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  • 1
    Publication Date: 2009-11-23
    Keywords: AERODYNAMICS
    Type: NAS 1.26:184996 , NASA-CR-184996
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  • 2
    facet.materialart.
    In:  CASI
    Publication Date: 2009-11-16
    Description: Aeroelasticity and unsteady flow problems of aircraft
    Keywords: AERODYNAMICS
    Type: PERFORMANCE AND DYN. OF AEROSPACE VEHICLES 1971; P 289-374
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  • 3
    Publication Date: 2011-08-24
    Description: An analytic solution of the thickness problem of a rectangular wing with parabolic airfoil section in three-dimensional flow is presented. The free-air solution is obtained by integrating the equation of the axial perturbation velocity. The Prandtl-Glauert rule can be used to derive the subsonic solution. Parts of the free-air solution are verified by taking the limit of the axial peturbation velocity on the model surface as the wing span goes to infinity. Pressure coefficients are studied for a selected wing geometry in the flow field. The solution of the thickness problem of a rectangular wing in a rectangular wind tunnel is derived using the free-air solution and the method of images.
    Keywords: AERODYNAMICS
    Type: ; : Algorithmic trends
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  • 4
    Publication Date: 2011-08-24
    Description: Unsteady flowfields around oscillating Boeing VR7 airfoil with and without a leading-edge slat were numerically investigated by a novel zonal method using a conformal mapping technique. Numerical aero-dynamic hysteresis loops show that the leading-edge slat prevents the airfoil dynamic stall at reduced frequency of 0.15, Reynolds number of 1 million, and the oscillation range of 5 deg to 25 deg.
    Keywords: AERODYNAMICS
    Type: ; : Algorithmic trends
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  • 5
    Publication Date: 2011-08-24
    Description: Fixed-wing code development is now aimed primarily at the solution of problems dominated by separation--based on the assumptions that the ability to solve such problems implies the ability to solve all other problems and that present inviscid method are already adequate for most other problems. Neither of the above assumptions are correct for rotary wing problems. This is because of the unique and overriding importance of wake modeling to rotor problems and also due to the well-known numerical diffusion problems which convectional Eulerian Computational fluid dynamics (CFD) method encounter when called on to convect strong vortical regions for long distances. The need for accurate wake analyses is probably the most fundamental difference between rotory and fixed-wing aerodynamics. In addition, rotary wing complexity requires a much more intimate relationship between test and analysis than is common in fixed-wing work. With these issues in mind, this paper will review some of our recent experience in using a unique-Eulerian-Lagrangian Computational fluid dynamics (CFC) method for the solution of a critical rotor-wake problem--the prediction of hover performance.
    Keywords: AERODYNAMICS
    Type: ; : Algorithmic trends
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  • 6
    Publication Date: 2011-08-24
    Description: The use of the External propulsion Accelerator (EPA) for launching models of hypersonic aerodynamic configurations into an instrumented ballistic range is discussed. The aerodynamic model is encased inside an axisymmetric projectile designed to be accelerated to high speed in the EPA. Accelerator lengths required to achieve hypersonic speeds are estimated to vary from 10 meters for Mach 7, 40 meters for Mach 10, 150 meters for Mach 15, and 700 meters for Mach 30, assuming a limit of 50,000 g's acceleration. For a model span of 10 cm to 25 cm, the launch tube diameters are 40 cm and 100 cm, respectively. Using this EPA launcher will enable exact simulation of hypersonic flight in ground facilities where both the gas composition and pressure can be controlled in the ballistic range.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 95-6138 , AIAA, Aerospace Planes and Hypersonics Technologies Conference; Apr. 3-7, 1995; Chattanooga, TN; United States|; 5 p.
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  • 7
    Publication Date: 2011-08-24
    Description: A method for analyzing the mutual aerodynamic interaction between a rotor and an airframe model has been developed. This technique models the rotor implicitly through the source terms of the momentum equations. A three-dimensional, incompressible, laminar, Navier-Stokes solver in cylindrical coordinates was developed for analyzing the rotor-airframe problem. The calculations are performed on a simplified rotor-airframe model at an advance ratio of 0.1. The airframe surface pressure predictions are found to be in good agreement with wind tunnel test data. Results are also presented for velocity and pressure field distributions in the wake of the rotor.
    Keywords: AERODYNAMICS
    Type: American Helicopter Society, Journal (ISSN 0002-8711); 40; 2; p. 57-67
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  • 8
    Publication Date: 2011-08-24
    Description: An assessment is given of existing shock wave/tubulent boundary-layer interaction experiments having sufficient quality to guide turbulence modeling and code validation efforts. Although the focus of this work is hypersonic, experiments at Mach numbers as low as 3 were considered. The principal means of identifying candidate studies was a computerized search of the AIAA Aerospace Database. Several hundred candidate studies were examined and over 100 of these were subjected to a rigorous set of acceptance criteria for inclusion in the data-base. Nineteen experiments were found to meet these criteria, of which only seven were in the hypersonic regime (M is greater than 5).
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 32; 7; p. 1377-1383
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  • 9
    Publication Date: 2011-08-24
    Description: The objective of the present investigation is to assess the effect of the spatial order of accuracy used for the evaluation of the inviscid fluxes on the resolution of higher order quantitites, such as velocity gradients. The viscous terms are computed as second-order accurate with central difference formulas, even though for the explicit part of the algorithm higher order approximations may be used. A viscous/inviscid method is used, and the outer part of the flowfield is computed with the inviscid flow equations. The viscous boundary-layer type flow region close to the body surface is computed with an algebraic eddy viscosity model. Results obtained with the conservative and nonconservative formulations and the viscous/inviscid approach are compared with available experimental data. The effect of grid refinement on the accuracy of the solution is also presented.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 32; 12; p. 2471-2474
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  • 10
    Publication Date: 2010-11-08
    Description: Stability derivatives have been computed for twisted wings of different planforms that include variations in both the wing taper and the aspect ratio. Taper ratios of 1.0, 0.50, and 0.25 are considered for each of three aspect ratios: 6, 10, and 16. The specific derivatives for which results are given are the rolling moment and the yawing moment derivatives with respect to rolling velocity, yawing velocity, and angle of sideslip. In addition to the stability derivatives, results are included for determining the theoretical rolling moment due to aileron deflection and a series of influence lines is given by which the loading across the span may be determined for any angle-of-attack distribution that may occur on the wing planforms considered.
    Keywords: AERODYNAMICS
    Type: REPT-635 , Collected Works of Robert T. Jones; p 147-165
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  • 11
    Publication Date: 2010-11-08
    Description: The two control operation of a conventional airplane is treated by means of the theory of disturbed motions. The consequences of this method of control are studied with regard to the stability of the airplane in its unconstrained components of motion and the movements set up during turn maneuvers. It is found that the motion of a conventional airplane is more stable when an arbitrary kinematic constraint is imposed in banking than when such constraint is imposed in yawing. Several hypothetical assumptions of piloting procedure, each of which is considered to represent a component of the actual procedure, are studied. Different means of two control operation are also discussed and it is concluded that a reliable rolling moment control that does not give the usual adverse secondary yawing moment should be most satisfactory.
    Keywords: AERODYNAMICS
    Type: REPT-579 , Collected Works of Robert T. Jones; p 73-90
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  • 12
    Publication Date: 2010-11-08
    Description: In theory, antisymmetric arrangements of wings and bodies can have smaller wave drag than corresponding mirror-symmetric arrangements. Thus, a long narrow oblique wing which presents the same aspect for two opposite directions of flight is potentially more efficient than corresponding (i.e., structurally equivalent) swept wing. The single continuous wing panel also adapts itself more readily to varying angles of obliquity, and hence, to varying flight speeds. Previous work on the aerodynamics and flight stability of oblique wing combinations is reviewed and a possible mode of application to transport aircraft operating at moderate supersonic speeds is suggested.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 657-664
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  • 13
    Publication Date: 2010-11-08
    Description: The items discussed are: (1) a recently proposed correction formula for the effect of compressibility in two dimensional subsonic flow; (2) the equivalence rule and the area rule for transonic speeds; (3) reciprocal relations in linearized wing theory; and (4) some general results connected with the problem of minimum wave resistance. The paper concludes with an example showing indentation of the fuselage to obtain favorable interference with the wing at supersonic speeds.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 601-608
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  • 14
    Publication Date: 2010-11-08
    Description: The assumptions of the thin airfoil theory are found to provide certain necessary conditions for the minimum drag of airfoils having a given total lift, a given maximum thickness, or a given volume. The conditions are applicable to steady or unsteady motions and to subsonic or supersonic speeds without restriction on the planform. The computation of drag and the statement of the conditions for minimum drag depend on the consideration of a combined flow field, which is obtained by superimposing the disturbance velocities in forward and reversed motions. If the planform of the airfoil and its total lift are given, it is found that, for minimum drag, the lift must be distributed in such a way that the downwash in the combined field is constant over the entire planform. If the planform is given and the thickness of the airfoil is required to contain a specified volume, then the thickness must be distributed over the planform in such a way that the pressure gradient of the combined field in the direction of flight is constant at all points of the wing.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 557-565
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  • 15
    Publication Date: 2010-11-08
    Description: A method is suggested for predicting the stability of automatically controlled aircraft by a comparison of calculated frequency-response curves for the aircraft and experimentally determined frequency-response curves for the automatic pilot. The method is applied only to stabilization in roll. The method is expected to be useful as a means of establishing the specifications of the performance required of the automatic control device for pilotless aircraft designed as missiles.
    Keywords: AERODYNAMICS
    Type: TN-1901 , Collected Works of Robert T. Jones; p 515-532
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  • 16
    facet.materialart.
    In:  CASI
    Publication Date: 2010-11-08
    Description: The analysis indicates that for aerodynamic efficiency, wings designed for flight at supersonic speeds should be swept back at an angle greater than the Mach angle and the angle of sweepback should be such that the component of velocity normal to the leading edge is less than the critical speed of the airfoil sections. This principle may also be applied to wings designed for subsonic speeds near the speed of sound, for which the induced velocities resulting from the thickness might otherwise be sufficiently great to cause shock waves.
    Keywords: AERODYNAMICS
    Type: REPT-863 , Collected Works of Robert T. Jones; p 377-383
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  • 17
    Publication Date: 2010-11-08
    Description: A method is presented for predicting the amplitude and frequency, under certain simplifying conditions, of the hunting oscillations of an automatically controlled aircraft with lag in the control system or in the response of the aircraft to the controls. If the steering device is actuated by a simple right-left type of signal, the series of alternating fixed amplitude signals occuring during the hunting may ordinarily be represented by a square wave. Formulas are given expressing the response to such a variation of signal in terms of the response to a unit signal.
    Keywords: AERODYNAMICS
    Type: REPT-801 , Collected Works of Robert T. Jones; p 363-368
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  • 18
    Publication Date: 2010-11-08
    Description: In the wing section theory the magnitude of the circulation, and hence of the lift, is determined by the velocity that would be induced near the trailing edge of the section in a non-lifting potential flow. In three dimensional flow the problem is complicated by the presence of the wake and no simple basic solution has been found. Treatment of the problem of a wing of finite span is reported on the basis of the two dimensional theory, corrected for the effect of the wake.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 245-249
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  • 19
    Publication Date: 2010-11-08
    Description: The method of operators is used in the application of nonuniform lift theory to problems of airplane dynamics. The method is adapted to the determination of the lift under prescribed conditions of motion or to the determination of the motions with prescribed disturbing forces.
    Keywords: AERODYNAMICS
    Type: TN-667 , Collected Works of Robert T. Jones; p 179-191
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  • 20
    Publication Date: 2010-11-08
    Description: The conditions essential to the stability of an airplane with free control surfaces are studied. Calculations are based on typical airplane characteristics with certain factors varied to cover a range of current designs. The effects of reducing the chord and of eliminating the floating tendency of the surface, of changing the wing loading and of decreasing the radius of gyration of the airplane are indicated. An investigation has also been made of the nature of the motion of the airplane with controls free and of the modes of instability that may occur. Stability with the controls free generally depends more critically on the design of the control system than on the stability characteristics of the airplane.
    Keywords: AERODYNAMICS
    Type: REPT-709 , Collected Works of Robert T. Jones; p 221-234
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  • 21
    Publication Date: 2010-11-08
    Description: An analysis of the principal results of recent lateral control research is made. Two things are considered of primary importance in judging the effectiveness of different control devices: The (calculated) banking and yawing motion of a typical small airplane caused by a deflection of the control, and the stick force required to produce this deflection. The report includes a table in which a number of different lateral control devices are compared on these bases. Test flights demonstrated that satisfactory lateral control at high angles of attack depends as much on the retention of stability as on aileron effectiveness.
    Keywords: AERODYNAMICS
    Type: REPT-605 , Collected Works of Robert T. Jones; p 117-145
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  • 22
    Publication Date: 2010-11-08
    Description: It is shown that the control force of ordinary ailerons may be reduced to zero over a range of deflections and at a given flight condition by the use of an appropriate differential movement. Approximations to the ideal motion obtainable with a simple linkage are discussed and a chart that enables the selection of an appropriate crank arrangement is presented. Various aspects of the practical application of the system are discussed and it is concluded that a small fixed tab, deflected to trim both ailerons upward, would be advantageous.
    Keywords: AERODYNAMICS
    Type: TN-586 , Collected Works of Robert T. Jones; p 91-115
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  • 23
    Publication Date: 2011-09-13
    Description: Discontinuous flows and free streamline solutions for axisymmetric bodies at zero and small angles of attack
    Keywords: AERODYNAMICS
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  • 24
    Publication Date: 2010-11-08
    Description: In linearized flow theory, certain very interesting extremal properties of wings can be derived under rather broad conditions without the use of a complicated mathematical apparatus. The present chapter reviews certain results of this theory and indicates some rather obvious extensions to incorporate various auxiliary conditions. Several examples illustrating the relation between the geometrical features of the wing and the lift distribution for minimum drag are given.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Res. Center Collected Works of Robert T. Jones; p 645-656
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  • 25
    facet.materialart.
    In:  CASI
    Publication Date: 2010-11-08
    Description: Recent theoretical and experimental work in supersonic aerodynamics is reviewed with its practical application in mind. Several arrangements of supporting surfaces and bodies are discussed and in some cases comparisons of theory and experiment are made. Finally, certain phenomena connected with lift and drag in a rarefied medium are considered briefly.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 625-644
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  • 26
    Publication Date: 2010-11-08
    Description: Comparisons of wing-body combinations may not disclose the full effect of a loss in aerodynamic efficiency. If the thrust needs to be increased at a given altitude then more or larger engines will have to be used and the possibility of concealing them becomes less. In this process the lift drag ratio of the complete airplane may become still more unfavorable than indicated by the comparison. Primarily aerodynamic and structural considerations point toward the development of turbojet engines specifically adapted to operation in an atmosphere of one tenth normal density. In addition to the numerous other technological problems associated with operation at these high altitudes, the problems of safe descent and effective limitation to low speeds at low altitudes seem important.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 579-592
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  • 27
    Publication Date: 2010-11-08
    Description: Some of the recent advances in the theory of thin airfoils are presented with particular reference to extensions of the theory to three dimensional flows and to supersonic speeds. The problem discussed herein is the calculation of the small disturbance velocities u, v, and w in the external field produced by the flight velocity V of the airfoil.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 483-497
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  • 28
    Publication Date: 2010-11-08
    Description: Recent developments in supersonic flow theory are applied to obtain estimates of the lift-drag ratios that may be achieved by aircraft employing sweptback wings. Lift-drag ratios greater than 10 to 1 can be maintained up to a Mach number of 1.4by the use of large angles of sweep and high aspect ratios. As the speed increases in the supersonic range the attainable lift-drag ratios decrease and the gain due to sweepback also appears to diminish. An efficient configuration for M = 1.4 would require about 60 deg sweepback, an aspect ratio of 4 and a wing loading of one third the atmospheric pressure. For a wing loading of 50 pounds per square foot the cruising altitude would be 60,000 feet and the indicated airspeed 290 miles per hour.
    Keywords: AERODYNAMICS
    Type: TN-1350 , Collected Works of Robert T. Jones; p 425-471
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  • 29
    Publication Date: 2010-11-08
    Description: Low aspect ratio wings having pointed planforms are treated on the assumption that the flow potentials in planes at right angles of the long axis of the airfoils are similar to the corresponding two dimensional potentials. For the limiting case of small angles of attack and low aspect ratios the theory brings out the following significant properties: (1) the lift of a slender pointed airfoil moving in the direction of its long axis depends on the increase in width of the sections in a downstream direction; (2) spanwise loading of such an airfoil is independent of planform and approaches the distribution giving a minimum induced drag; and (3) lift distribution of a pointed airfoil traveling point-foremost is relatively unaffected by the compressibility of the air below or above the speed of sound.
    Keywords: AERODYNAMICS
    Type: REPT-835 , Collected Works of Robert T. Jones; p 369-375
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  • 30
    Publication Date: 2010-11-08
    Description: An analysis was made to show the relative effectiveness of streamline external fuel tanks, a fuel tank in the form of a wing mounted in a biplane position, and auxiliary wing panels attached at the wing tips to increase the span as temporary means for increasing the range of a fighter-type airplane. Figures and charts for the various devices considered show the results of calculations of range, duration of flight, and take-off distance for both land base and carrier operation. The results indicated that the wing tip extensions were the most promising of the devices considered.
    Keywords: AERODYNAMICS
    Type: L-223 , Collected Works of Robert T. Jones; p 315-334
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  • 31
    Publication Date: 2010-11-08
    Description: By a generalization of the Joukowski method, a procedure is developed for effecting localized modifications of airfoil shapes and for determining graphically the resultant changes in the pressure distribution. The application of the procedure to the determination of the pressure distribution over airfoils of original design is demonstrated. Formulas for the lift, the moment, and the aerodynamic center are also given.
    Keywords: AERODYNAMICS
    Type: REPT-722 , Collected Works of Robert T. Jones; p 235-243
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  • 32
    Publication Date: 2010-11-08
    Description: It is shown that the drag of any semi-infinite airfoil section in purely subsonic inviscid flow follows precisely the Prandtl-Glauert compressibility rule. The result for the parabola has application to leading edge corrections in thin airfoil theory.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 619-623
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  • 33
    Publication Date: 2010-11-08
    Description: A method is reported for determining mathematically the combined disturbance field, and in certain cases the minimum drag, of wings at supersonic speeds. The simplest analytic example is provided by the wing of elliptic planform, which achieves its minimum drag when the lift is distributed uniformly over the surface. With a symmetrical distribution of thickness, the requirement of minimum drag for a given total volume is found to lead to profiles of constant curvature.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 567-578
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  • 34
    facet.materialart.
    In:  CASI
    Publication Date: 2010-11-08
    Description: Prandtl's theory is used to determine the airflow over bodies and wings adapted to supersonic flight. By making use of these results, and by incorporating in them an allowance for the probable skin friction, some estimates of expected lift-drag ratios are made for various flight speeds with the best configuration. At each speed a slender body and wings having the best angle of sweepback are considered. For the range of supersonic speeds shown an airplane of normal density and loading would be required to operate at an altitude of the order of 60,000 feet. The limiting value of 1-1/2 times the speed of sound corresponds to a flight speed of 1000 miles per hour. At this speed about 1.5 miles per gallon of fuel are expected. It is interesting to note that this value corresponds to a value of more than 15 miles per gallon when the weight is reduced to correspond to that of an ordinary automobile.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 499-514
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  • 35
    Publication Date: 2010-11-08
    Description: Methods of thin airfoil theory have been extended to oblique or sweptback airfoils of finite aspect ratio moving at supersonic speeds. The cases considered thus far are symmetrical airfoils at zero lift having planforms bounded by straight lines. Because of the conical form of the elementary flow fields, the results are comparable in simplicity to the results of the two dimensional thin airfoil theory for subsonic speeds. In the case of untapered airfoils swept back behind the Mach cone the pressure distribution at the center section is similar to that given by the Ackeret theory for a straight airfoil. With increasing distance from the center section the distribution approaches the form given by the subsonic flow theory. The pressure drag is concentrated chiefly at the center section and or long wings a slight negative drag may appear on outboard sections.
    Keywords: AERODYNAMICS
    Type: REPT-851 , Collected Works of Robert T. Jones; p 403-413
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  • 36
    Publication Date: 2010-11-08
    Description: A theoretical analysis is made to determine the optimum chord distribution, location, and extent of control surfaces, with the ratio of hinge moment to effectiveness as the criterion. Expressions for the effectiveness - for ailerons, the rolling moment, and for tail surfaces, the change of lift on the tail due to deflection of the surface-were derived from lifting line theory. Solutions found for a range of airfoil planforms indicate that, regardless of the characteristics of the tail surface, the chord of the rudder or the elevator should be very nearly constant over its span. The optimum ailerons are also of a characteristics shape, varying little with the planform of the wing.
    Keywords: AERODYNAMICS
    Type: REPT-731 , Collected Works of Robert T. Jones; p 307-314
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  • 37
    Publication Date: 2010-11-08
    Description: Problems involved in the stability and control of tailless airplanes are discussed. Such factors as the location of the aerodynamic center and its effect on the longitudinal stability, longitudinal trim with high lift devices, the effects of various changes in the shape of the wing on lateral stability, and the effects of nacelles are covered. It appears that sufficient stability and controllability can be secured without sweepback. With sweepback, a flap over the center section of the wing may be used to serve the dual purpose of elevator control and high lift device. Sweepback introduces undesirable stalling characteristics, however, and may require auxiliary devices to prevent stalling of the tips.
    Keywords: AERODYNAMICS
    Type: TN-837 , Collected Works of Robert T. Jones; p 251-280
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  • 38
    Publication Date: 2010-11-08
    Description: The effect of wing wake on the lift of the horizontal tail surfaces was studied. In the development of expressions for this effect, the growth of wing circulation and wing wake, the time interval represented by the tail length, and the development of lift by the tail were considered. The theory has been applied to a specific case to show the magnitude of the effect to be expected. It is shown that, for motions below a certain frequency, the development of lift by the tail may be represented by a simple lag function. The lag is, however, somewhat greater than that indicated by the tail length.
    Keywords: AERODYNAMICS
    Type: TN-771 , Collected Works of Robert T. Jones; p 203-220
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  • 39
    Publication Date: 2010-11-08
    Description: The application of mathematical advances made in electricity and other branches to problems of airplane dynamics is demonstrated. The Heaviside-Bromwich methods of solution of linear differential equations are described and it is shown how these methods avoid the consideration of boundary conditions and of particular or complementary integrals. It is pointed out that if the solution of the differential equation is obtained for the case of a unit disturbance, the effect of varying disturbances may be found therefrom by Carson's theorem. A graphical solution of Carson's integral for irregular disturbances is given. The procedure of obtaining unit solutions of the equations is then taken up and the analogy between Heaviside's symbolic series solution and a physical procedure of approximation is shown. It is suggested that a fictitious impulsive disturbance be used in the treatment of initial motions.
    Keywords: AERODYNAMICS
    Type: Collected Works of Robert T. Jones; p 21-29
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  • 40
    facet.materialart.
    In:  CASI
    Publication Date: 2011-10-14
    Description: The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration have been measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60 deg cropped delta wing platform, blended fuselage and two circular in-line jets that exit perpendicularly from the flat lower surface. The measured flows have a maximum freestream Mach number of 0.2. Model angle of attack is varied between -10 and +20 deg. The flow is ambient temperature in both jet exits and the nozzle pressure ratios are varied between 1 and 3. The data presented includes forces and moments, pressures measured at 281 surface pressure ports and the pressures of the jets. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exits to guide boundary condition selection for the planned computations. Flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.
    Keywords: AERODYNAMICS
    Type: AGARD, A Selection of Experimental Test Cases for the Validation of CFD Codes, Volume 2; 16 p
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  • 41
    Publication Date: 2011-10-14
    Description: This test was initiated to provide validation data on low aspect ratio wings at transonic speeds. The test was conducted so that the data obtained would be useful in the validation of codes, and all boundary condition data required would be measured as part of the test. During the conduct of the test, the measured quantities were checked for repeatability, and when the data would not repeat, the cause was tracked down and either eliminated or included in the measurement uncertainty. The accuracy of the data was in the end limited by wall imperfections of the wind tunnel in which the test was run.
    Keywords: AERODYNAMICS
    Type: AGARD, A Selection of Experimental Test Cases for the Validation of CFD Codes, Volume 2; 11 p
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  • 42
    Publication Date: 2011-10-14
    Description: A new technique is reported for calculating the entire flow field on spherically blunted cones at high angles of attack and high laminar Reynolds numbers. An approximate system of parabolic equations obtained from the steady Navier-Stokes equations by assuming the viscous, streamwise derivative terms are small compared to the viscous normal and circumferential derivatives is the basis of the calculations. These equations are valid for both the inviscid and viscous regions, including the circumferential separation zone that develops on the leeward side at high angles of attack. Two different methods are used to obtain the initial conditions for these equations at the sphere cone tangency plane. For small nose Reynolds numbers, an axisymmetric merged layer solution around a sphere is rotated to provide a three-dimensional initial plane of data. For large nose Reynolds numbers, the nose region is solved using an inviscid, three dimensional time dependent solution combined with a boundary layer solution for the viscous flow. The computed flowfield including the leeward separation region is described and compared with data for a 7 deg half angle cone at 10 deg angle of attack, and a blunt 15 deg half angle cone at 15 deg angle of attack.
    Keywords: AERODYNAMICS
    Type: AGARD Flow Separation; 11 p
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  • 43
    Publication Date: 2011-10-14
    Description: In-flight studies of the overall and local components of drag of many types of aircraft were conducted. The primary goal of these studies was to evaluate wind-tunnel and semiempirical prediction methods. Some evaluations are presented in this paper which may be summarized by the following observations: Wind-tunnel predictions of overall vehicle drag can be accurately extrapolated to flight Reynolds numbers, provided that the base drag is removed and the boattail areas on the vehicle are small. The addition of ablated roughness to lifting body configurations causes larger losses in performance and stability than would be expected from the added friction drag due to the roughness. Successful measurements of skin friction have been made in flight to Mach numbers above 4. A reliable inflatable deceleration device was demonstrated in flight which effectively stabilizes and decelerates a lifting aircraft at supersonic speeds.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. Drag; 12 p
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  • 44
    Publication Date: 2013-08-31
    Description: An existing cold-jet facility at NASA Lewis Research Center was modified to produce swirling flows with controllable initial tangential velocity distribution. Two extreme swirl profiles, i.e., one with solid-body rotation and the other predominated by a free-vortex distribution, were produced at identical swirl number of 0.48. Mean centerline velocity decay characteristics of the solid-body rotation jet flow exhibited classical decay features of a swirling jet with S - 0.48 reported in the literature. However, the predominantly free-vortex distribution case was on the verge of vortex breakdown, a phenomenon associated with the rotating flows of significantly higher swirl numbers, i.e., S sub crit greater than or equal to 0.06. This remarkable result leads to the conclusion that the integrated swirl effect, reflected in the swirl number, is inadequate in describing the mean swirling jet behavior in the near field. The relative size (i.e., diameter) of the vortex core emerging from the nozzle and the corresponding tangential velocity distribution are also controlling factors. Excitability of swirling jets is also investigated by exciting a flow with a swirl number of 0.35 by plane acoustic waves at a constant sound pressure level and at various frequencies. It is observed that the cold swirling jet is excitable by plane waves, and that the instability waves grow about 50 percent less in peak r.m.s. amplitude and saturate further upstream compared to corresponding waves in a jet without swirl having the same axial mass flux. The preferred Strouhal number based on the mass-averaged axial velocity and nozzle exit diameter for both swirling and nonswirling flows is 0.4.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:180895 , NASA-CR-180895
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  • 45
    Publication Date: 2013-08-31
    Description: An experimental study has been conducted to investigate the far-field, self-similar properties of a flat plate wake. A plane turbulent wake was generated at the trailing edge of a smooth splitter plate separating two legs of a Mixing Layer Wind Tunnel, with both initial boundary layers tripped. For the present study, both legs were operated at a free-steam velocity in the test section of 15 m/s, giving a Reynolds number based on wake momentum thickness of about 1750. Single profile measurements were obtained at five streamwise locations using a Pitot probe for the mean velocity measurements and a single cross-wire probe for the turbulence data, which included statistics up to third order. The mean flow data indicated a self-similar behavior beyond a streamwise distance equivalent to about 350 wake momentum thicknesses. However, the turbulence data show better collapse beyond a distance equivalent to about 500 momentum thicknesses, with all the measured peak Reynolds stresses achieving constant, asymptotic levels. The asymptotic mean flow behavior and peak primary stress levels agree well with theoretical predictions based on a constant eddy viscosity model. The present data also agree reasonably well with previous measurements, of which only one set extends into the self-similar region. Detailed comparisons with previous data are presented and discussed in this report.
    Keywords: AERODYNAMICS
    Type: NASA-CR-185917 , JIAA-TR-95 , NAS 1.26:185917
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  • 46
    Publication Date: 2013-08-31
    Description: For fully supersonic flows, an efficient strategy for obtaining numerical solutions is to employ space marching techniques. A full potential marching technique, known as the SIMP code and capable of handling such embedded subsonic regions, has achieved some success analyzing low supersonic Mach number flows. The extension of the full potential approach to the Euler equation which model the exact nonlinear inviscid gas dynamic flow processes is presented. Within the assumption of an inviscid flow, such an Euler marching solver can be applied to a wide class of shocked flows including the hypersonic range. The intent is to maintain some of the basic features of the full potential SIMP code within the Euler solver in dealing with geometry input, gridding techniques, and input/output routines including post processing of results. An Euler marching code known as EMTAC was developed. Results obtained for a variety of configurations involving canard, wing, horizontal tail, flow-through inlet, and fuselage using both the EMTAC and SIMP codes are reported. For shocked cases satisfying the isentropic assumption, the EMTAC and SIMP codes produced practically the same results. In terms of execution time, the EMTAC code is slower.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4085 , NAS 1.26:4085
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  • 47
    Publication Date: 2013-08-31
    Description: Inlets for tractor installations of advanced turboprop propulsion systems were tested in three phases, covering a period from November, 1982 to January, 1984. Nacelle inlet configuration types included single scoop, twin scoop, and annular arrangements. Tests were performed with and without boundary layer diverters and several different diverter heights were tested for the single scoop inlet. This same inlet was also tested at two different axial positions. Test Mach numbers ranged from Mach 0.20 to 0.80. Types of data taken were: (1) internal and external pressures, including inlet throat recoveries; (2) balance forces, including thrust-minus-drag; and (3) propellar blade stresses.
    Keywords: AERODYNAMICS
    Type: NASA-CR-174937 , NAS 1.26:174937 , LG85ER0105
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  • 48
    Publication Date: 2013-08-31
    Description: There are many fluid flows where the onset of transition can be caused by different instability mechanisms which compete among themselves. The interaction is considered of two types of instability mode (at an asymptotically large Reynolds number) which can occur in the flow above a rotating disc. In particular, the interaction is examined between lower branch Tollmien-Schlichting (TS) waves and the upper branch, stationary, inviscid crossflow vortex whose asymptotic structure has been described by Hall (1986). This problem is studied in the context of investigating the effect of the vortex on the stability characteristics of a small TS wave. Essentially, it is found that the primary effect is felt through the modification to the mean flow induced by the presence of the vortex. Initially, the TS wave is taken to be linear in character and it is shown (for the cases of both a linear and a nonlinear stationary vortex) that the vortex can exhibit both stabilizing and destabilizing effects on the TS wave and the nature of this influence is wholly dependent upon the orientation of this latter instability. Further, the problem is examined with a larger TS wave, whose size is chosen so as to ensure that this mode is nonlinear in its own right. An amplitude equation for the evolution of the TS wave is derived which admits solutions corresponding to finite amplitude, stable, traveling waves.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181859 , ICASE-89-34 , NAS 1.26:181859
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  • 49
    Publication Date: 2013-08-31
    Description: Aerodynamic analysis using computational fluid dynamics (CFD) is most fruitful when it is combined with a thorough program of wind tunnel testing. The understanding of aerodynamic phenomena is enhanced by the synergistic use of both analysis methods. A technique is described for an integrated approach to determining the forces and moments acting on a wind tunnel model by using a combination of experimentally measured pressures and CFD predictions. The CFD code used was FLO57 (an Euler solver) and the wind tunnel model was a heavily instrumented delta wing with 62.5 deg of leading-edge sweep. A thorough comparison of the CFD results and the experimental data is presented for surface pressure distributions and longitudinal forces and moments. The experimental pressures were also integrated over the surface of the model and the resulting forces and moments are compared to the CFD and wind tunnel results. The accurate determination of various drag increments via the combined use of the CFD and experimental pressures is presented in detail.
    Keywords: AERODYNAMICS
    Type: A-89145 , NASA-TM-102195 , NAS 1.15:102195
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  • 50
    Publication Date: 2013-08-31
    Description: Engine airframe integration effects are investigated for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic through-flow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic through-flow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic through-flow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum lift-to-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from 1.10 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration.
    Keywords: AERODYNAMICS
    Type: NASA-CR-185140 , NAS 1.26:185140 , AIAA PAPER 89-2140 , E-5068 , Aircraft Design, Systems and Operations Conference; 31 Jul. - 2 Aug. 1989; Seattle, Wa; United States
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  • 51
    Publication Date: 2013-08-31
    Description: A cooperative program was undertaken by research organizations in England, France, Australia and the U.S. to study the capabilities of computational fluid dynamics codes (CFD) to predict the aerodynamic loading on helicopter rotor blades. The program goal is to compare predictions with experimental data for flight tests of a research Puma helicopter with rectangular and swept tip blades. Two topics are studied. First, computed results from three CFD codes are compared for flight test cases where all three codes use the same partial inflow-angle boundary conditions. Second, one of the CFD codes (FPR) is iteratively coupled with the CAMRAD/JA helicopter performance code. These results are compared with experimental data and with an uncoupled CAMRAD/JA solution. The influence of flow field unsteadiness is found to play an important role in the blade aerodynamics. Alternate boundary conditions are suggested in order to properly model this unsteadiness in the CFD codes.
    Keywords: AERODYNAMICS
    Type: NASA-TM-102226 , USAAVSCOM-TM-89-A-001 , NAS 1.15:102226 , A-89223
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  • 52
    Publication Date: 2013-08-31
    Description: An examination of the potential flow computer code VSAERO to model leading edge separation over a delta wing is examined. Recent improvements to the code suggest that it may be capable of predicting pressure coefficients on the body. Investigation showed that although that code does predict the vortex roll-up, the pressure coefficients have significant error. The program is currently unsatisfactory, but with some additional development it may become a useful tool for this application.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:185892 , NASA-CR-185892
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  • 53
    Publication Date: 2013-08-31
    Description: The experimental program for validating real gas hypersonic flow codes at NASA Ames Rsearch Center is described. Ground-based test facilities used include ballistic ranges, shock tubes and shock tunnels, arc jet facilities and heated-air hypersonic wind tunnels. Also included are large-scale computer systems for kinetic theory simulations and benchmark code solutions. Flight tests consist of the Aeroassist Flight Experiment, the Space Shuttle, Project Fire 2, and planetary probes such as Galileo, Pioneer Venus, and PAET.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100093 , A-88135 , NAS 1.15:100093 , AGARD Symposium on Validation of Computational Fluid Dynamics; 2-5 May 1988; Lisbon; Portugal
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  • 54
    Publication Date: 2013-08-31
    Description: This program is to develop an analytic method for reducing SUMS data for the determination of the undisturbed atmosphere conditions ahead of the shuttle along its descending trajectory. It is divided into an internal flow problem, an external flow problem and their matching conditions. Since the existing method of Direct Simulation Monte Carlo (DSMC) failed completely for the internal flow problem, the emphasis is on the internal flow of a highly non-equilibrium, rarefied air through a short tube of a diameter much less than the gaseous mean free path. A two fluid model analysis of this internal flow problem has been developed and studied with typical results illustrated. A computer program for such an analysis and a technical paper published in Lecture Notes in Physics No. 323 (1989) are included as Appendices 3 and 4. A proposal for in situ determination of the surface accommodation coefficients sigma sub t and sigma e is included in Appendix 5 because of their importance in quantitative data reduction. A two fluid formulation for the external flow problem is included as Appendix 6 and a review article for AIAA on Hypersonic propulsion, much dependent on ambient atmospheric density, is also included as Appendix 7.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:181826 , NASA-CR-181826
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  • 55
    Publication Date: 2013-08-31
    Description: An engineering approach was used to include the nonlinear effects of thickness and camber in an analytical aeroelastic analysis of cascades in supersonic acial flow (supersonic leading-edge locus). A hybrid code using Lighthill's nonlinear piston theory and Lanes's linear potential theory was developed to include these nonlinear effects. Lighthill's theory was used to calculate the unsteady pressures on the noninterference surface regions of the airfoils in cascade. Lane's theory was used to calculate the unsteady pressures on the remaining interference surface regions. Two airfoil profiles was investigated (a supersonic throughflow fan design and a NACA 66-206 airfoil with a sharp leading edge). Results show that compared with predictions of Lane's potential theory for flat plates, the inclusion of thickness (with or without camber) may increase or decrease the aeroelastic stability, depending on the airfoil geometry and operating conditions. When thickness effects are included in the aeroelastic analysis, inclusion of camber will influence the predicted stability in proportion to the magnitude of the added camber. The critical interblade phase angle, depending on the airfoil profile and operating conditions, may also be influenced by thickness and camber. Compared with predictions of Lane's linear potential theory, the inclusion of thickness and camber decreased the aerodynamic stifness and increased the aerodynamic damping at Mach 2 and 2.95 for a cascade of supersonic throughflow fan airfoils oscillating 180 degrees out of phase at a reduced frequency of 0.1.
    Keywords: AERODYNAMICS
    Type: NAS 1.15:101949 , E-4642 , NASA-TM-101949
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  • 56
    Publication Date: 2013-08-31
    Description: A very efficient direct particle simulation algorithm for hypersonic rarefied flows is presented and its implmentation on a Connection Machine is described. The implementation simulates ideal diatomic Maxwell molecules with three translational and two rotational degrees of freedom. Results for a 2-D simulation of supersonic flow over a 30 deg wedge are presented and used for validation.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:185428 , NASA-CR-185428 , RIACS-TR-88.46
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  • 57
    Publication Date: 2013-08-31
    Description: A 2.1 m diameter, 1/6-scale model helicopter main rotor was tested in hover in the test section of the NASA Ames 40- by 80-foot wind tunnel. Performance and noise data on a small-scale rotor at various thrust coefficients and tip Mach numbers were obtained for comparison with existing data on similar full-scale helicopter rotors. These data form part of a data base to permit the estimation of scaling effects on various rotor noise mechanisms. Another objective was to contribute to a data base that will permit the estimation of facility effects on acoustic testing. Acoustic 1/3-octave-band spectra are presented, together with variations of overall acoustic levels with rotor performance, microphone distance, and directivity angle.
    Keywords: AERODYNAMICS
    Type: A-89015 , NASA-TM-101058 , NAS 1.15:101058
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  • 58
    Publication Date: 2013-08-31
    Description: The steady-state and transient effects of simulated heavy rain on the subsonic aerodynamic characteristics of a wing model were determined in the Langley 14- by 22-Foot Subsonic Tunnel. The 1.29 foot chord wing was comprised of a NACA 23015 airfoil and had an aspect ratio of 6.10. Data were obtained while test variables of liquid water content, angle of attack, and trailing edge flap angle were parametrically varied at dynamic pressures of 10, 30, and 50 psf (i.e., Reynolds numbers of .76x10(6), 1.31x10(6), and 1.69x10(6)). The experimental results showed reductions in lift and increases in drag when in the simulated rain environment. Accompanying this was a reduction of the stall angle of attack by approximately 4 deg. The transient aerodynamic performance during transition from dry to wet steady-state conditions varied between a linear and a nonlinear transition.
    Keywords: AERODYNAMICS
    Type: L-16576 , NASA-TP-2932 , NAS 1.60:2932
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  • 59
    Publication Date: 2013-08-31
    Description: A comprehensive test program is defined which is being implemented in the NASA/Ames 3.5 foot Hypersonic Wind Tunnel for obtaining data on a generic all-body hypersonic vehicle for computational fluid dynamics (CFD) code validation. Computational methods (approximate inviscid methods and an upwind parabolized Navier-Stokes code) currently being applied to the all-body model are outlined. Experimental and computational results on surface pressure distributions and Pitot-pressure surveys for the basic sharp-nose model (without control surfaces) at a free-stream Mach number of 7 are presented.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:185347 , NASA-CR-185347 , National Aero-Space Plane Technology Symposium; 24-28 Apr. 1989; Sunnyvale, CA; United States
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  • 60
    Publication Date: 2013-08-31
    Description: A wind tunnel test was conducted in the NASA-Ames 7 x 10 ft wind tunnel to investigate the lift distribution on a semispan wing with a discontinuous change in spanwise twist. The semispan wing had a tip with an adjustable pitch angle independent on the inboard section pitch angle simulating the free tip rotor blade when its free tip is at a deflected position. The spanwise lift distribution over the wing and the tip were measured and three component velocity surveys behind the wing were obtained with a 3-D laser Doppler velocimeter (LV) with the wing at one angle of attack and the tip deflected at different pitch angles. A six-component internal strain gage balance was also used to measure total forces and moments on the tip. The 3-D lift was computed from the 2-D lift distributions obtained from the LV and from the strain gage balance. The results from both experimental methods are shown to be in agreement with predictions made by a steady, 3-D panel code, VSAERO.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:177532 , NASA-CR-177532
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  • 61
    Publication Date: 2013-08-31
    Description: Calculated results for the aerodynamic coefficients over the range of + or - 90 deg in both pitch and yaw attitude angles for the Aeroassist Flight Experiment (AFE) vehicle in free molecule flow are presented. The AFE body is described by a large number of small flat plate surface elements whose orientations are established in a wind axes coordinate system through the pitch and yaw attitude angles. Lift force, drag force, and three components of aerodynamic moment about a specified point are computed for each element. The elemental forces and moments are integrated over the entire body, and total force and moment coefficients are computed. The coefficients are calculated for the two limiting gas-surface molecular collision conditions, namely, specular and diffuse, which assume zero and full thermal accommodation of the incoming gas molecules with the surface, respectively. The individual contribution of the shear stress and pressure terms are calculated and also presented.
    Keywords: AERODYNAMICS
    Type: NAS 1.15:101600 , NASA-TM-101600
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  • 62
    Publication Date: 2013-08-31
    Description: Two dimensional problems are solved using numerical techniques. Navier-Stokes equations are studied both in the vorticity-stream function formulation which appears to be the optimal choice for two dimensional problems, using a storage approach, and in the velocity pressure formulation which minimizes the number of unknowns in three dimensional problems. Analysis shows that compact centered conservative second order schemes for the vorticity equation are the most robust for high Reynolds number flows. Serious difficulties remain in the choice of turbulent models, to keep reasonable CPU efficiency.
    Keywords: AERODYNAMICS
    Type: VKI, Unsteady Aerodynamics, Volume 2; 120 p
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  • 63
    Publication Date: 2013-08-31
    Description: A zonal method for modeling powered-lift aircraft flow fields is based on the coupling of a three-dimensional Navier-Stokes code to a potential flow code. By minimizing the extent of the viscous Navier-Stokes zones the zonal method can be a cost effective flow analysis tool. The successful coupling of the zonal solutions provides the viscous/inviscid interations that are necessary to achieve convergent and unique overall solutions. The feasibility of coupling the two vastly different codes is demonstrated. The interzone boundaries were overlapped to facilitate the passing of boundary condition information between the codes. Routines were developed to extract the normal velocity boundary conditions for the potential flow zone from the viscous zone solution. Similarly, the velocity vector direction along with the total conditions were obtained from the potential flow solution to provide boundary conditions for the Navier-Stokes solution. Studies were conducted to determine the influence of the overlap of the interzone boundaries and the convergence of the zonal solutions on the convergence of the overall solution. The zonal method was applied to a jet impingement problem to model the suckdown effect that results from the entrainment of the inviscid zone flow by the viscous zone jet. The resultant potential flow solution created a lower pressure on the base of the vehicle which produces the suckdown load. The feasibility of the zonal method was demonstrated. By enhancing the Navier-Stokes code for powered-lift flow fields and optimizing the convergence of the coupled analysis a practical flow analysis tool will result.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:177521 , NASA-CR-177521
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  • 64
    Publication Date: 2013-08-31
    Description: Conservative algorithms for boundary interfaces of overlaid grids are presented. The basic method is zeroth order, and is extended to a higher order method using interpolation and subcell decomposition. The present method, strictly based on a conservative constraint, is tested with overlaid grids for various applications of unsteady and steady supersonic inviscid flows with strong shock waves. The algorithm is also applied to a multi-level grid adaptation in which the next level finer grid is overlaid on the coarse base grid with an arbitrary orientation.
    Keywords: AERODYNAMICS
    Type: NAS 1.15:102080 , NASA-TM-102080 , E-4842 , AIAA PAPER 89-1980 , Computational Fluid Dynamics Conference; 13-15 Jun. 1989; Buffalo, NY; United States
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  • 65
    Publication Date: 2013-08-31
    Description: The three-dimensional finite element modeling techniques developed for the thermal analysis of a hypersonic wing test structure (HWTS) are described. The computed results are compared to measured test data. In addition, the results of a NASA two-dimensional parameter finite difference local thermal model and the results of a contractor two-dimensional lumped parameter finite difference local thermal model will be presented.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:185319 , NASA-CR-185319
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  • 66
    Publication Date: 2013-08-31
    Description: A hover test was conducted on a small scale rotor model for two sets of tapered rotor blades. The baseline rotor blade set used a NACA 0012 airfoil section, whereas the second rotor blade set had advanced rotorcraft airfoils distributed along the radius. The experiment was conducted for a range of thrust coefficients and tip speeds, and the data were compared to the predictions of three analytical methods. The data show the advantage of the advanced airfoils at the higher rotor thrust levels; two of the analyses predicted the correct data trends.
    Keywords: AERODYNAMICS
    Type: L-16407 , NASA-TP-2832 , AVSCOM-TP-88-B-001 , NAS 1.60:2832
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  • 67
    Publication Date: 2013-08-31
    Description: Static longitudinal and lateral-directional forces and moments were measured for elliptic- and crescent-wing models at high angles-of-attack in the NASA Langley 14 by 22-Ft Subsonic Tunnel. The forces and moments were obtained for an angle-of-attack range including stall and post-stall conditions at a Reynolds number based on the average wing chord of about 1.8 million. Flow-visualization photographs using a mixture of oil and titanium-dioxide were also taken for several incidence angles. The force and moment data and the flow-visualization results indicated that the crescent wing model with its highly swept tips produced much better high angle-of-attack aerodynamic characteristics than the elliptic model. Leading-edge separation-induced vortex flow over the highly swept tips of the crescent wing is thought to produce this improved behavior at high angles-of-attack. The unique planform design could result in safer and more efficient low-speed airplanes.
    Keywords: AERODYNAMICS
    Type: NASA-CR-184992 , NAS 1.26:184992
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  • 68
    Publication Date: 2013-08-31
    Description: Progress in a recently started project aimed at the prediction of transition to turbulence in hypersonic flow is briefly discussed. The prediction of transition to turbulence is a very important issue in the design of space vessels. Two space vehicles currently under investigation, namely the aeroassisted transfer vehicle (AOTV) and the trans-atmospheric vehicle (TAV), suffer from strong aerodynamic heating. This heating is strongly influenced by the boundary layer structure. These aerospace vehicles fly in the upper atmospheric layer at a Mach number between 10 and 30 at very low atmospheric pressures. At very high altitudes the flow is laminar, but when the space vessel returns to a lower orbit, the flow becomes turbulent and the heating is dramatically increased. The prediction of this transition process is commonly done by means of experiments. The experimental facilities available nowadays cannot model the hypersonic flow field accurately enough by limitations in Mach and Reynolds number. These facilities also have a large free stream disturbance level which makes it very difficult to investigate transition accurately. An alternative approach is to study transition by theoretical means. Up to now numerical studies of hypersonic flow only discussed steady laminar or turbulent flow. This theoretical approach is extended to the study of transition in hypersonic flow by means of direct numerical simulations and additional theoretical investigations to explain the mechanisms leading to transition. A brief outline of how this research is to be performed is given.
    Keywords: AERODYNAMICS
    Type: Annual Research Briefs, 1988; p 115-119
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  • 69
    Publication Date: 2013-08-31
    Description: A modeling technique for predicting the axial and transverse velocity characteristics of rectangular nozzle plumes is developed. In this technique, modeling of the plume cross section is initiated at the nozzle exit plane. The technique is demonstrated for the plume issuing from a rectangular nozzle having an aspect ratio of 6.0 and discharging into quiescent air. Application of the present procedures to a nozzle discharging into a moving airstream (flight effect) are then demonstrated. The effects of plume shear layer structure modification on the velocity flowfield are discussed and modeling procedures are illustrated by example.
    Keywords: AERODYNAMICS
    Type: E-4739 , NASA-TM-102047 , AIAA PAPER 89-2357 , NAS 1.15:102047 , Joint Propulsion Conference; 10-12 Jul. 1989; Monterey, CA; United States
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  • 70
    Publication Date: 2013-08-31
    Description: The development of the variational approach for the solution of inviscid aerodynamic problems using solution adaptive grids is discussed. The formulation of a new, directional weighting, functional has been shown to have desirable properties. The scheme has been applied to compute the transonic flow past two-dimensional airfoils using the Euler equations of inviscid, compressible flow. Transonic flows in quasi-one-dimensional nozzles and over the two dimensional airfoils are solved on the various solution-adaptive-grids to demonstrate the applicability of the proposed directional-concentration functional and the grid adaptation process from the stand point of improving the solution accuracy and demonstrating the overall convergence.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:184824 , NASA-CR-184824
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  • 71
    Publication Date: 2013-08-31
    Description: An iterative method for wall interference assessment and/or correction is presented for transonic flow conditions in wind tunnels equipped with two component velocity measurements on a single interface. The iterative method does not require modeling of the test article and tunnel wall boundary conditions. Analytical proof for the convergence and stability of the iterative method is shown in the subsonic flow regime. The numerical solutions are given for both 2-D and axisymmetrical cases at transonic speeds with the application of global Mach number correction.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 2; p 853-866
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  • 72
    Publication Date: 2013-08-31
    Description: Wind tunnel wall interference assessment and correction (WIAC) concepts, applications, and typical results are discussed in terms of several nonlinear transonic codes and one panel method code developed for and being implemented at NASA-Langley. Contrasts between 2-D and 3-D transonic testing factors which affect WIAC procedures are illustrated using airfoil data from the 0.3 m Transonic Cryogenic Tunnel and Pathfinder 1 data from the National Transonic Facility. Initial results from the 3-D WIAC codes are encouraging; research on and implementation of WIAC concepts continue.
    Keywords: AERODYNAMICS
    Type: Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 2; p 817-851
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  • 73
    Publication Date: 2013-08-31
    Description: Three dimensional linear secondary instability theory is extended for compressible boundary layers on a flat plate in the presence of finite amplitude Tollmien-Schlichting waves. The focus is on principal parametric resonance responsible for strong growth of subharmonics in low disturbance environment.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 2; p 691-704
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  • 74
    Publication Date: 2013-08-31
    Description: When three-dimensional separation occurs on a body immersed in a flow governed by the incompressible Navier-Stokes equations, the geometrical surfaces formed by the three vector fields (velocity, vorticity and the skin-friction) and a scalar field (pressure) become interrelated through topological maps containing their respective singular points and extremal points. A mathematically consistent description of these singular points becomes inevitable when we want to study the geometry of the separation. A separated stream surface requires, for example, the existence of a saddle-type singular point on the skin-friction surface. This singular point is actually, in the proper language of mathematics, a saddle of index two. The index is a measure of the dimension of the outset (set leaving the singular point). Hence, when a saddle of index two is specified, a two dimensional surface that becomes separated from the osculating plane of the saddle is implied. The three-dimensional singular point is interpreted mathematically and the most common aerodynamical singular points are discussed through this perspective.
    Keywords: AERODYNAMICS
    Type: USAAVSCOM-TR-87-A-14 , A-88029 , NASA-TM-100045 , AD-A197978 , NAS 1.15:100045
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  • 75
    Publication Date: 2013-08-31
    Description: A cell-vertex scheme is outlined for solving the flow about a delta wing with M (sub infinity) is greater than 1. Embedded regions of mesh refinement allow solutions to be obtained which have much higher resolution than those achieved to date. Effects of mesh refinement and artificial viscosity on the solutions are studied, to determine at what point leading-edge vortex solutions are grid-converged. A macroscale and a microscale for the size of the vortex are defined, and it is shown that the macroscale (which includes the wing surface properties) is converged on a moderately refined grid, while the microscale is very sensitive to grid spacing. The level of numerical diffusion in the core of the vortex is found to be substantial. Comparisons with the experiment are made for two cases which have transonic cross-flow velocities.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 231-259
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  • 76
    Publication Date: 2013-08-31
    Description: The application of computational fluid dynamics (CFD) to fighter aircraft design and development is discussed. Methodology requirements for the aerodynamic design of fighter aircraft are briefly reviewed. The state-of-the-art of computational methods for transonic flows in the light of these requirements is assessed and the techniques found most adequate for the subject application are identified. Highlights from some proof-of-feasibility Euler and Navier-Stokes computations about a complete fighter aircraft configuration are presented. Finally, critical issues and opportunities for design application of CFD are discussed.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 153-173
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  • 77
    Publication Date: 2013-08-31
    Description: A brief survey is given on the study of transonic shock/boundary layer effects in flight. Then the possibility of alleviating the adverse shock effects through passive shock control is discussed. A Swedish flight experiment on a swept wing attack aircraft is used to demonstrate how it is possible to reduce the extent of separated flow and increase the drag-rise Mach number significantly using a moderate amount of perforation of the surface.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 61-77
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  • 78
    Publication Date: 2013-08-31
    Description: Performance optimization for rotors in hover and axial flight is a topic of continuing importance to rotorcraft designers. The aim of this Phase 1 effort has been to demonstrate that a linear optimization algorithm could be coupled to an existing influence coefficient hover performance code. This code, dubbed EHPIC (Evaluation of Hover Performance using Influence Coefficients), uses a quasi-linear wake relaxation to solve for the rotor performance. The coupling was accomplished by expanding of the matrix of linearized influence coefficients in EHPIC to accommodate design variables and deriving new coefficients for linearized equations governing perturbations in power and thrust. These coefficients formed the input to a linear optimization analysis, which used the flow tangency conditions on the blade and in the wake to impose equality constraints on the expanded system of equations; user-specified inequality contraints were also employed to bound the changes in the design. It was found that this locally linearized analysis could be invoked to predict a design change that would produce a reduction in the power required by the rotor at constant thrust. Thus, an efficient search for improved versions of the baseline design can be carried out while retaining the accuracy inherent in a free wake/lifting surface performance analysis.
    Keywords: AERODYNAMICS
    Type: NAS 1.26:177524 , NASA-CR-177524
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  • 79
    Publication Date: 2013-08-31
    Description: The stability of compressible 2-D and 3-D boundary layers is reviewed. The stability of 2-D compressible flows differs from that of incompressible flows in two important features: There is more than one mode of instability contributing to the growth of disturbances in supersonic laminar boundary layers and the most unstable first mode wave is 3-D. Whereas viscosity has a destabilizing effect on incompressible flows, it is stabilizing for high supersonic Mach numbers. Whereas cooling stabilizes first mode waves, it destabilizes second mode waves. However, second order waves can be stabilized by suction and favorable pressure gradients. The influence of the nonparallelism on the spatial growth rate of disturbances is evaluated. The growth rate depends on the flow variable as well as the distance from the body. Floquet theory is used to investigate the subharmonic secondary instability.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 2; p 629-689
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  • 80
    Publication Date: 2013-08-31
    Description: Algorithms are described for the generation and adaptation of unstructured grids in two and three dimensions, as well as Euler solvers for unstructured grids. The main purpose is to demonstrate how unstructured grids may be employed advantageously for the economic simulation of both geometrically as well as physically complex flow fields.
    Keywords: AERODYNAMICS
    Type: Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 377-408
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  • 81
    Publication Date: 2013-08-31
    Description: For many internal transonic flows of practical interest, some of the relevant nondimensional parameters typically are small enough that a perturbation scheme can be expected to give a useful level of numerical accuracy. A variety of steady and unsteady transonic channel and cascade flows is studied with the help of systematic perturbation methods which take advantage of this fact. Asymptotic representations are constructed for small changes in channel cross-section area, small flow deflection angles, small differences between the flow velocity and the sound speed, small amplitudes of imposed oscillations, and small reduced frequencies. Inside a channel the flow is nearly one-dimensional except in thin regions immediately downstream of a shock wave, at the channel entrance and exit, and near the channel throat. A study of two-dimensional cascade flow is extended to include a description of three-dimensional compressor-rotor flow which leads to analytical results except in thin edge regions which require numerical solution. For unsteady flow the qualitative nature of the shock-wave motion in a channel depends strongly on the orders of magnitude of the frequency and amplitude of impressed wall oscillations or fluctuations in back pressure. One example of supersonic flow is considered, for a channel with length large compared to its width, including the effect of separation bubbles and the possibility of self-sustained oscillations. The effect of viscosity on a weak shock wave in a channel is discussed.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 261-291
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  • 82
    Publication Date: 2013-08-31
    Description: A computer analysis was developed for calculating steady (or unsteady) three-dimensional aircraft component flow fields. This algorithm, called ENS3D, can compute the flow field for the following configurations: diffuser duct/thrust nozzle, isolated wing, isolated fuselage, wing/fuselage with or without integrated inlet and exhaust, nacelle/inlet, nacelle (fuselage) afterbody/exhaust jet, complete transport engine installation, and multicomponent configurations using zonal grid generation technique. Solutions can be obtained for subsonic, transonic, or hypersonic freestream speeds. The algorithm can solve either the Euler equations for inviscid flow, the thin shear layer Navier-Stokes equations for viscous flow, or the full Navier-Stokes equations for viscous flow. The flow field solution is determined on a body-fitted computational grid. A fully-implicit alternating direction implicit method is employed for the solution of the finite difference equations. For viscous computations, either a two layer eddy-viscosity turbulence model or the k-epsilon two equation transport model can be used to achieve mathematical closure.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 175-194
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  • 83
    Publication Date: 2013-08-31
    Description: Numerous computational fluid dynamics (CFD) codes are available that solve any of several variations of the transonic flow equations from small disturbance to full Navier-Stokes. The design philosophy at General Dynamics Fort Worth Division involves use of all these levels of codes, depending on the stage of configuration development. Throughout this process, drag calculation is a central issue. An overview is provided for several transonic codes and representative test-to-theory comparisons for fighter-type configurations are presented. Correlations are shown for lift, drag, pitching moment, and pressure distributions. The future of applied CFD is also discussed, including the important task of code validation. With the progress being made in code development and the continued evolution in computer hardware, the routine application of these codes for increasingly more complex geometries and flow conditions seems apparent.
    Keywords: AERODYNAMICS
    Type: NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 1; p 109-132
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  • 84
    Publication Date: 2013-08-31
    Description: Experimental data for a series of two- and three-dimensional shock wave/turbulent boundary layer interaction flows at Mach 7 are presented. Test bodies, composed of simple geometric shapes, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface-pressure and heat-transfer distributions as well as limited mean-flow-field surveys in both the undisturbed and the interaction regimes. The data are presented in a convenient form for use in validating existing or future computational models of these generic hypersonic flows.
    Keywords: AERODYNAMICS
    Type: NAS 1.15:101075 , NASA-TM-101075 , A-89048
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  • 85
    Publication Date: 2013-08-31
    Description: In the past decade, there has been much activity in the development of computational methods for the analysis of unsteady transonic aerodynamics about airfoils and wings. Significant features are illustrated which must be addressed in the treatment of computational transonic unsteady aerodynamics. The flow regimes for an aircraft on a plot of lift coefficient vs. Mach number are indicated. The sequence of events occurring in air combat maneuvers are illustrated. And further features of transonic flutter are illustrated. Also illustrated are several types of aeroelastic response which were encountered and which offer challenges for computational methods. The four cases illustrate problem areas encountered near the boundaries of aircraft envelopes, as operating condition change from high speed, low angle conditions to lower speed, higher angle conditions.
    Keywords: AERODYNAMICS
    Type: Transonic Unsteady Aerodynamics and Aeroelasticity 1987, Part 2; p 631-637
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  • 86
    Publication Date: 2013-08-31
    Description: Present flutter analysis methods do not accurately predict the flutter speeds in the transonic flow region for wings with supercritical airfoils. Aerodynamic programs using computational fluid dynamic (CFD) methods are being developed, but these programs need to be verified before they can be used with confidence. A wind tunnel test was performed to obtain all types of data necessary for correlating with CFD programs to validate them for use on high aspect ratio wings. The data include steady state and unsteady aerodynamic measurements on a nominal stiffness wing and a wing four times that stiffness. There is data during forced oscillations and during flutter at several angles of attack, Mach numbers, and tunnel densities.
    Keywords: AERODYNAMICS
    Type: Transonic Unsteady Aerodynamics and Aeroelasticity 1987, Part 2; p 543-570
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  • 87