywords: AERODYNAMICS
    Type: Journal of Sound and Vibration; 64; May 8
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  • 64
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    Reynolds number effects on the turbulence field in compressible boundary layers (1979)
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    Publication Date: 2011-08-17
    Description: Detailed experiments were conducted in a zero pressure gradient, supersonic turbulent boundary layer, including measurements of the three components of velocity fluctuations and the turbulent shear stress, for Reynolds numbers ranging from 11.7 million to 105 million at a freestream Mach number of 2.3. The mean flow measurements established the fully developed and equilibrium nature of the boundary layer. Measurements of the turbulence field show that the vertical and transverse fluctuations are essentially equal throughout the boundary layer at all Reynolds numbers, a feature that is different from observations in incompressible flows. The data show that the boundary layer exhibits similarity in the turbulence profiles for the entire Reynolds number range and agrees with previous compressible and incompressible data using Morkovin's scaling to account for compressibility effects.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 17; Apr. 197
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  • 65
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    Unsteady transonic flow in cascades (1984)
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    Publication Date: 2011-08-19
    Description: There is a need for methods to predict the unsteady air loads associated with flutter of turbomachinery blading at transonic speeds. The results of such an analysis in which the steady relative flow approaching a cascade of thin airfoils is assumed to be transonic, irrotational, and isentropic is presented. The blades in the cascade are allowed to undergo a small amplitude harmonic oscillation which generates a small unsteady flow superimposed on the existing steady flow. The blades are assumed to oscillate with a prescribed motion of constant amplitude and interblade phase angle. The equations of motion are obtained by linearizing about a uniform flow the inviscid nonheat conducting continuity and momentum equations. The resulting equations are solved by employing the Weiner Hopf technique. The solution yields the unsteady aerodynamic forces acting on the cascade at Mach number equal to 1. Making use of an unsteady transonic similarity law, these results are compared with the results obtained from linear unsteady subsonic and supersonic cascade theories. A parametric study is conducted to find the effects of reduced frequency, solidity, stagger angle, and position of pitching axis on the flutter.
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    Transonic viscous-inviscid interaction using Euler and inverse boundary-layer equations (1984)
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    Publication Date: 2011-08-19
    Description: The paper is concerned with the use of a zonal method for the computation of transonic viscous-inviscid interacting flow about airfoils. The inviscid portion of the flow is treated by using an Euler equation solution method, while an inverse integral compressible turbulent boundary-layer solution method is used for the viscous portion of the flow. The matching of the viscous and inviscid solutions is discussed, and some numerical results as well as comparisons with experimental data are presented.
    Keywords: AERODYNAMICS
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    Approximate-factorization algorithms - Theory and applications in viscous-flow computations (1984)
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    Publication Date: 2011-08-19
    Description: A systematic development of implicit approximate-factorization algorithms in delta form for both unsteady and steady viscous flow is presented. The algorithms are cast in conservation-law form and simplified by using a thin-layer approximation to the governing equations. The implementation of implicit surface viscous boundary conditions is discussed in detail, and an example is presented illustrating the advantage of using the implicit boundary conditions. Three-dimensional results from the steady form of the algorithm are presented and compared with experimental data.
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    Finite difference simulation of unsteady interactive flows (1984)
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    Publication Date: 2011-08-19
    Description: The development of time-dependent numerical simulations of unsteady interactive flows of an aerodynamic nature is reviewed with emphasis on compressible flows at flight Reynolds numbers and noniterative schemes based on Navier-Stokes equations. The importance of writing the equations in strong conservation-law form for a generalized body-oriented coordinate system is pointed out. The discussion covers time and length scales and numerical methods currently in use. Some computed results are presented and compared with experimental data.
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    Nonunique solutions to the transonic potential flow equation (1984)
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    Publication Date: 2011-08-18
    Description: Steinhoff and Jameson (1981) have shown that within a certain range of angle of attack and freestream Mach number, numerical solutions of the full-potential equation for flow past an airfoil are not unique. This study was mainly concerned with showing that the anomaly is inherent to the partial-differential equation governing the flow and not a result of its discrete representation. Steinhoff and Jameson conjectured that the anomaly may have a physical basis. The present investigation has two objectives. Results are to be presented which indicate that the anomaly is due to a breakdown in the potential approximation, rather than a phenomenon associated with the inviscid flowfield. The second objective is to show that the lift coefficient, predicted by the potential equation, is a smooth but multivalued function of the angle of attack.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 145
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  • 70
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    Computations and aeroelastic applications of unsteady transonic aerodynamics about wings (1984)
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 13, p. 2016, Accession no. A82-30157
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 37-43
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  • 71
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    Adaptation of two-dimensional transonic analysis code, TSFOIL, for use in modeling adaptive wall test sections (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The effects of the elastic deformation of the wind tunnel wall held to a streamline-like shape were simulated numerically. The wall itself is simulated by a finite element model and is allowed to deform under the pressure loading developing in the wind tunnel with an airfoil model present. A modified version of the transonic analysis program, TSFOIL, is then used to determine the resulting flow field with the effects of the deformed wall included, the shape of the wall and the flow field. Once a result from a particular operating condition was obtained, the pressure distribution on the airfoil in the wind tunnel model is compared with solutions generated by TSFOIL in the free air mode.
    Keywords: AERODYNAMICS
    Type: A Coop. Program to Stimulate Student Involvement through the MIT Undergraduate Research Opportunity Program; 19 p
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  • 72
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    Supersonic wing design concepts employing nonlinear flows (1984)
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    Publication Date: 2011-08-18
    Description: Three nonlinear flow concepts for the design of supersonic wings are reviewed. The specific concepts are: leading-edge thrust, supercritical crossflow, and leading-edge vortex flow. The major results of the experimental-theoretical studies supporting the development of these concepts are presented and discussed. Also, supporting aerodynamic prediction methods are described and example applications are given. Recommendations for further development of each concept are made.
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  • 73
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    Validation of a transonic analysis code for use in preliminary design of advanced transport configurations (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The WBPPW code has the capability of analyzing flow-field effects about configurations which include wing pylons and engine nacelles or pods in addition to the basic wing/fuselage combination. Using the concept of grid embedding, the code solves the extended small disturbance transonic flow equation for complex flow interactions of the various configuration components. A general description of the code and solution algorithm is included. Results are presented and compared with experiment for various configurations which encompass the code capabilities. These include wing planform and wing contour modifications and variations in nacelle position beneath a high-aspect-ratio wing. Results are analyzed in the light of preliminary design, where the capability to accurately compute flow-field effects resulting from various configuration perturbations is important. The comparisons show that the computational results are sensitive to subtle design modifications and that the code could be used as an effective guide during the design process for transport configurations.
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  • 74
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    Prediction of transonic separated flows (1984)
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    Publication Date: 2011-08-18
    Description: Johnson et al. (1982) have provided a detailed comparison between a thoroughly documented transonic flow with shock-induced separations and solutions of the flow using the Navier-Stokes equations. According to this comparison, there were several deficiencies in the computations. The present investigation takes into account new experimental data which have been obtained in a larger wind tunnel with the same test model for a wider range of freestream Mach numbers. The results of new Navier-Stokes computations using more compatible boundary conditions are shown, and the effects of the turbulence model choice on predicting Mach number trends are assessed.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1001-100
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  • 75
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    Improved design of subcritical and supercritical cascades using complex characteristics and boundary-layer correction (1984)
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    Publication Date: 2011-08-18
    Description: The method of complex characteristics and hodograph transformation for the design of shockless airfoils was extended to design supercritical cascades with high solidities and large inlet angles. This capability was achieved by introducing a conformal mapping of the hodograph domain onto an ellipse and expanding the solution in terms of Tchebycheff polynomials. A computer code was developed based on this idea. A number of airfoils designed with the code are presented. Various supercritical and subcritical compressor, turbine and propeller sections are shown. The lag-entrainment method for the calculation of a turbulent boundary layer was incorporated to the inviscid design code. The results of this calculation are shown for the airfoils described. The elliptic conformal transformation developed to map the hodograph domain onto an ellipse can be used to generate a conformal grid in the physical domain of a cascade of airfoils with open trailing edges with a single transformation. A grid generated with this transformation is shown for the Korn airfoil. Previously announced in STAR as N83-24474
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 950-956
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  • 76
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    Turbulent boundary-layer relaxation with application to skin-friction drag reduction (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 05, p. 584, Accession no. A83-16633
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 871
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  • 77
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    Surface pressures on a flat plate with dual jet configurations (1984)
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 17, p. 2456, Accession no. A83-38677
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 484-490
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  • 78
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    Two-dimensional wind-tunnel interference from measurements on two contours (1984)
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    Publication Date: 2011-08-18
    Description: This paper describes how wall-induced velocities near a model in a two-dimensional wind tunnel can be estimated from upwash distributions measured along two contours surrounding a model. The method is applicable to flows that can be represented by linear theory. It was derived by applying the Schwarz Integral Formula separately to the two contours and by exploiting the free-air relationship between upwashes along the contours. Advantages of the method are that only one flow quantity need by measured and no representation of the model is required. A weakness of the method is that it assumes streamwise interference velocity vanishes far upstream of the model. This method was applied to a simple theoretical model of flow in a solid-wall wind tunnel. The theoretical interference velocities and the velocities computed using the method were in excellent agreement. The method was then used to analyze experimental data acquired during adaptive-wall experiments at Ames Research Center. This analysis confirmed that the wall adjustments reduced wall-induced velocities near the model.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 414-419
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  • 79
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    Effect of boundary layers on solid walls in three-dimensional subsonic wind tunnels (1984)
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 05, p. 580, Accession no. A83-16553
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 365-371
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  • 80
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    Effect of a buried-wire gage on the separation bubble Numerical study (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A numerically simulated buried-wire separation gage is investigated with emphasis on its effect on the separation bubble. The conjugated problem of a supersonic, time-dependent, two-dimensional flowfield above a conductive solid wall with an embedded heat source is solved using implicit finite difference algorithms. Steady-state and transient cases were computed for different locations of the heat source within the bubble. Results show that by using a steady heat source, the flow direction near the wall can be detected, without distorting the flowfield, only if the source is located in regions where the bubble is thick (i.e., not too close to the separation). The flow direction near separation can be detected by using a temperature pulse at the solid/fluid interface with insignificant distortion of the flowfield.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1539-154
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  • 81
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    Rotary-wing aerodynamics. Volume 1 - Basic theories of rotor aerodynamics (With application to helicopters). Volume 2 Performance prediction of helicopters (2nd revised and enlarged edition) (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Basic theories of rotor aerodynamics are presented and applied to the performance prediction of helicopters. The very simple physicomathematical model of the rotor offered by momentum theory is addressed first, followed by the combined blade-element and momentum theory. Vortex theory is discussed, and a rotor blade is modeled by means of a vortex filament or vorticity surface. Considerations of airfoil sections suitable for rotors are examined. Detailed performance techniques for a single-rotor helicopter in hover, vertical ascent, and forward flight are described, and winged and tandem-rotor helicopter performance calculations are presented as extensions and modifications of single-rotor methodology. Computer data based on the vortex theory are compared with approximate results obtained from the simplified momentum theory and the blade element solution.
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  • 82
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    Resonance in flows with vortex sheets and edges (1984)
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    Publication Date: 2011-08-19
    Description: It is shown that the vortex sheet in a slot between two semi-infinite plates does not admit incompressible resonant perturbations. The semi-infinite vortex sheet entering a duct does admit incompressible resonance. These results indicate that the vortex-sheet approximation is less useful for impinging shear flows than for non-impinging flows. They also suggest an important role of downstream vortical disturbances in resonant flows. The general solution for perturbations to flow with a vortex sheet and edges is written in terms of a Cauchy integral. Requirements on the behavior of this solution at edges and at downstream infinity fix the criteria for resonance.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics (ISSN 0022-1120); 145; 275-285
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  • 83
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    Generation of instability waves in flows separating from smooth surfaces (1984)
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    Publication Date: 2011-08-19
    Description: This paper analyses the coupling between an imposed disturbance and an instability wave that propagates downstream on a shear layer which emanates from a separation point on a smooth surface. Since the wavelengths of the most-amplified instability waves will generally be small compared with the streamwise body dimensions, the analysis is restricted to this 'high-frequency' limit and the solution is obtained by using matched asymptotic expansions. An 'inner' solution, valid near the separation point, is matched onto an outer solution, which represents an instability wave on a slowly diverging mean flow. The analysis relates the amplitude of this instability to that of the imposed disturbance.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics (ISSN 0022-1120); 145; 71-94
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  • 84
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    A time-split finite-volume algorithm for three-dimensional flowfield simulation (1984)
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    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1564-157
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  • 85
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    A fast viscous correction method for full-potential transonic wing analysis (1984)
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    Publication Date: 2011-08-18
    Description: An analysis of the transonic flowfield around a three-dimensional wing is carried out using a strip method. Attention is given to the boundary layer growth in the streamwise direction. A viscous correction technique is defined for the TWING code for solving the full potential equations. A viscous ramp at the base of a shock is superimposed on the boundary layer displacement thickness generated by an integral boundary layer method. A relationship is then obtained between the effective displacement thickness and a vertical component of the surface velocity, a transpirational boundary condition. The viscous correction is found to be unnecessary in weak shock conditions but gives a better shock position and pressure distribution in a strong shock condition when compared with data from an ONERA M6 airfoil and the Hinson and Burdges (1980) Wing A.
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    A new aerodynamic integral equation based on an acoustic formula in the time domain (1984)
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    Publication Date: 2011-08-18
    Description: An aerodynamic integral equation for bodies moving at transonic and supersonic speeds is presented. Based on a time-dependent acoustic formula for calculating the noise emanating from the outer portion of a propeller blade travelling at high speed (the Ffowcs Williams-Hawking formulation), the loading terms and a conventional thickness source terms are retained. Two surface and three line integrals are employed to solve an equation for the loading noise. The near-field term is regularized using the collapsing sphere approach to obtain semiconvergence on the blade surface. A singular integral equation is thereby derived for the unknown surface pressure, and is amenable to numerical solutions using Galerkin or collocation methods. The technique is useful for studying the nonuniform inflow to the propeller.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1337-134
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  • 87
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    Viscous-inviscid interactive procedure for rotational flow in cascades of airfoils (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1281
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  • 88
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    Hypersonic interactions with surface mass transfer. I - Steady flow. II - Unsteady flow (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 15, p. 2345, Accession no. A82-31944
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 449-452
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    Impact of computers on aerodynamics research and development (1984)
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    Publication Date: 2011-08-18
    Description: Factors motivating the development of computational aerodynamics as a discipline are traced back to the limitations of the tools available to the aerodynamicist before the development of digital computers. Governing equations in exact and approximate forms are discussed together with approaches to their numerical solution. Example results obtained from the successively refined forms of the equations are presented and discussed, both in the context of levels of computer power required and the degree of the effect that their solution has on aerodynamic research and development. Factors pacing advances in computational aerodynamics are identified, including the amount of computational power required to take the next major step in the discipline. Finally, the Numerical Aerodynamic Simulation (NAS) Program - with its 1987 target of achieving a sustained computational rate of 1 billion floating-point operations per second operating on a memory of 240 million words - is briefly discussed in terms of its projected effect on the future of computational aerodynamics.
    Keywords: AERODYNAMICS
    Type: IEEE, Proceedings (ISSN 0018-9219); 72; 68-79
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  • 90
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    Some steady and oscillating airfoil test results, including the effects of sweep, from the tunnel spanning wing (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: A large scale tunnel spanning wing was built and tested. The model can be operated as either a swept or unswept wing and can be tested in steady state or oscillated sinusoidally in pitch about its quarter chord. Data is taken at mid-span with an internal 6-component balance and is also obtained from miniature pressure transducers distributed near the center span region. A description is given of the system and a brief discussion of some of the steady and unsteady results obtained to date. These are the steady load behavior to Mach numbers of approximately 1.1 and unsteady loads, including drag, at a reduced frequency of approximately 0.1.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 445-458
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    A new airfoil research capability (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: The design and construction of a self streamlining wall test section for the Langley 0.3 meter transonic cryogenic tunnel was included in the fiscal year 1978 construction of facilities budget for Langley Research Center. The design is based on the research being carried out by M. J. Goodyer at the University of Southampton, Southampton, England, and is supported by Langley Research Center. This paper presents a brief description of the project. Included are some of the design considerations, anticipated operational envelope, and sketches showing the detail design concepts. Some details of the proposed operational mode, safety aspects, and preliminary schedule are presented.
    Keywords: AERODYNAMICS
    Type: Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 425-432
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  • 92
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    Design and calibration of slotted walls for transonic airfoil wind tunnels (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: The traditional procedure for estimating the performance of slotted walls for airfoil wind tunnels is reviewed, and a modification which improves the accuracy of this procedure is described. Unlike the traditional procedure, the modified procedure indicates that the design of airfoil wind-tunnel walls which induce minimal blockage and streamline-curvature effects is feasible. The design and testing of such a slotted wall is described. It is shown experimentally that the presence of a model can affect the plenum pressure and thus make the use of the plenum pressure as a calibration reference questionable. Finally, an ONERA experiment which shows the effect of the sidewall boundary layer on the measured model normal force is discussed.
    Keywords: AERODYNAMICS
    Type: Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 433-443
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    Inverse boundary-layer technique for airfoil design (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: A description is presented of a technique for the optimization of airfoil pressure distributions using an interactive inverse boundary-layer program. This program allows the user to determine quickly a near-optimum subsonic pressure distribution which meets his requirements for lift, drag, and pitching moment at the desired flow conditions. The method employs an inverse turbulent boundary-layer scheme for definition of the turbulent recovery portion of the pressure distribution. Two levels of pressure-distribution architecture are used - a simple roof top for preliminary studies and a more complex four-region architecture for a more refined design. A technique is employed to avoid the specification of pressure distributions which result in unrealistic airfoils, that is, those with negative thickness. The program allows rapid evaluation of a designed pressure distribution off-design in Reynolds number, transition location, and angle of attack, and will compute an airfoil contour for the designed pressure distribution using linear theory.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 383-397
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    Langley's two-dimensional research facilities: Capabilities and plans (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: The current capabilities and the forthcoming plans for Langley's two-dimensional research facilities are described. The characteristics of the Langley facilities are discussed in terms of Reynolds number, Mach number, and angle-of-attack capabilities. Comments are made with regard to the approaches which have been investigated to alleviate typical problem areas such as wall boundary effects. Because of the need for increased Reynolds number capability at high subsonic speeds, a considerable portion of the paper deals with a description of the 20 by 60 cm two-dimensional test section of the Langley 0.3 meter transonic cryogenic tunnel which is currently in the calibration and shakedown phase.
    Keywords: AERODYNAMICS
    Type: Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 399-414
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    The prediction of two-dimensional airfoil stall progression (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: A generalized boundary condition potential flow calculation method was combined with a momentum integral boundary layer method and a base flow theory of separation to predict airfoil viscous-inviscid interference up to and beyond stall. The resultant program considers laminar and turbulent separation and is, therefore, applicable to thin or thick airfoil stall. The calculated flow field includes the airfoil and the separation bubble recombination region behind the airfoil. Calculated pressure distributions and equivalent airfoil shapes, including the displacement thickness of the viscous regions, are compared with flow field measurements for several airfoils. The measured displacement thicknesses and wake centerlines corroborate the calculated shape. The comparison also suggests the use of the analytical solution to evaluate the measurements.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 335-345
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    Application of numerical optimization to the design of advanced supercritical airfoils (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: An application of numerical optimization to the design of advanced airfoils for transonic aircraft showed that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while minimizing the drag coefficient at the design Mach number. This multiple design-point numerical optimization has been implemented with the use of airfoil shape functions which permit a wide range of attainable profiles during the optimization process. Analytical data for the starting airfoil shape, a single design-point optimized shape, and a double design-point optimized shape are presented. Experimental data obtained in the NASA Ames two-by two-foot wind tunnel are also presented and discussed.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 315-325
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    Improved prediction of laminar leading edge separation (1979)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: Research was conducted to provide a definite criterion for the prediction of the bubble burst on airfoils typical of those used for fighter wings. The approach taken was to correlate existing airfoil bubble burst data using various parameters at the laminar separation point. The method due to Weber was modified to provide a continuous analytic solution for the velocity distribution around the airfoil leading edge. Coupling the modified Weber method with the Stratford laminar separation prediction method leads to a universal chart giving the conditions at separation as a function of stagnation location and leading edge radius. Application of the combined method to available two-dimensional airfoil data resulted in an empirical criterion presenting the limiting local velocity gradient at separation as a function of the boundary layer momentum thickness at separation for bubble burst. The correlation leads as well to the qualitative explanation of two types of laminar stall: thin airfoil and leading edge. The validity of the correlation is demonstrated by predicting the lift coefficient and angle of attack for stall on airfoils with leading edge or trailing edge flaps.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 327-334
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    Optimization of multi-element airfoils for maximum lift (1979)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Two theoretical methods are presented for optimizing multi-element airfoils to obtain maximum lift. The analyses assume that the shapes of the various high lift elements are fixed. The objective of the design procedures is then to determine the optimum location and/or deflection of the leading and trailing edge devices. The first analysis determines the optimum horizontal and vertical location and the deflection of a leading edge slat. The structure of the flow field is calculated by iteratively coupling potential flow and boundary layer analysis. This design procedure does not require that flow separation effects be modeled. The second analysis determines the slat and flap deflection required to maximize the lift of a three element airfoil. This approach requires that the effects of flow separation from one or more of the airfoil elements be taken into account. The theoretical results are in good agreement with results of a wind tunnel test used to corroborate the predicted optimum slat and flap positions.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt.1; p 237-253
    Format: application/pdf
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    Improvements in surface singularity analysis and design methods (1979)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The coupling of the combined source vortex distribution of Green's potential flow function with contemporary numerical techniques is shown to provide accurate, efficient, and stable solutions to subsonic inviscid analysis and design problems for multi-element airfoils. The analysis problem is solved by direct calculation of the surface singularity distribution required to satisfy the flow tangency boundary condition. The design or inverse problem is solved by an iteration process. In this process, the geometry and the associated pressure distribution are iterated until the pressure distribution most nearly corresponding to the prescribed design distribution is obtained. Typically, five iteration cycles are required for convergence. A description of the analysis and design method is presented, along with supporting examples.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 221-236
    Format: application/pdf
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    Prospects for computing airfoil aerodynamics with Reynolds averaged Navier-Stokes codes (1979)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The Reynolds averaged Navier-Stokes equations are solved numerically for a variety of transonic airfoil configurations where viscous phenomena are important. Illustrative examples include flows past sensitive geometries, Reynolds number effects, and buffet phenomena.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Advanced Technol. Airfoil Res., Vol. 1, Pt. 1; p 119-131
    Format: application/pdf
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