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  • AERODYNAMICS  (383)
  • 1975-1979  (383)
  • 1935-1939
  • 1977  (383)
  • 1
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    Transonic/supersonic inlet R and D past, present, and future (1977)
    In:  CASI
    Details
    Publication Date: 2004-10-07
    Keywords: AERODYNAMICS
    Type: NASA Lewis Research Center Inlet Workshop; p 427-480
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  • 2
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    Comments on wind-tunnel-flight correlations for the F-8 supercritical wing configurations (1977)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Keywords: AERODYNAMICS
    Type: Supercritical Wing Technol.: A Report on Flight Evaluation; p 111-120
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  • 3
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    Aircraft (1977)
    In:  CASI
    Details
    Publication Date: 2006-01-16
    Description: The physical principles of flight, and the consideration of atmospheric composition and aerodynamic forces in the design and construction of various types of aircraft are discussed. Flight characteristics are described for helicopters, rotary-wing aircraft, short and vertical takeoff aircraft, and tailess or variable geometry wing aircraft. Flow characteristics at various speeds are also discussed.
    Keywords: AERODYNAMICS
    Type: Soviet Aircraft and Rockets (NASA-TT-F-770); p 24-80
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  • 4
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    Numerical solution for subcritical flows by a transonic integral equation method (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The method presented makes use of a division of the region of integration into closed rectangular elements. The velocity is taken to be constant in each element. The integral equation is reduced to a matrix equation which can be solved by an appropriate iteration approach. The derivation and solution of the matrix equation are discussed and the matrix elements are considered. The described concepts were implemented for a nonlifting parabolic-arc airfoil.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Mar. 197
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  • 5
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    Derivation of an integral equation for transonic flows (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: It is noted that the nonlinear partial differential equation for the perturbation velocity potential and boundary conditions describing steady inviscid compressible transonic flow past a thin two-dimensional airfoil can be transformed into a singular integrodifferential equation and that differentiation of the latter yields an integral equation. Two forms of this integral equation currently exist: one for the singularity that is enclosed in an infinitely long strip of vanishing thickness and the other for the singularity that is enclosed in a vanishing circle. In the present article, a more general integral equation is derived by enclosing the singularity in a vanishing rectangular cavity of arbitrary aspect ratio. The two existing forms of this equation are deduced as special cases distinguished by the respective values for the aspect ratio (infinity for the first form and unity for the second).
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Feb. 197
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  • 6
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    Flow characteristics in an expansion tunnel as inferred from velocity measurements (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A photoionization technique was used to study flow characteristics in an expansion tunnel. Vertical surveys of the axial component of flow velocity just downstream from the nozzle exit were obtained, and estimates of freestream density were inferred from the velocity measurement technique. The pitot pressure was measured and compared to the average axial component of velocity as a function of time for the two cases when air and CO2 were used as test gases. Vertical velocity and static density profiles at the nozzle exit are presented for the case when CO2 was used as test gas. Experimental results were used to determine the diameter and uniformity of the test core at the nozzle exit and the duration of the quasi-steady flow period. These data are relevant to evaluation of the suitability of operating an expansion tube as an expansion tunnel. The expansion tunnel is an expansion tube with a conical nozzle positioned at the exit of the acceleration section, so that nozzle entrance flow conditions are hypersonic and characterized by hypervelocity.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Sept
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  • 7
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    Low Reynolds number flow past a blunt axisymmetric body at angle of attack (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Low Reynolds number flow of an ideal gas over a blunt axisymmetric body of large half-angle at small angles of attack is investigated, for the case of laminar hypersonic flow. Time-varying viscous shock layer equations describing the flowfield are obtained from the full Navier-Stokes system by keeping terms to second order in the inverse square root of Re in both viscous and inviscid regions; the equations are valid for moderate to high Re. Drag, skin friction, and heating rates were obtained at small (or zero) angles of attack. Conditions experienced by planetary entry probes during the high-altitude (early) legs of an atmospheric entry trajectory are pertinent to the problem.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Aug. 197
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  • 8
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    Evaluation of flight spoilers for vortex alleviation (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The paper describes the facilities and test procedures used in a series of wind-tunnel and full-scale flight investigations of the effectiveness of flight spoilers currently existing on wide-bodied transport jet aircraft when used as trailing vortex hazard alleviation devices. Examples of the results of such studies include the variation of trailing wing rolling-moment coefficient with downstream distance behind a B-747 airplane model with various segments of its flight spoilers deflected 45 deg, and comparisons with models without spoilers deflected. It is concluded that the existing flight spoilers on the B-747 are effective as trailing vortex attenuators.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Aug. 197
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  • 9
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    The effect of ablation injection on radiative and convective heating (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A viscous shock-layer analysis for calculating high energy equilibrium flow fields about blunt axisymmetric bodies is applied to the problem of massive ablation injection with radiation transport. A nongray radiation model is used that accounts for both line and continuum radiation. The solution method is direct and provides both stagnation and downstream solutions. Results for shock heated air show that phenolic-nylon injection is substantially more effective in reducing the wall radiant flux than air injection. Also, for large included body angles, the wall radiative flux and the coupled phenolic-nylon injection rate do not continue to decrease with increasing distance downstream.
    Keywords: AERODYNAMICS
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  • 10
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    Comment on 'Drag coefficient of spheres in continuum and rarefied flows' (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A paper by Henderson (1976) provides a method of predicting experimental sphere drag data. This approach uses two equations for the drag coefficient, one for relative Mach number less than one, one for relative Mach number greater than 1.75. For relative Mach numbers between these limits a linear interpolation procedure is followed. In a comment on this paper, it is claimed, on the basis of comparing predictions with experimental results, that a method proposed by Walsh (1975) gives better predictions of the drag coefficient for relative Mach numbers less than 1.75, provided that a modification of the procedure is made for relative Mach numbers less than 0.1. For values over 1.75, both methods are considered equally accurate. In a reply to this comment, it is agreed that the Walsh method is more accurate when Reynolds numbers are within a range between 20 and 200, and Mach numbers are between 0.5 and 1.25. Presumed errors and possible limitations in the Walsh procedure for predicting drag coefficients are discussed.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; June 197
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  • 11
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    Vortex lattice prediction of subsonic aerodynamics of hypersonic vehicle concepts (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The vortex lattice method introduced by Lamar and Gloss (1975) was applied to the prediction of subsonic aerodynamic characteristics of hypersonic body-wing configurations. The reliability of the method was assessed through comparison of the calculated and observed aerodynamic performances of two National Hypersonic Flight Research Facility craft at Mach 0.2. The investigation indicated that a vortex lattice model involving 120 or more panel elements can give good results for the lift and induced drag coefficients of the craft, as well as for the pitching moment at angles of attack below 10 to 15 deg. Automated processes for calculating the local slopes of mean-camber surfaces may also render the method suitable for use in preliminary design phases.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Oct. 197
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  • 12
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    Cooling effectiveness of slot injection into a turbulent boundary layer (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The cooling effectiveness of injection through multiple flush slots at an angle of 10 deg was studied experimentally in a wind tunnel. Air was injected from one to four slots into a turbulent Mach 6 boundary layer. The slot mass flow ratio is defined, and data which describe the dependence of the cooling effectiveness on the slot mass flow ratio are presented. Experimental values are indicated graphically for various cases of single and multiple slot injection, where the total mass injection (i.e., the sum of flow rates from each slot) is the same for each case. The results show that, for a given coolant mass flow rate, thermal protection over the maximum surface area can be accomplished best by injecting the coolant flow through multiple slots.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Sept
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  • 13
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    Wall cooling effects on hypersonic transitional/turbulent boundary layers at high Reynolds numbers (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The characteristics of a thick hypersonic boundary layer turbulent for a length of 175 cm on a 4 deg sharp wedge were measured. The resulting boundary layer was free from transverse curvature effects and only mildly affected by upstream history effects caused by pressure and wall temperature gradients. Heat-transfer distributions were used to locate regions of laminar, transitional, and turbulent flow at an edge unit Reynolds number of 470,000 cm at wall-to-total temperature ratios from about 0.3 to 1. Wall cooling had little effect on the location of the transition region. Pitot and total temperature profiles and skin-friction measurements were obtained at several locations along the model longitudinal centerline. Mixing length and turbulent Prandtl number distributions were derived from the fully turbulent mean profiles.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Oct. 197
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  • 14
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    Application of the method of integral relations to laminar boundary layers in three dimensions (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The method of integral relations is extended to general three-dimensional compressible laminar boundary layer flows. The transformation employed to transform the basic three-dimensional compressible boundary layer equations into quasi-incompressible form is an extension of the Howarth transformation. The resulting system of differential equations is integrated numerically by the method of integral relations as proposed by Dorodnitsyn. To demonstrate the accuracy of the method, it is applied to calculation of the parabolic flow over a flat plate and the boundary flow over an infinite yawed cylinder, for which solutions are known. It is then applied to the flow over a flat plate disturbed by a cylinder normal to the plate, for which a finite-difference solution is available for comparison. It is finally applied to calculating the crossflow velocity variation for supersonic flow over a swept wedge.
    Keywords: AERODYNAMICS
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  • 15
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    A novel concept for subsonic inlet boundary-layer control (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A self-bleeding method for boundary layer control is described and tested for a subsonic inlet designed to operate in the flowfield generated by high angles of attack. Naturally occurring surface static pressure gradients are used to remove the boundary layer from a separation-prone region of the inlet and to reinject it at a less critical location with a net performance gain. The results suggest that this self-bleeding method for boundary-layer control might be successfully applied to other inlets operating at extreme aerodynamic conditions.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Apr. 197
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  • 16
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    Mach 6 flowfield survey at the engine inlet of a research airplane (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A flowfield survey was conducted to better define the nature of vehicle forebody flowfield at the inlet location of an airframe-integrated scramjet engine mounted on the lower surface of a high-speed research airplane to be air launched from a B-52 and rocket boosted to Mach 6. The tests were conducted on a 1/30-scale brass model in a Mach-6 20-in. wind tunnel at Reynolds number of 11,200,000 based on distance to engine inlet. Boundary layer profiles at five spanwise locations indicate that the boundary layer in the area of the forebody centerline is more than twice as thick as the boundary layer at three outboard stations. It is shown that the cold streak found in heating contours on the centerline of the forebody is caused by a thickening of the boundary layer on the centerline, and that this thickening decreases with angle of attack.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Apr. 197
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  • 17
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    Comparison of numerical and experimental 'conical' flow fields in supersonic corners with compression and/or expansion (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The flow field produced by the intersection of two plane solid surfaces in a supersonic stream is a complex interference flow. These flows can be fully compressive, fully expansive, or of mixed compression-expansion nature. This paper presents a comparison of the experimentally obtained flow-field structure in an axial corner with that predicted numerically by using a shock-capturing finite-difference method. The effect of sweep and surface deflection are evaluated, and the general influence of each is presented for the three classes of corner flow. The results show that the numerical method is a valuable aid in understanding the flow structure for simple configurations. In addition, confidence in the numerical method is gained for use in solving more general three-dimensional configurations where the flow is nonconical and several wave interaction may be presented.
    Keywords: AERODYNAMICS
    Type: British Aircraft Corp.
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  • 18
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    Unsteady flow in a supersonic cascade with strong in-passage shocks (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Linearized theory is used to study the unsteady flow in a supersonic cascade with in-passage shock waves. We use the Wiener-Hopf technique to obtain a closed-form analytical solution for the supersonic region. To obtain a solution for the rotational flow in the subsonic region we must solve an infinite set of linear algebraic equations. The analysis shows that it is possible to correlate quantitatively the oscillatory shock motion with the Kutta condition at the trailing edges of the blades. This feature allows us to account for the effect of shock motion on the stability of the cascade. Unlike the theory for a completely supersonic flow, the present study predicts the occurrence of supersonic bending flutter. It therefore provides a possible explanation for the bending flutter that has recently been detected in aircraft-engine compressors at higher blade loadings.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics; 83; Dec. 5
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  • 19
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    Correlation of turbulent shear layer attachment peak heating near Mach 6 (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A correlation of new turbulent two-dimensional data and peak heating data for attaching free shear layers is presented for a 2.54-cm and 5.08-cm diam cylindrical leading-edge slab 25.4 cm long, and 7.62 and 10.16 cm wide. A 30.48 x 25.4 cm sharp leading-edge flat plate set at 15 and 20 deg is used to generate plane impinging shocks. The freestream Mach number is 6 and the freestream Reynolds number varies from 3,300,000 to 25,600,000/m. Peak heating is measured on silica-based epoxy models with a phase change coating technique. A comparison of the free shear layer data with the transition data of Birch and Keyes (1972) reveals that the shear layer data are turbulent at attachment. The trend of the data shows that peak heating is strongly affected by the state of development at attachment. As the free shear layers become more fully developed, the data approach the two-dimensional correlation. Persistence of transitional flow structures for supersonic free shear flows is pointed out.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Dec. 197
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  • 20
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    Calculated heating on the afterbody nozzle of a hypersonic aircraft concept (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The present analysis was carried out to estimate the heating levels of the external nozzle of a scramjet/airframe-integrated research aircraft. A parametric examination of the effects of Mach number, reference length, and wall temperature showed that the heating rate distributions are independent of reference length and wall temperature. The initial heating rates obtained for a Mach 6 flight are in the (3 to 8) x 10 to the 5th power W/sq m range. Underlying the entire study is the question of nozzle boundary layer formation and growth, as well as the question of the reference length value that should be used in the computations. It is shown that the reference length is not the dominant factor setting the heating levels; an attempt to bound the actual length was made. A more detailed calculation of the rates requires further work to gain a better understanding of the combustor exit boundary layer.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; Dec. 197
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  • 21
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    Aerodynamic characteristics at high angles of attack (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: An overview is presented of the aerodynamic inputs required for analysis of flight dynamics in the high-angle-of-attack regime wherein large-disturbance, nonlinear effects predominate. An outline of the presentation is presented. The discussion includes: (1) some important fundamental phenomena which determine to a large extent the aerodynamic characteristics of airplanes at high angles of attack; (2) static and dynamic aerodynamic characteristics near the stall; (3) aerodynamics of the spin; (4) test techniques used in stall/spin studies; (5) applications of aerodynamic data to problems in flight dynamics in the stall/spin area; and (6) the outlook for future research in the area.
    Keywords: AERODYNAMICS
    Type: Von Karman Inst. for Fluid Dyn. Aerodyn. Inputs for Probl. in Aircraft Dyn., Vol. 2; 39 p
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  • 22
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    Vortex attenuation flight experiments (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Flight tests evaluating the effects of altered span loading, turbulence ingestion, combinations of mass and turbulence ingestion, and combinations of altered span loading turbulance ingestion on trailed wake vortex attenuation were conducted. Span loadings were altered in flight by varying the deflections of the inboard and outboard flaps on a B-747 aircraft. Turbulence ingestion was achieved in flight by mounting splines on a C-54G aircraft. Mass and turbulence ingestion was achieved in flight by varying the thrust on the B-747 aircraft. Combinations of altered span loading and turbulence ingestion were achieved in flight by installing a spoiler on a CV-990 aircraft and by deflecting the existing spoilers on a B-747 aircraft. The characteristics of the attenuated and unattenuated vortexes were determined by probing them with smaller aircraft. Acceptable separation distances for encounters with the attenuated and unattenuated vortexes are presented.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 369-403
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  • 23
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    Ground-based facilities for evaluating vortex minimization concepts (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: To determine the feasibility of altering the formation and decay of aircraft trailing vortexes through aerodynamic means, the test capabilities of two wind tunnels and two towing basins were used. The facilities, common models, and measurement techniques that were employed in the evaluation of vortex minimization concepts are described.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 129-156
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  • 24
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    Ground development and flight correlation of the vortex attenuating spline device (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Data presented indicate that the wing-mounted spline is a effective vortex-attenuating device. A comparison of the vortex induced rolling moment results at a separation scale distance of 0.70 km with those measured in full scale flight indicate good agreement for the unattenuated vortex configuration. The comparison also indicates that the spline effectiveness in flight was greater than in the ground facility test. The results of an applications study show that, for the heavy commercial jet aircraft studied, use of the splines does result in some degradation of the climb gradient and rate of climb, but the aircraft should meet certification requirements.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 271-303
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  • 25
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    Application of laser velocimetry to aircraft wake-vortex measurements (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The theory and use of a laser velocimeter that makes simultaneous measurements of vertical and longitudinal velocities while rapidly scanning a flow field laterally are described, and its direct application to trailing wake-vortex research is discussed. Pertinent measurements of aircraft wake-vortex velocity distributions obtained in a wind tunnel and water towing tank are presented. The utility of the velocimeter to quantitatively assess differences in wake velocity distributions due to wake dissipating devices and span loading changes on the wake-generating model is also demonstrated.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 157-192
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  • 26
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    Effect of compliant wall motion on turbulent boundary layers (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A critical analysis of available wall data which indicated drag reduction under turbulent boundary layers. Detailed structural dynamic calculations suggest the surfaces responded in a resonant, rather than compliant, manner. Alternate explanations are given for drag reductions observed in two classes of experiments: flexible pipe flown, and waterbacked membranes in air. Analysis indicates the wall motion for the remaining data is typified by short wave lengths in agreement with the requirement of a possible compliant wall drag reduction mechanism recently suggested by Langley.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on Concepts for Drag Reduction; 26 p
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  • 27
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    Methods for Reducing Subsonic Drag Due to Lift (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: The results of repeat experimental research on methods for reducing subsonic drag due to lift are discussed. The NASA supercritical airfoils and their application to structurally practical wings with increased aspect radio are described. A design approach and experimental results for wing-tip-mounted winglets are presented. Several methods for utilizing the thrust of jet engines to provide reductions in the drag due to lift are also discussed.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on Concepts for Drag Reduction; 17 p
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  • 28
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    Efficient solution of unsteady transonic flows about airfoils (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: An implicit finite difference procedure was developed for the efficient solution of unsteady transonic flow fields. Sample computations illustrate applications of procedures to aerodynamic problems. Solutions are presented that illustrate three types of shock wave motion that can result from airfoil control surface oscillations. The significant effect of wind tunnel wall conditions on these shock wave motions is demonstrated. Solutions are also presented for a simple aeroelastic problem in which the flow field equations and the structural motion equations are integrated simultaneously in time. Both stable and unstable aeroelastic interactions are considered. The procedure is adapted to compute unsteady aerodynamic force coefficients by the indicial method.
    Keywords: AERODYNAMICS
    Type: AGARD Unsteady Airloads in Separated and Transonic Flow; 11 p
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  • 29
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    Assessment of existing analytic methods for prediction of high angle-of-attack loads on delta wings at supersonic speeds (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: An assessment of the applicability of four loading prediction methods to high angle-of-attack conditions for simplified wing-body configurations is provided. The methods are: The tangent wedge approximation, the linear theory methods of Middleton and Woodward, and a shock-fitting finite-difference technique. Estimates obtained by these methods were compared with experimental pressure data on delta wings to examine the effects of Mach number, camber, sweep angle, and angle of attack. Results indicate that all of the methods provided reasonable estimates at moderate angles of attack. At these moderate angles of attack, the methods of Middleton and Woodward provided good estimates at Mach numbers higher than those usually associated with linear theory. Only the finite-difference method provided reasonable load estimates at high angles of attack.
    Keywords: AERODYNAMICS
    Type: AGARD Prediction of Aerodynamic Loading; 7 p
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  • 30
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    Comparisons of Theoretical and Experimental Pressure Distributions on an Arrow-Wing Configuration at Subsonic, Transonic, and Supersonic Speeds (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A wind tunnel test of an arrow wing body configuration consisting of flat and twisted wings, as well as a variety of leading- and trailing-edge control surface deflection, has been conducted at Mach numbers from 0.40 to 2.50 to provide an experimental data base for comparison with theoretical methods. Theory-to-experiment comparisons of detailed pressure distributions have been made using current state-of-the-art attached- and separated-flow methods. The purpose of these comparisons was to delineate conditions under which these theories are valid for aeroelastic calculations and to explore the use of empirical methods to correct the theoretical methods where theory is deficient. It was determined that current state-of-the-art attached flow and empirical methods were inadequate to predict aeroelastic loads for this configuration.
    Keywords: AERODYNAMICS
    Type: AGARD Prediction of Aerodynamic Loading; 14 p
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  • 31
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    Inviscid modeling of aircraft trailing vortices (1977)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: A survey is presented of inviscid theoretical methods that are useful in the study of lift-generated vortices. Concepts derived using these invisicid theories are cited which have helped to guide research directed at alleviating the velocities and rolling moments imposed on aircraft entering these wakes.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 9-60
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  • 32
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    Unsteady aerodynamic modeling for arbitrary motions (1977)
    In:  Other Sources
    Details
    Publication Date: 2015-11-05
    Description: Results indicating that unsteady aerodynamic loads derived under the assumption of simple harmonic motions executed by airfoil or wing can be extended to arbitrary motions are summarized. The generalized Theodorsen (1953) function referable to loads due to simple harmonic oscillations of a wing section in incompressible flow, the Laplace inversion integral for unsteady aerodynamic loads, calculations of root loci of aeroelastic loads, and analysis of generalized compressible transient airloads are discussed.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; Volume 15; Apr. 1977
    Format: text
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  • 33
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    The development and use of spoilers as vortex attenuators (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Results are presented of groundbased and flight investigations performed to develop spoilers as trailing vortex alleviation devices. Based on the results obtained in these investigations, it was found that the induced rolling moment on a trailing model can be reduced by spoilers located near the midsemispan of a vortex generating wing. Substantial reductions in induced rolling moment occur when the spoiler vortex attenuator is located well forward on both unswept and swept wing models. In addition, it was found that existing flight spoilers on the jumbo-jet transport aircraft can be effective as trailing vortex attenuators.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 339-368
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  • 34
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    Unsuccessful Concepts for Aircraft Wake Vortex Minimization (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Exploratory concepts are described which were investigated to achieve a reduction in the vortex induced rolling upsets produced by heavy aircraft trailing vortexes. The initial tests included the use of mass injection, oscillating devices, wingtip shape design, interacting multiple vortexes, and end plates. Although later refinements of some of these concepts were successful, initial test results did not indicate a capability of these concepts to significantly alter the vortex induced rolling upset on a following aircraft.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 221-250
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  • 35
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    Flight test techniques for wake-vortex minimization studies (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Flight test techniques developed for use in a study of wake turbulence and used recently in flight studies of wake minimization methods are discussed. Flow visualization was developed as a technique for qualitatively assessing minimization methods and is required in flight test procedures for making quantitative measurements. The quantitative techniques are the measurement of the upset dynamics of an aircraft encountering the wake and the measurement of the wake velocity profiles. Descriptions of the instrumentation and the data reduction and correlation methods are given.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 193-220
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  • 36
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    Aircraft wake-vortex minimization by use of flaps (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: A survey was made of research on the alleviation of the trailing vortex hazard by altering span loading with flaps on the generator airplane. Flap configurations of the generator that shed multiple vortices were found to have wakes that dispersed by vortex merging and sinusoidal instability. Reductions of approximately 50 percent in both the wake rolling moment imposed on a following aircraft and the aircraft separation requirement were achieved in the ground based and flight test experiments by deflecting the trailing edge flaps more inboard than outboard. Significantly, this configuration did not increase the drag or vibration on the generating aircraft compared to the conventional landing configuration. Ground based results of rolling moment measurement and flow visualization are shown, using a water tow facility, an air tow facility, and a wind tunnel. Flight test results are also shown, using a full scale B-747 airplane. General agreement was found among the results of the various ground based facilities and the flight tests.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 305-338
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  • 37
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    Thrust-augmented vortex attenuation (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: An experimental investigation was conducted to determine the vortex attenuating effect of engine thrust. Tests were made using a 0.03-scale model of the Boeing 747 transport aircraft as a vortex generating model. A Learjet-class probe model was used to measure the vortex induced rolling moment at a scale separation distance of 1.63 km. These tests were conducted at a lift coefficient of 1.4 at a model velocity of 30.48 m/s. The data presented indicate that engine thrust is effective as a vortex attenuating device when the engines are operated at high thrust levels and are positioned to direct the high energy engine wake into the core of the vortex. The greatest thrust vortex attenuation was obtained by operating the inboard engine thrust reversers at one-quarter thrust and the outboard engines at maximum forward thrust.
    Keywords: AERODYNAMICS
    Type: Wake Vortex Minimization; p 251-270
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  • 38
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    Slot injection for skin-friction drag reduction (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A description and analysis of slot injection in low-speed flow, slot injection in high-speed flow, a discussion of aircraft applications, and possibilities for future improvements of slot drag reduction capability are presented.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on Concepts for Drag Reduction; 11 p
    Format: text
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  • 39
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    An overview of concepts for aircraft drag reductions (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A current overview of aerodynamic drag reduction concepts which have potential for reducing aircraft fuel consumption is presented. The discussion shows where the greatest percentages of aircraft fuel is burned and what areas have the greatest potential for fuel conservation. The paper deals with aerodynamic improvements and touches only briefly on structural and propulsion improvements. Concepts for reducing pressure drag (i.e., roughness, wave, interference, and separation drag), drag due to lift/induced drag, and skin-friction drag at subsonic and supersonic speeds are emphasized.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on Concepts for Drag Reduction; 30 p
    Format: text
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  • 40
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    Laminar flow control laminarization (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A practical aerodynamically and structurally reasonably efficient laminar flow control (LFC) suction method, removing the slowest boundary layer particles through many closed spaced fine slots, was developed and subsequently applied to a second F94 LFC wing glove in flight: 100 percent laminar flow was observed up to the F94 test limit. Laminar flow on LFC wings in flight is thus possible at a much higher Reynold's number than even in the best low turbulence tunnels as a result of the negligible influence of the atmospheric microscale turbulence on transition. The F94 LFC glove comparison experiments, with suction starting at 0.03c and 0.4c, verified the theoretically predicted boundary layer stabilization by suction starting at 0.08c, thus maintaining laminar flow at substantially higher C sub L numbers as compared to boundary layer stabilization by flow acceleration; i.e., geometry alone without suction upstream of 0.4c.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on Concepts for Drag Reduction; 75 p
    Format: text
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  • 41
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    Aerodynamic loads near cranks, apexes, and tips of thin, lifting wings in incompressible flow (1977)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: The calculation of the incompressible and irrotational flow in the vicinity of tips and corners of thin, lifting wings is considered. It is shown that the important characteristics of the flow are governed by an eigenvalue problem, which is nonlinear at the trailing edge because of the shed wake (assumed to be in the wing plane). A new solution method was devised because either the existing methods were not valid for the trailing edge case or they would have required excessive amounts of computer time. The new method, which is fundamentally different than the previous ones, was used to calculate solutions for a number of cases, including some for which correct answers had not previously been obtained. Two of these solutions were used to determine the validity of drag and leading-edge-suction distributions near the tips of a delta wing and a swept wing as calculated by using both the vortex lattice method and a kernel function method. The calculations for the swept wing resolved the question of whether or not the induced drag should be zero at the wing tip.
    Keywords: AERODYNAMICS
    Type: AGARD Prediction of Aerodynamic Loading; 12 p
    Format: text
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  • 42
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    Experimental studies of flow separation of the GA(W)-2 airfoil at low speeds (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Wind tunnel tests have been conducted on a NASA GA(W)-2 airfoil section at Reynolds number of 2.2 x 10(exp 6) and Mach number of 0.13. Detailed measurements of flow fields associated with turbulent boundary layers have been obtained at angles of attack of 10.3, 14.4, and 18.3 deg. Pre- and post-separated velocity and pressure survey results over the airfoil and in the associated wake are presented. Extensive force, pressure, tuft survey, hot-film survey, local skin friction, and boundary layer data are also included. Pressure distributions and separation point locations show good agreement with theory for the two lower angles of attack. Boundary layer displacement thickness, momentum thickness, and shape factor agree well with theory up to the point of separation. There is considerable disparity between extent of flow reversal in the wake as measured by pressure and hot-film probes. The difference is attributed to the intermittent nature of the flow reversal.
    Keywords: AERODYNAMICS
    Type: NASA-CR-197254 , NAS 1.26:197254 , AR77-4
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  • 43
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    Computer program to prepare airfoil characteristic data for use in helicopter performance calculations (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A computer program developed to prepare wind tunnel generated airfoil data for input into helicopter performance prediction programs is described. The program provides for numerically cross plotting the data, plotting the data, and tabulating and punching the tabulated result into computer cards for use in the rotorcraft flight simulation model.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78627 , L-11608
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  • 44
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    Two-dimensional wind-tunnel tests of a NASA supercritical airfoil with various high-lift systems. Volume 1: Data analysis (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: High-lift systems for a NASA, 9.3%, method for calculating the viscous flow about two-dimensional multicomponent airfoils was evaluated by comparing its predictions with test data. High-lift systems derived from supercritical airfoils were compared in terms of performance to high-lift systems derived from conventional airfoils. The high-lift systems for the supercritical airfoil were designed to achieve maximum lift and consisted of: a single-slotted flap; a double-slotted flap and a leading-edge slat; and a triple-slotted flap and a leading-edge slat. Agreement between theoretical predictions and experimental results are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2214 , D6-41063-1-VOL-1
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  • 45
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    Two-dimensional wind-tunnel tests of a NASA supercritical airfoil with various high-lift systems. Volume 2: Test data (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Three high lift systems for a 9.3 percent blunt based, supercritical airfoil were designed, fabricated, and wind tunnel tested. A method for calculating the viscous flow about two dimensional multicomponent airfoils was evaluated by comparing its predictions with test data. A comparison of high lift systems derived from supercritical airfoils with high lift systems derived from conventional airfoils is presented. The high lift systems for the supercritical airfoil were designed to achieve maximum lift and consisted of: (1) a single slotted flap, (2) a double slotted flap and a leading edge slat, and (3) a triple slotted flap and a leading edge slat. Aerodynamic force and moment data and surface pressure data are presented for all configurations and boundary layer and wake profiles for the single slotted flap configuration. The wind-tunnel models, test facilities and instrumentation, and data reduction are described.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2215 , D6-41063-2-VOL-2
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  • 46
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    Comparison of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an analysis code (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Comparisons of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an advanced analysis code. Comparisons are made at near-design normal-force coefficients for Reynolds numbers from 2 to 11 million. Comments are made concerning various input parameters to the code.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-74041 , NAS 1.15:X-74041
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  • 47
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    Effect of winglets on a first-generation jet transport wing. 1: Longitudinal aerodynamic characteristics of a semispan model at subsonic speeds (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of winglets and a simple wing-tip extension on the aerodynamic forces and moments and the flow-field cross flow velocity vectors behind the wing tip of a first generation jet transport wing were investigated in the Langley 8-foot transonic pressure tunnel using a semi-span model. The test was conducted at Mach numbers of 0.30, 0.70, 0.75, 0.78, and 0.80. At a Mach number of 0.30, the configurations were tested with combinations of leading- and trailing-edge flaps.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8473 , L-11354
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  • 48
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    Mean Velocity, Turbulence Intensity and Turbulence Convection Velocity Measurements for a Convergent Nozzle in a Free Jet Wind Tunnel. Comprehensive Data Report (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effect of flight on the mean flow and turbulence properties of a 0.056m circular jet were determined in a free jet wind tunnel. The nozzle exit velocity was 122 m/sec, and the wind tunnel velocity was set at 0, 12, 37, and 61 m/sec. Measurements of flow properties including mean velocity, turbulence intensity and spectra, and eddy convection velocity were carried out using two linearized hot wire anemometers. This report contains the raw data and graphical presentations. The final technical report includes a description of the test facilities, test hardware, along with significant test results and conclusions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-135238 , PWA-5516
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  • 49
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    Two-dimensional wind tunnel test of an oscillating rotor airfoil, volume 2 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N78-16003.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2915 , D-210-11188-2
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  • 50
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    Two-dimensional wind tunnel test of an oscillating rotor airfoil, volume 1 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A two dimensional wind tunnel test was conducted to obtain the quasisteady and unsteady characteristics of an advanced airfoil designed for helicopter rotor applications. Differential pressures were measured at 17 locations along the chord of the airfoil model. The airfoil motions were sinusoidal forced-pitch oscillations about the quarter chord at amplitudes varying from 2.5 to 10.0 degrees and at frequencies from 23 Hz to 90 Hz. The quasisteady tests were conducted at Mach numbers from 0.2 to 0.9, and the oscillatory tests between M = 0.2 and M = 0.7. At quasisteady conditions a limited number of drag measurements were made with a wake-traversing probe.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2914 , D210-11188-1-VOL-1
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  • 51
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    A generalized vortex lattice method for subsonic and supersonic flow applications (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: If the discrete vortex lattice is considered as an approximation to the surface-distributed vorticity, then the concept of the generalized principal part of an integral yields a residual term to the vorticity-induced velocity field. The proper incorporation of this term to the velocity field generated by the discrete vortex lines renders the present vortex lattice method valid for supersonic flow. Special techniques for simulating nonzero thickness lifting surfaces and fusiform bodies with vortex lattice elements are included. Thickness effects of wing-like components are simulated by a double (biplanar) vortex lattice layer, and fusiform bodies are represented by a vortex grid arranged on a series of concentrical cylindrical surfaces. The analysis of sideslip effects by the subject method is described. Numerical considerations peculiar to the application of these techniques are also discussed. The method has been implemented in a digital computer code. A users manual is included along with a complete FORTRAN compilation, an executed case, and conversion programs for transforming input for the NASA wave drag program.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2865 , LR-28112
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  • 52
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    Longitudinal aerodynamic characteristics at Mach 0.60 to 2.86 of a fighter configuration with strut braced wing (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was made to determine the effects on longitudinal aerodynamic characteristics of utilizing struts to brace the wing to allow the wing thickness reduction on the LFAX-8 fighter configuration. Structural and load analysis indicated that the maximum airfoil thickness could be reduced from 4.5 to 3.1 percent with the strut brace concept. Wave drag theory indicated that reducing the wing maximum thickness from 4.5 percent to 3.1 percent would yield a significant reduction in zero-lift wave drag of about 28 percent at the design Mach number of 1.60. Strut arrangements designed and tested included, a single straight strut, a single swept strut, and a set of tandem straight struts. In addition, a wire of approximately the same cross sectional area replaced the single straight strut on one series of runs. The original LFAX-8 with the 4.5-percent-thick wing was retested to serve as a base line for this investigation.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1102 , L-11801
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  • 53
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    Aerodynamic characteristics of a monoplanar missile concept with bodies of circular and elliptical cross sections (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Experimental aerodynamic characteristics of a low-drag missile concept with a body of circular cross section were compared to one with a body of 3:1 elliptical cross section, the bodies having identical cross section area distributions. The concepts were of monowing design with constant wing span. Tail surfaces were located flush at the body base with plus or minus 30 deg dihedral. Wind tunnel tests were performed at Mach numbers from 0.5 to 4.63 and at angles of attack from about -5 deg to 28 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74079 , L-11423
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  • 54
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    Improvement of maneuver aerodynamics by spanwise blowing (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Spanwise blowing was used to test a generalized wind-tunnel model to investigate component concepts in order to provide improved maneuver characteristics for advanced fighter aircraft. Primary emphasis was placed on performance, stability, and control at high angles of attack and subsonic speeds. Test data were obtained in the Langley high speed 7 by 10 foot tunnel at free stream Mach numbers up to 0.50 for a range of model angles of attack, jet momentum coefficients, and leading and trailing edge flap deflection angles. Spanwise blowing on a 44 deg swept trapezoidal wing resulted in leading edge vortex enhancement with subsequent large vortex induced lift increments and drag polar improvements at the higher angles of attack. Small deflections of a leading edge flap delayed these lift and drag benefits to higher angles of attack. In addition, blowing was more effective at higher Mach numbers. Spanwise blowing in conjunction with a deflected trailing edge flap resulted in lift and drag benefits that exceeded the summation of the effects of each high lift device acting alone. Asymmetric blowing was an effective lateral control device at the higher angles of attack.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1065 , L-11642
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  • 55
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    Aerodynamic characteristics of a small-scale straight and swept-back wing with knee-blown jet flaps (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two sting-mounted, 50.8 cm (20 in.) span, knee-blown, jet-flap models were tested in a large (2.1- by 2.5-m (7- by 10-ft) subsonic wind tunnel. A straight- and swept-wing model were tested with fixed flap deflection with various combinations of full-span leading-edge slats. The swept-wing model was also tested with wing tip extensions. Data were taken at angles-of-attack between 0 deg and 40 deg, at dynamic pressures between 143.6 N/sq m (3 lb/sq ft) and 239.4 N/sq m (5 lb/sq ft), and at Reynolds numbers (based on wing chord) ranging from 100,000 to 132,000. Jet flap momentum blowing coefficients up to 10 were used. Lift, drag, and pitching-moment coefficients, and exit flow profiles for the flap blowing are presented in graphical form without analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78427 , A-7161
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  • 56
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    Computation of transonic boattail flow with separation (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The relaxation procedure of South and Jameson for the full potential transonic flow equation was coupled to a modified Reshotko-Tucker integral boundary-layer technique with an empirical model for separated flow. The viscous and inviscid flows were solved iteratively until convergence was obtained. This iterative method was then applied to the subsonic and transonic flow over a series of axisymmetric circular-arc boattails with solid jet plume simulators. Comparisons of theoretical and experimental surface pressures and boattail drag are presented over a free-stream Mach numbers below 0.90. The qualitative variation of boattail drag with free-stream Mach number and boattail angle well into the region of transonic drag rise was correctly predicted; however, the absolute drag levels were significantly underpredicted. For separated flows, the empirical discriminating streamline model gives good results up to a free-stream Mach number of about 0.90 and allows reasonable predictions for shock-induced separation if the proper separation location and separation turning angle are known.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1070 , L-11669
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  • 57
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    Effects of roughness size on the position of boundary-layer transition and on the aerodynamic characteristics of a 55 deg. swept delta wing at supersonic speeds (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted to determine the effects of roughness size on the position of boundary layer transition and on the aerodynamic characteristics of a 55 deg swept delta wing model. Results are presented and discussed for wind tunnel tests conducted at free stream Mach numbers from 1.50 to 4.63, Reynolds numbers per meter from 3,300,000 to 1.6 x 10 to the 7th power, angles of attack from -8 to 16 deg, and roughness sizes ranging from 0.027 cm sand grit to 0.127 cm high cylinders. Comparisons were made with existing flat plate data. An approximate method was derived for predicting the drag of roughness elements used in boundary layer trips.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1027 , L-11496
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  • 58
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    Theoretical and experimental analysis of longitudinal and lateral aerodynamic characteristics of skewed wings at subsonic speeds to high angles of attack (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of sweep and aspect ratio on the longitudinal and lateral-directional aerodynamic characteristics of low-aspect-ratio skewed (oblique) wings having separation-induced vortex flows along leading and side edges were investigated in the Langley high-speed 7- by 10-foot tunnel at a low-subsonic Mach number. The theoretical analysis used the vortex-lattice method for estimating attached-flow aerodynamic characteristics and the leading-edge suction analogy of Polhamus for estimating separation induced vortex-flow effects. Experimental results were compared with asymmetric, separated, vortex flow theory.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8512 , L-11230
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  • 59
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    Effects of deflected thrust on the longitudinal aerodynamic characteristics of a close-coupled wing-canard configuration (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of power on the longitudinal aerodynamic characteristics of a close-coupled wing-canard fighter configuration with partial-span rectangular nozzles at the trailing edge of the wing were investigated. Data were obtained on a basic wing-strake configuration for nozzle and flap deflections from 0 deg to 30 deg and for nominal thrust coefficients from 0 to 0.30. The model was tested over an angle-of-attack range from -2 deg to 40 deg at Mach numbers of 0.15 and 0.18. Results show substantial improvements in lift-curve slope, in maximum lift, and in drag-due-to-lift efficiency when the canard and strakes have been added to the basic wing-fuselage (wing-alone) configuration. Addition of power increased both lift-curve slope and maximum lift, improved longitudinal stability, and reduced drag due to lift on both the wing-canard and wing-canard-strake configurations. These beneficial effects are primarily derived from boundary-layer control due to moderate thrust coefficients which delay flow separation on the nozzle and inboard portion of the wing flaps.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1090 , L-11886
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  • 60
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    Investigation of compressible vortex flow characteristics (1977)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: The nature of intense air vortices was studied and the factors which determine the intensity and rate of decay of both single and pairs of vortices were investigated. Vortex parameters of axial pressure differential, circulation, outflow rates, separation distance and directions of rotation were varied. Unconfined vortices, generated by a single rotating cage, were intensified by an increasing axial pressure gradient. Breakdown occurred when the axial gradient became negligible. The core radius was a function of the axial gradient. Dual vortices, generated by two counterrotating cages, rotated opposite to the attached cages. With minimum spacing only one vortex was formed which rotated in a direction opposite to the attached cage. When one cage rotated at half the speed of the other cage, one vortex formed at the higher speed cage rotating in the cage direction.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145261 , CRINC-FRL-260-1
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    Subsonic longitudinal and lateral directional static stability characteristics of a variable sweep fighter configuration employing various control surface deflections at angles of attack of 0 deg to 50 deg and large sideslip angles (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A variable sweep fighter configuration with the wing in the 22 deg sweep position having leading edge slats and trailing edge flaps on the outboard panels was tested at a Mach number of 0.15 in the Langley 7- by 10-foot high speed tunnel. The angle of attack range was 0 deg to 50 deg and the sideslip angle range was -20 deg to 20 deg. Pitch, roll, and yaw control effectiveness were studied as well as the effects of spoilers. The data are presented without analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74050
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  • 62
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    Computations of supersonic viscous flow over a finite-width plate (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Finite difference methods were applied to solve the parabolic Navier-Stokes equations for the flow over a finite width plate at 0 deg and 10 deg angles of attack. The methods were developed on the basis of the operator factorization concept resulting in the split of a three dimensional equation into successive two dimensional equations. Backward and centered implicit factorization schemes, were used and their results were compared. Available numerical solutions and experimental data obtained at low Reynolds number conditions were also used for comparison. The backward implicit method provides a more successful solution, which ranges from the merged layer to the strong interaction regimes. Detailed structures were revealed of the shear layer around and behind the side edge.
    Keywords: AERODYNAMICS
    Type: NASA-CR-151534 , JSC-13140 , TR-7004
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  • 63
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    Ground simulation and tunnel blockage for a swept, jet-flapped wing tested to very high lift coefficients (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Ground effects experiments and large/small tunnel interference studies were carried out on a model with a 20 inch (50.8 cm) 25 degree swept wing. The wing is slatted, has a 60 degree knee-blown flap and can be fitted with unflapped tips. A tail rake of pitch-yaw probes can be fitted to the fuselage. Certain check tests were also made with a very similar straight-wing model.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152032
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  • 64
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    Supersonic aerodynamic characteristics of a Sparrow 3 type missile model with wing controls and comparison with existing tail-control results (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted on a model of a wing control version of the Sparrow III type missile to determine the static aerodynamic characteristics over an angle of attack range from 0 deg to 40 deg for Mach numbers from 1.50 to 4.60.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1078 , L-11715
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  • 65
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    Spin-tunnel investigation of the spinning characteristics of typical single-engine general aviation airplane designs. 2: Low-wing model A; tail parachute diameter and canopy distance for emergency spin recovery (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A spin tunnel study is reported on a scale model of a research airplane typical of low-wing, single-engine, light general aviation airplanes to determine the tail parachute diameter and canopy distance (riser length plus suspension-line length) required for energency spin recovery. Nine tail configurations were tested, resulting in a wide range of developed spin conditions, including steep spins and flat spins. The results indicate that the full-scale parachute diameter required for satisfactory recovery from the most critical conditions investigated is about 3.2 m and that the canopy distance, which was found to be critical for flat spins, should be between 4.6 and 6.1 m.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1076 , L-11804
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  • 66
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    An experimental and analytical investigation of proprotor whirl flutter (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of an experimental parametric investigation of whirl flutter are presented for a model consisting of a windmilling propeller-rotor, or proprotor, having blades with offset flapping hinges mounted on a rigid pylon with flexibility in pitch and yaw. The investigation was motivated by the need to establish a large data base from which to assess the predictability of whirl flutter for a proprotor since some question has been raised as to whether flutter in the forward whirl mode could be predicted with confidence. To provide the necessary data base, the parametric study included variation in the pylon pitch and yaw stiffnesses, flapping hinge offset, and blade kinematic pitch-flap coupling over a large range of advance ratios. Cases of forward whirl flutter and of backward whirl flutter are documented. Measured whirl flutter characteristics were shown to be in good agreement with predictions from two different linear stability analyses which employed simple, two dimensional, quasi-steady aerodynamics for the blade loading. On the basis of these results, it appears that proprotor whirl flutter, both forward and backward, can be predicted.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1047 , L-11656
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  • 67
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    Performance of twin two-dimensional wedge nozzles including thrust vectoring and reversing effects at speeds up to Mach 2.20 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Transonic tunnel and supersonic pressure tunnel tests were reformed to determine the performance characteristics of twin nonaxisymmetric or two-dimensional nozzles with fixed shrouds and variable-geometry wedges. The effects of thrust vectoring, reversing, and installation of various tails were also studied. The investigation was conducted statically and at flight speeds up to a Mach number of 2.20. The total pressure ratio of the simulated jet exhaust was varied up to approximately 26 depending on Mach number. The Reynolds number per meter varied up to 13.20 x 1 million. An analytical study was made to determine the effect on calculated wave drag by varying the mathematical model used to simulate nozzle jet-exhaust plume.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8449 , L-11277
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  • 68
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    Aerodynamic characteristics at Mach numbers from 0.33 to 1.20 of a wing-body design concept for a hypersonic research airplane (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation of the static aerodynamic characteristics of a model of one design concept for the proposed National Hypersonic Flight Research Facility was conducted in the Langley 8 foot transonic pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, and a three module or six module scramjet engine. The freestream test Mach numbers were 0.33, 0.80, 0.90, 0.95, 0.98, 1.10, and 1.20 at Reynolds numbers per meter ranging from 4.8 x 1 million to 10.4 x 1 million. The test angle of attack range was approximately -4 deg to 22 deg at constant angles of sideslip of 0 deg and 4 deg; the angle of sideslip ranged from about -6 deg to 6 deg at constant angles of attack of 0 deg and 17 deg. The elevons were deflected 0 deg, -10 deg, and -20 deg with rudder deflections of 0 deg and 15.6 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1044 , L-11723
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  • 69
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    On the logarithmic-singularity correction in the kernel function method of subsonic lifting-surface theory (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A logarithmic-singularity correction factor is derived for use in kernel function methods associated with Multhopp's subsonic lifting-surface theory. Because of the form of the factor, a relation was formulated between the numbers of chordwise and spanwise control points needed for good accuracy. This formulation is developed and discussed. Numerical results are given to show the improvement of the computation with the new correction factor.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8513 , L-11142
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  • 70
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    Computation of unsteady transonic flows through rotating and stationary cascades. 2: User's guide to FORTRAN program B2DATL (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Documentation for the FORTRAN program B2DATL is provided. The program input, output, and operational procedures are described; a dictionary of the principal FORTRAN variables is provided; the function of all subroutines; is outlined and flow charts of the principal subroutines and the main program are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2901 , ATL-TR-205-VOL-2
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  • 71
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    Computation of unsteady transonic flows through rotating and stationary cascades. 3: Acoustic far-field analysis (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A small perturbation type analysis has been developed for the acoustic far field in an infinite duct extending upstream and downstream of an axial turbomachinery stage. The analysis is designed to interface with a numerical solution of the near field of the blade rows and, thereby, to provide the necessary closure condition to complete the statement of infinite duct boundary conditions for the subject problem. The present analysis differs from conventional inlet duct analyses in that a simple harmonic time dependence was not assumed, since a transient signal is generated by the numerical near-field solution and periodicity is attained only asymptotically. A description of the computer code developed to carry out the necessary convolutions numerically is included, as well as the results of a sample application using an impulsively initiated harmonic signal.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2902 , ATL-TR-205-VOL-3
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  • 72
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    Longitudinal aerodynamic characteristics at Mach numbers from 1.60 to 2.86 for a fixed-span missile with three wing planforms (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Effects of wing planform modifications on the longitudinal aerodynamic characteristics of a fixed span, maneuverable cruciform missile configuration were determined. A basic delta planform and two alternate trapezoidal planforms having progressively increasing tip chords were included. Data were obtained for angles of attack up to approximately -32 deg, model roll angles of 0 deg to 45 deg, and tail control deflections of 0 deg and -20 deg. The experimental drag due to lift was compared with linear values.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74088 , L-11916
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  • 73
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    Supersonic longitudinal aerodynamic characteristics of two space shuttle orbiter configurations (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted to determine the supersonic longitudinal aerodynamic characteristics of 0.015 scale models of the Rockwell International 089B and 139B space shuttle orbiter configurations and the 139B orbiter with a modifier forebody. The models each had a 45 deg swept delta wing that was blended into the body with an 81 deg swept fillet to form a double delta planform. The vertical tail had a split rudder deflected 27.5 deg on each side to form a speed brake. Tests were conducted at Mach numbers of 2.5, 3.9, and 4.6 at a Reynolds number, based on the body length of the 089B model, of 4,150,000. Angles of attack varied from -4 deg to 44 deg at 0 deg sideslip.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74074 , L-11787
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  • 74
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    Viscous effects in aircraft trailing vortices (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The mechanism of merging of like-signed aircraft vortices leading to a rapid redistribution of trailed vorticity in a wake through both convective and turbulent processes was investigated. Research was done experimentally in a small wind tunnel and analytically through the use of a code which computes turbulent transport using a second-order closure turbulent model. Computations are reported which demonstrate the merging phenomenon, and comparisons are made with experimental results. The usefulness of point vortex computations in predicting merging was explored. Limited computations showed that jet exhaust does not appreciably alter the merging phenomenon. The effect of ambient atmospheric turbulence on the aging of an aircraft wake was investigated at a constant turbulent dissipation rate. It was shown that under stable atmospheric conditions, when atmospheric macroscales are less than or equal to the vortex spacing, misleading results may be obtained.
    Keywords: AERODYNAMICS
    Type: NASA, Washington Wake Vortex Minimization; p 61-128
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  • 75
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    Supersonic separated turbulent boundary - layer over a wavy wall (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A prediction method is developed for calculating distributions of surface heating rates, pressure and skin friction over a wavy wall in a two-dimensional supersonic flow. Of particular interest is the flow of thick turbulent boundary layers. The surface geometry and the flow conditions considered are such that there exists a strong interaction between the viscous and inviscid flow. First, using the interacting turbulent boundary layer equations, the problem is formulated in physical coordinates and then a reformulation of the governing equations in terms of Levy-Lees variables is given. Next, a numerical scheme for solving interacting boundary layer equations is adapted. A number of modifications which led to the improvement of the numerical algorithm are discussed. Finally, results are presented for flow over a train of up to six waves at various flow conditions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-155329 , AFL-77-11-36
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  • 76
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    A technique for simulating turbulence for aerospace vehicle flight simulation studies (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An atmospheric turbulence model which accommodates variability of turbulence properties along an aerospace vehicle trajectory was developed. The technique involves the use of Dryden spectral forms in which the defining parameters are the standard deviations (sigma) and integral scales (L) of turbulence. These spectra are expressed as nondimensional functions of the nondimensional frequency Omega = omega L/V where omega is dimensional radian frequency and V is the true air speed of the aerospace vehicle. The nondimensional spectra are factored by standard techniques to obtain nondimensional linear recursive filters in the time domain whereby band-limited white-like noise can be operated upon to obtain nondimensional longitudinal, lateral, and vertical turbulence velocities, as functions of nondimensional time, tV/L, where t is time. Application of the technique to the simulation of the space shuttle orbiter entry flight phase is discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78141
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  • 77
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    A user's guide for V174, a program using a finite difference method to analyze transonic flow over oscillating wings (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design and usage of a pilot program using a finite difference method for calculating the pressure distributions over harmonically oscillating wings in transonic flow are discussed. The procedure used is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The steady velocity potential which must be obtained from some other program, is required for input. The unsteady differential equation is linear, complex in form with spatially varying coefficients. Because sinusoidal motion is assumed, time is not a variable. The numerical solution is obtained through a finite difference formulation and a line relaxation solution method.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145214
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  • 78
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    Static aerodynamic characteristics of a 0.035-scale model of a modified NKC-135 airplane at a Mach number of 0.28 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 0.035-scale model fo a modified NKC-135 airplane was tested in 12-foot pressure wind tunnel to determine the effects on the static aerodynamic characteristics of modifications to the basic aircraft. Modifications investigated included: nose, lower fuselage, and upper fuselage radomes; wing pylons and pods; overwing probe; and air conditioning inlets. The investigation was performed at a Mach number of 0.28 over a Reynolds number range from 6.6 to 26.2 million per meter. Angles of attack and sideslip varied from -8 deg to 20 deg and from -18 deg to 8 deg, respectively, for various combinations of flap, aileron, and rudder deflections. A limited analysis of the test results indicates that the addition of the radomes reduces lateral-directional stability and control effectiveness of the basic aircraft.
    Keywords: AERODYNAMICS
    Type: NASA-TM-73250 , A-7068
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  • 79
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    Effect of coolant flow ejection on aerodynamic performance of low-aspect-ratio vanes. 2: Performance with coolant flow ejection at temperature ratios up to 2 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aerodynamic performance of a 0.5 aspect ratio turbine vane configuration with coolant flow ejection was experimentally determined in a full annular cascade. The vanes were tested at a nominal mean section ideal critical velocity ratio of 0.890 over a range of primary to coolant total temperature ratio from 1.0 to 2.08 and a range of coolant to primary total pressure ratio from 1.0 to 1.4 which corresponded to coolant flows from 3.0 to 10.7 percent of the primary flow. The variations in primary and thermodynamic efficiency and exit flow conditions with circumferential and radial position were obtained.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1057 , E-9213
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  • 80
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    Experimental performance of a 13.65-centimeter-tip-diameter tandem-bladed sweptback centrifugal compressor designed for a pressure ratio of 6 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 13.65 cm tip diameter backswept centrifugal impeller having a tandem inducer and a design mass flow rate of 0.907 kg/sec was experimentally investigated to establish stage and impeller characteristics. Tests were conducted with both a cascade diffuser and a vaneless diffuser. A pressure ratio of 5.9 was obtained near surge for the smallest clearance tested. Flow range at design speed was 6.3 percent for the smallest clearance test. Impeller exit to shroud axial clearance at design speed was varied to determine the effect on stage and impeller performance.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1091
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  • 81
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    Wind tunnel investigation of rotor lift and propulsive force at high speed. Test data appendix (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N78-10020.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145217-APP-2 , D210-11135-2
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  • 82
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    Splined version of FLEXSTAB: A critical analysis of alternate schemes (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Recommendations are made for improved aerodynamic models and numerical schemes to be considered for inclusion into the FLEXSTAB computer program system. These recommendations are based on a critical analysis of numerical technology.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152030
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  • 83
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    An Exploratory Investigation of the Effects of a Thin Plastic Film Cover on the Profile Drag of an Aircraft Wing Panel (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Exploratory wind tunnel tests were conducted on a large chord aircraft wing panel to evaluate the potential for drag reduction resulting from the application of a thin plastic film cover. The tests were conducted at a Mach number of 0.15 over a Reynolds number range from about 7 x 10 to the 6th power to 63 x 10 to the 6th power.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74073
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  • 84
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    Unsteady aerodynamic modeling and active aeroelastic control (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Unsteady aerodynamic modeling techniques are developed and applied to the study of active control of elastic vehicles. The problem of active control of a supercritical flutter mode poses a definite design goal stability, and is treated in detail. The transfer functions relating the arbitrary airfoil motions to the airloads are derived from the Laplace transforms of the linearized airload expressions for incompressible two dimensional flow. The transfer function relating the motions to the circulatory part of these loads is recognized as the Theodorsen function extended to complex values of reduced frequency, and is termed the generalized Theodorsen function. Inversion of the Laplace transforms yields exact transient airloads and airfoil motions. Exact root loci of aeroelastic modes are calculated, providing quantitative information regarding subcritical and supercritical flutter conditions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-148019 , SUDAAR-504
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  • 85
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    Identification of state variables and dynamic inflow from rotor model dynamic tests (1977)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: The paper describes methods for extracting unknown state variables and parameters from dynamic rotor model tests given transient cyclic pitch stirring inputs, blade root flap-bending measurements, and the form of the dynamic rotor equations, including a rotor dynamic inflow description, when none of the physical parameters are known. A simplified version of the maximum likelihood method seems best suited for this purpose. The measurement equation error covariance matrix is assumed constant during each iteration, but updated for the subsequent iteration. A detailed analysis of the suitability of the derived techniques for studying various rotor dynamic inflow effects is provided.
    Keywords: AERODYNAMICS
    Type: American Helicopter Society; vol. 22
    Format: text
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  • 86
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    A numerical study of the unsteady leading edge separation bubble on an oscillating airfoil (1977)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: A preliminary study of the unsteady viscous flow in the region of an airfoil leading edge was performed, in which the interaction between the viscous and inviscid flow fields is neglected. The solution method uses the finite difference form of the governing equations throughout the separated flow field and incorporates a transition model based on the integral turbulence kinetic energy equation. The validity of the numerical procedure is verified by making comparisons with analytical solutions to several test problems, including unsteady flow over a plate oscillating in its own plane. The method was then applied to the problem of unsteady viscous flow over a NACA 0012 airfoil oscillating sinusoidally in pitch. The flow field characteristics were in qualitative agreement with experimental results. The bubble moved forward on the airfoil and decreased in size as incidence was increased. Viscous flow in the leading edge region was found to be quasi-steady, while bubble height varied inversely with Reynolds number.
    Keywords: AERODYNAMICS
    Type: Computer Methods in Applied Mechanics and Engineering; 11; Apr. 197
    Format: text
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  • 87
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    Effect of Winglets on a First-Generation Jet Transport Wing. 2: Pressure and Spanwise Load Distributions for a Semispan Model at High Subsonic Speeds (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Pressure and spanwise load distributions on a first-generation jet transport semispan model at high subsonic speeds are presented for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. Selected data are discussed to show the general trends and effects of the various configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8474 , L-11026
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  • 88
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    Preliminary study of the use of the STAR-100 computer for transonic flow calculations (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An explicit method for solving the transonic small-disturbance potential equation is presented. This algorithm, which is suitable for the new vector-processor computers such as the CDC STAR-100, is compared to successive line over-relaxation (SLOR) on a simple test problem. The convergence rate of the explicit scheme is slower than that of SLOR, however, the efficiency of the explicit scheme on the STAR-100 computer is sufficient to overcome the slower convergence rate and allow an overall speedup compared to SLOR on the CYBER 175 computer.
    Keywords: AERODYNAMICS
    Type: NASA-TM-74086
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  • 89
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    Effect of a simulated engine jet blowing above an arrow wing at Mach 2.0 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of a gas jet simulating a turbojet engine exhaust blowing above a cambered and twisted arrow wing were investigated. Tests were conducted in the Langley 4-foot supersonic pressure tunnel at a Mach number of 2.0. Nozzle pressure ratios from 1 to 64 were tested with both helium and air used as jet gases. The tests were conducted at angles of attack from -2 deg to 8 deg at a Reynolds number of 9,840,000 per meter. Only the forces and moments on the wing were measured. Results of the investigation indicated that the jet blowing over the wing caused reductions in maximum lift-drag ratio of about 4 percent for helium and 6 percent for air at their respective design nozzle pressure ratios, relative to jet-off data. Moderate changes in the longitudinal, vertical, or angular positions of the jet relative to the wing had little effect on the wing aerodynamic characteristics.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1050 , L-11751
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  • 90
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    Finite element analysis of low speed viscous and inviscid aerodynamic flows (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A weak interaction solution algorithm was established for aerodynamic flow about an isolated airfoil. Finite element numerical methodology was applied to solution of each of differential equations governing potential flow, and viscous and turbulent boundary layer and wake flow downstream of the sharp trailing edge. The algorithm accounts for computed viscous displacement effects on the potential flow. Closure for turbulence was accomplished using both first and second order models. The COMOC finite element fluid mechanics computer program was modified to solve the identified equation systems for two dimensional flows. A numerical program was completed to determine factors affecting solution accuracy, convergence and stability for the combined potential, boundary layer, and parabolic Navier-Stokes equation systems. Good accuracy and convergence are demonstrated. Each solution is obtained within the identical finite element framework of COMOC.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2908 , COMOC-77TR-2
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  • 91
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    Effects of canard location on the aerodynamic characteristics of a sharp-nosed missile at Mach numbers of 1.5 and 2.0 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A sharp-nosed missile model with nose-mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-foot wind tunnel to determine the contributions of the component aerodynamic surfaces on the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and a Reynolds number of 1,000,000 based on body diameter. Effect of canard location (at three different axial positions) was investigaed. Data were obtained at angles of attack ranging from 3 deg to 12 deg, roll angles of 0 deg and 45 deg, and canard-deflection angles from 3 deg to 15 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). The canard in the most aft position was least effective in trimming the model. With movement of the canard to the most forward position, the model was trimmable at angles of attack up to about 5 deg or 6 deg with panel deflections of 9 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TM-73221 , A-6959
    Format: application/pdf
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  • 92
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    Wind tunnel investigation of rotor lift and propulsive force at high speed: Test data appendix (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N78-10020.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145217-APP-3 , D210-11135-3-APP
    Format: application/pdf
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  • 93
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    Tetroon evaluation program (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The actual volume of a constant volume superpressured tetrahedron shaped balloon changes as the amount of superpressure is changed. The experimental methods used to measure these changes in volume are described and results are presented. The basic equations used to determine the amount of inflation gas required for a tetroon to float at a predetermined flight level are presented and inflation techniques discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-150437
    Format: application/pdf
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  • 94
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    Effects of lengthwise lift distribution on sonic boom of SST configurations (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Sonic Boom signatures produced by possible SST configurations during cruise were investigated. It is shown that optimization based on a far field analysis is not necessarily the optimum for these conditions. For an airplane length of 300 ft, near-field effects can be obtained when sufficient lift is generated near the nose of the airplane. Because of the near-field effects, sonic booms with maximum overpressures of the order of 1 lb/square foot can be obtained with possible airplane configurations having the same flight conditions at cruise.
    Keywords: AERODYNAMICS
    Type: Selected Papers on Advanced Design of Air Vehicles; p 51-54
    Format: text
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  • 95
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    General purpose computer program for interacting supersonic configurations. User's manual (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The input data required to execute the computer program ISCON are described. The program generates a numerical procedure for the determination of unsteady aerodynamic forces on arbitrarily interacting wings and tails in supersonic flow. A velocity potential gradient method is used. Constant Mach number is assumed throughout the flow field. Lifting surfaces are represented by trapezoidal elements which can be generated automatically by the program. The wake field is represented by rectangular strip elements. The formulation is reviewed as well as input overview and input format. Instruction on how to use ISCON, a sample problem, and the restart feature are discussed. Program size limitations, computer program flow, and error messages are also included along with a description of the SS31 program used to compute the coefficients of surface spline.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145128
    Format: application/pdf
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  • 96
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    Numerical solutions for viscous and potential flow about arbitrary two-dimensional bodies using body-fitted coordinate systems (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: A procedure for numerical solution of the time-dependent, incompressible Navier-Stokes equations for the flow about arbitrarily shaped two-dimensional bodies is given. This solution is based on a technique of automatic numerical generation of a curvilinear coordinate system having a coordinate line coincident with the body contour regardless of its shape. The implicit solution utilizes the vorticity-stream function formulation with a false-position iterative adjustment of the surface vorticity in satisfaction of the no-slip boundary condition. Excellent agreement with the Blasius boundary layer solution is obtained for a semi-infinite flat plate. Results are presented for Reynolds numbers up to 2000 for several airfoils and a cambered rock.
    Keywords: AERODYNAMICS
    Type: Journal of Computational Physics; 24; July 197
    Format: text
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  • 97
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    Compressible laminar boundary layers with suction on swept and tapered wings (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: In this paper we present a numerical method for solving the compressible laminar boundary-layer equations with suction on swept and tapered wings. The method employs an efficient two-point finite-difference method to solve the governing equations, and a very convenient similarity transformation which removes the wall normal velocity as a boundary condition and places it into the governing equations as a parameter. In this way the awkward nonlinear boundary condition which couples all the variables is avoided. To test and demonstrate the method, we present a sample calculation for a typical laminar-flow-control (LFC) wing.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; July 197
    Format: text
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  • 98
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    Load distributions on slender delta wings having vortex flow (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: Results obtained through application of potential-flow theory based on a suction flow analogy for incompressible flow are compared to attached-flow theory and to experimental data acquired for sharp-edged delta wings. The error incurred in predicting structural loads on the basis of attached-flow theory is investigated. Force and moment coefficients and load distributions are used in the comparisons. The root bending moment can be estimated by assuming vortex lift concentrated at the wing leading edge. While this may entail some error at large angles of attack (owing to inboard movement of the center of vortex lift), it may furnish an upper bound in evaluating results based on other theories.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 14; July 197
    Format: text
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  • 99
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    Cold air performance of a 12.766-centimeter-tip-diameter axial-flow cooled turbine. 2: Effect of air ejection on turbine performance (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An air cooled version of a single-stage, axial-flow turbine was investigated to determine aerodynamic performance with and without air ejection from the stator and rotor blades surfaces to simulate the effect of cooling air discharge. Air ejection rate was varied from 0 to 10 percent of turbine mass flow for both the stator and the rotor. A primary-to-air ejection temperature ratio of about 1 was maintained.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1018
    Format: application/pdf
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  • 100
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    Study of design and analysis methods for transonic flow (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An airfoil design program and a boundary layer analysis were developed. Boundary conditions were derived for ventilated transonic wind tunnels and performing transonic windtunnel wall calculations. A computational procedure for rotational transonic flow in engine inlet throats was formulated. Results and conclusions are summarized.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152041 , NOTE-118
    Format: application/pdf
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