### Investigations of Compression Shocks and Boundary Layers in Gases Moving at High Speed

**Date:**January 1, 1947

**Creator:**Ackeret, J.; Feldmann, F. & Rott, N.

**Description:**The mutual influences of compression shocks and friction boundary layers were investigated by means of high speed wind tunnels.Schlieren optics provided a clear picture of the flow phenomena and were used for determining the location of the compression shocks, measurement of shock angles, and also for Mach angles. Pressure measurement and humidity measurements were also taken into consideration.Results along with a mathematical model are described.

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### The Calculation of Compressible Flows with Local Regions of Supersonic Velocity

**Date:**March 1, 1947

**Creator:**Goethert, B. & Kawalki, K. H.

**Description:**This report addresses a method for the approximate calculation of compressible flows about profiles with local regions of supersonic velocity. The flow around a slender profile is treated as an example.

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### Drag Reduction by Suction of the Boundary Layer Separated Behind Shock Wave Formation at High Mach Numbers

**Date:**July 1, 1947

**Creator:**Regenscheit, B.

**Description:**With an approach of the velocity of flight of a ship to the velocity of sound, there occurs a considerable increase of the drag. The reason for this must be found in the boundary layer separation caused by formation of shock waves. It will be endeavored to reduce the drag increase by suction of the boundary layer. Experimental results showed that drag increase may be considerably reduced by this method. It was, also, observed that, by suction, the position of shock waves can be altered to a considerable extent.

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### On Possible Similarity Solutions for Three-Dimensional Incompressible Laminar Boundary-Layer Flows Over Developable Surfaces and with Proportional Mainstream Velocity Components

**Date:**September 1, 1958

**Creator:**Hansen, Arthur G.

**Description:**Analysis is presented on the possible similarity solutions of the three-dimensional, laminar, incompressible, boundary-layer equations referred to orthogonal, curvilinear coordinate systems. Requirements of the existence of similarity solutions are obtained for the following: flow over developable surface and flow over non-developable surfaces with proportional mainstream velocity components.

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### On the Contribution of Turbulent Boundary Layers to the Noise Inside a Fuselage

**Date:**December 1, 1956

**Creator:**Corcos, G. M. & Liepmann, H. W.

**Description:**The following report deals in preliminary fashion with the transmission through a fuselage of random noise generated on the fuselage skin by a turbulent boundary layer. The concept of attenuation is abandoned and instead the problem is formulated as a sequence of two linear couplings: the turbulent boundary layer fluctuations excite the fuselage skin in lateral vibrations and the skin vibrations induce sound inside the fuselage. The techniques used are those required to determine the response of linear systems to random forcing functions of several variables. A certain degree of idealization has been resorted to. Thus the boundary layer is assumed locally homogeneous, the fuselage skin is assumed flat, unlined and free from axial loads and the 'cabin' air is bounded only by the vibrating plate so that only outgoing waves are considered. Some of the details of the statistical description have been simplified in order to reveal the basic features of the problem. The results, strictly applicable only to the limiting case of thin boundary layers, show that the sound pressure intensity is proportional to the square of the free stream density, the square of cabin air density and inversely proportional to the first power of the damping constant ...

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### Flow and Force Equations for a Body Revolving in a Fluid

**Date:**December 1, 1979

**Creator:**Zahm, A. F.

**Description:**A general method for finding the steady flow velocity relative to a body in plane curvilinear motion, whence the pressure is found by Bernoulli's energy principle is described. Integration of the pressure supplies basic formulas for the zonal forces and moments on the revolving body. The application of the steady flow method for calculating the velocity and pressure at all points of the flow inside and outside an ellipsoid and some of its limiting forms is presented and graphs those quantities for the latter forms. In some useful cases experimental pressures are plotted for comparison with theoretical. The pressure, and thence the zonal force and moment, on hulls in plane curvilinear flight are calculated. General equations for the resultant fluid forces and moments on trisymmetrical bodies moving through a perfect fluid are derived. Formulas for potential coefficients and inertia coefficients for an ellipsoid and its limiting forms are presented.

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### The Inertia Coefficients of an Airship in a Frictionless Fluid

**Date:**December 1, 1979

**Creator:**Bateman, H.

**Description:**The apparent inertia of an airship hull is examined. The exact solution of the aerodynamical problem is studied for hulls of various shapes with special attention given to the case of an ellipsoidal hull. So that the results for the ellipsoidal hull may be readily adapted to other cases, they are expressed in terms of the area and perimeter of the largest cross section perpendicular to the direction of motion by means of a formula involving a coefficient kappa which varies only slowly when the shape of the hull is changed, being 0.637 for a circular or elliptic disk, 0.5 for a sphere, and about 0.25 for a spheroid of fineness ratio. The case of rotation of an airship hull is investigated and a coefficient is defined with the same advantages as the corresponding coefficient for rectilinear motion.

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### Calculations and Experimental Investigations on the Feed-Power Requirement of Airplanes with Boundary-Layer Control

**Date:**September 1, 1947

**Creator:**Krueger, W.

**Description:**Calculations and test results are given about the feed-power requirement of airplanes with boundary-layer control. Curves and formulas for the rough estimate of pressure-loss and feed-power requirement are set up for the investigated arrangements which differ structurally and aerodynamically. According to these results the feed power for three different designs is calculated at the end of the report.

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### A Theoretical Investigation of the Drag of Generalized Aircraft Configurations in Supersonic Flow

**Date:**January 1, 1957

**Creator:**Graham, E. W.; Lagerstrom, P. A.; Licher, R. M. & Beane, B. J.

**Description:**It seems possible that, in supersonic flight, unconventional arrangements of wings and bodies may offer advantages in the form of drag reduction. It is the purpose of this report to consider the methods for determining the pressure drag for such unconventional configurations, and to consider a few of the possibilities for drag reduction in highly idealized aircraft. The idealized aircraft are defined by distributions of lift and volume in three-dimensional space, and Hayes' method of drag evaluation, which is well adapted to such problems, is the fundamental tool employed. Other methods of drag evaluation are considered also wherever they appear to offer amplifications. The basic singularities such as sources, dipoles, lifting elements and volume elements are discussed, and some of the useful inter-relations between these elements are presented. Hayes' method of drag evaluation is derived in detail starting with the general momentum theorem. In going from planar systems to spatial systems certain new problems arise. For example, interference between lift and thickness distributions generally appears, and such effects are used to explain the difference between the non-zero wave drag of Sears-Haack bodies and the zero wave drag of Ferrari's ring wing plus central body. Another new feature of the spatial ...

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### Turbulence in the Wake of a Thin Airfoil at Low Speeds

**Date:**January 1, 1957

**Creator:**Campbell, George S.

**Description:**Experiments have been made to determine the nature of turbulence in the wake of a two-dimensional airfoil at low speeds. The experiments were motivated by the need for data which can be used for analysis of the tail-buffeting problem in aircraft design. Turbulent intensity and power spectra of the velocity fluctuations were measured at a Reynolds number of 1.6 x 10(exp 5) for several angles of attack. Total-head measurements were also obtained in an attempt to relate steady and fluctuating wake properties. Mean-square downwash was found to have nearly the same dependence on vertical position in the wake as that shown by total-head loss. For this particular wing, turbulent intensity, integrated across the wake, increased roughly as the 3/2 power of the drag coefficient. Power-spectrum measurements indicated a decrease in frequency as wing angle of attack was increased. The average frequency in the wake was proportional to the ratio of mean wake velocity to wake width.

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### Two-Dimensional Wing Theory in the Supersonic Range

**Date:**June 1, 1949

**Creator:**Hoenl, H.

**Description:**The plane problem of the vibrating airfoil in supersonic flow is dealt with and solved within the scope of a linearized theory by the method of the acceleration potential.

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### Systematic Investigations of the Effects of Plan Form and Gap between the Fixed Surface and Control Surface on Simple Flapped Wings

**Date:**May 1, 1949

**Creator:**Goethert & Roeber

**Description:**Four component measurements of 12 wings of symmetric profile having flaps with chord ratios t(sub R)/t(sub L) = 0.3 and t(sub R)/t(sub L) = 0.2 are treated in this report. As a result of the investigations, the effects of plan form and gap between fixed surface and control surface have been clarified. Lift, drag, pitching moment, and hinge moment were measured in the control-surface deflection range: -23 deg < or = beta < or = 23 deg and the range of angle of attack: -20 deg < or = alpha < or = 20 deg. Six wings with flaps of small chord (t(sub R)/t(sub L) < 0.1) were investigated at large flap settings.

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### Test Report on Measurements on a Series of Tapered Wings of Small Aspect Ratio (Trapezoidal Wing with Fuselage)

**Date:**July 1, 1947

**Creator:**Lange & Wacke

**Description:**This is the second of a series of six reports dealing with three- and six-component measurements on the tapering series at small aspect ratio. The present report concerns the trapezoidal wing with fuselage.

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### Pressure-Distribution Measurements on a Straight and on a 35 Degree Swept-Back Tapered Wing

**Date:**January 1, 1947

**Creator:**Thiel, A. & Weissinger, J.

**Description:**The spanwise lift-distribution measurements in straight air flow on a straight and a 35 deg swept-back tapered wing (NACA airfoil section 0012) are compared with theory for two angles of attack each (alpha approx. 6 deg and alpha approx. 12 deg) in the unstalled range of flow. The complete pressure distribution for the greater of the two angles is indicated.

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### Experimental Flights for Testing of a Reactor as an Expedient for the Termination of Dangerous Spins

**Date:**July 1, 1949

**Creator:**Hoehler, P. & Koeppen, I. v.

**Description:**In the Institute for Flight Mechanics of the DVL a reactor arrangement with a maximum output of 100 kg was investigated as an expedient for the termination of dangerous spins on an airplane of the FW 56 type. reproduce the influence of a disturbance of the steady spin condition by a pitching or yawing moment. The tests were meant to reproduce the influence of a disturbance of the steady spin condition by a pitching and yawing moment.

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### Flow and Force Equations for a Body Revolving in a Fluid

**Date:**January 1, 1930

**Creator:**Zahm, A F

**Description:**Part I gives a general method for finding the steady-flow velocity relative to a body in plane curvilinear motion, whence the pressure is found by Bernoulli's energy principle. Integration of the pressure supplies basic formulas for the zonal forces and moments on the revolving body. Part II, applying this steady-flow method, finds the velocity and pressure at all points of the flow inside and outside an ellipsoid and some of its limiting forms, and graphs those quantities for the latter forms. Part III finds the pressure, and thence the zonal force and moment, on hulls in plane curvilinear flight. Part IV derives general equations for the resultant fluid forces and moments on trisymmetrical bodies moving through a perfect fluid, and in some cases compares the moment values with those found for bodies moving in air. Part V furnishes ready formulas for potential coefficients and inertia coefficients for an ellipsoid and its limiting forms. Thence are derived tables giving numerical values of those coefficients for a comprehensive range of shapes.

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### The Inertial Coefficients of an Airship in a Frictionless Fluid

**Date:**January 1, 1924

**Creator:**Bateman, H

**Description:**This report deals with the investigation of the apparent inertia of an airship hull. The exact solution of the aerodynamical problem has been studied for hulls of various shapes and special attention has been given to the case of an ellipsoidal hull. In order that the results for this last case may be readily adapted to other cases, they are expressed in terms of the area and perimeter of the largest cross section perpendicular to the direction motion by means of a formula involving a coefficient K which varies only slowly when the shape of the hull is changed, being 0.637 for a circular or elliptic disk, 0.5 for a sphere, and about 0.25 for a spheroid of fineness ratio 7. For rough purposes it is sufficient to employ the coefficients, originally found for ellipsoids, for hulls otherwise shaped. When more exact values of the inertia are needed, estimates may be based on a study of the way in which K varies with different characteristics and for such a study the new coefficient possesses some advantage over one which is defined with reference to the volume of fluid displaced. The case of rotation of an airship hull has been investigated ...

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### Effect of Ice and Frost Formations on Drag of NACA 65(sub 1) -212 Airfoil for Various Modes of Thermal Ice Protection

**Date:**June 1, 1953

**Creator:**Gray, V. H. & Von Glahn, U. H.

**Description:**The effects of primary and. runback icing and frost formations on the drag of an 8-foot-chord NACA 651-212 airfoil section were investigated over a range of angles of attack from 20 to 80 and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.25 to 1.4 grams per cubic meter and datum air temperatures of -30 to 30 F. The results showed that glaze-ice formations, either primary or runback, on the upper surface near the leading edge of the airfoil caused large and rapid increases in drag, especially at datum air temperatures approaching 32 F and in the presence of high rates of water catch. Ice formations at lower temperatures (rime ice) did not appreciably increase the drag coefficient over the initial (standard roughness) drag coefficient. Cyclic de-icing of the primary Ice formations on the airfoil leading-edge section permitted the drag coefficient to return almost to the bare airfoil drag value. Runback icing on the lower surface did not present a serious drag problem except when heavy spanwise ridges of runback ice occurred aft of the heatable area. Frost formations caused rapid and large increases in drag with incipient stalling of the airfoil.

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### Effect of Ice Formations on Section Drag of Swept NACA 63A-009 Airfoil with Partial-Span Leading-Edge Slat for Various Modes of Thermal Ice Protection

**Date:**March 15, 1954

**Creator:**VonGlahn, Uwe H. & Gray, Vernon H.

**Description:**The effects of primary and runback ice formations on the section drag of a 36 deg swept NACA 63A-009 airfoil section with a partial-span leading-edge slat were studied over a range of angles of attack from 2 to 8 deg and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.39 to 1.23 grams per cubic meter and datum air temperatures from 10 to 25 F. The results with slat retracted showed that glaze-ice formations caused large and rapid increases in section drag coefficient and that the rate of change in section drag coefficient for the swept 63A-009 airfoil was about 2-1 times that for an unswept 651-212 airfoil. Removal of the primary ice formations by cyclic de-icing caused the drag to return almost to the bare-airfoil drag value. A comprehensive study of the slat icing and de-icing characteristics was prevented by limitations of the heating system and wake interference caused by the slat tracks and hot-gas supply duct to the slat. In general, the studies showed that icing on a thin swept airfoil will result in more detrimental aerodynamic characteristics than on a thick unswept airfoil.

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### Effects of surface roughness and extreme cooling on boundary-layer transition for 15 deg cone-cylinder in free flight at Mach numbers to 7.6

**Date:**March 5, 1958

**Creator:**Rabb, L. & Krasnican, M. J.

**Description:**None

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### Effects of extreme surface cooling on boundary-layer transition

**Date:**October 1, 1957

**Creator:**Jack, J. R.; Wisniewski, R. J. & Diaconis, N. S.

**Description:**None

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### Correlations Among Ice Measurements, Impingement Rates Icing Conditions, and Drag Coefficients for Unswept NACA 65A004 Airfoil

**Date:**February 1, 1958

**Creator:**Gray, Vernon H.

**Description:**An empirical relation has been obtained by which the change in drag coefficient caused by ice formations on an unswept NACA 65AO04 airfoil section can be determined from the following icing and operating conditions: icing time, airspeed, air total temperature, liquid-water content, cloud droplet impingement efficiencies, airfoil chord length, and angles of attack. The correlation was obtained by use of measured ice heights and ice angles. These measurements were obtained from a variety of ice formations, which were carefully photographed, cross-sectioned, and weighed. Ice weights increased at a constant rate with icing time in a rime icing condition and at progressively increasing rates in glaze icing conditions. Initial rates of ice collection agreed reasonably well with values predicted from droplet impingement data. Experimental droplet impingement rates obtained on this airfoil section agreed with previous theoretical calculations for angles of attack of 40 or less. Disagreement at higher angles of attack was attributed to flow separation from the upper surface of the experimental airfoil model.

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### Wind-Tunnel Investigation of a Low-Drag Airfoil Section with a Double Slotted Flap

**Date:**September 1, 1943

**Creator:**Bogdonoff, Seymour M.

**Description:**Tests were made of an 0.309-chord double-slotted flap on an NACA 65, 3-118, a equals 1.0 airfoil section to determine drag, lift, and pitching-moment characteristics for a range of flap deflections. Results indicate that combination of a low-drag airfoil and a double-slotted flap, of which the two parts moved as a single unit, gave higher maximum lift coefficients than have been obtained with plain, split, or slotted flaps on low-drag airfoils. Pitching moments were comparable to those obtained with other high-lift devices on conventional airfoils for similar lift coefficients.

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### A Transonic Wind-Tunnel Investigation of the Longitudinal Aerodynamic Characteristics of a Model of the Lockheed XF-104 Airplane

**Date:**January 1, 1954

**Creator:**Hieser, Gerald & Reid, Charles F.

**Description:**The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached ...

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