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**Partner:**UNT Libraries Government Documents Department

**Decade:**1950-1959

**Collection:**Technical Report Archive and Image Library

### An investigation of the experimental aerodynamic loading on a model helicopter rotor blade

**Date:**May 1, 1953

**Creator:**Meyer, John R , Jr; Falabella, Gaetano, Jr & NACA

**Description:**None

**Contributing Partner:**UNT Libraries Government Documents Department

**Permallink:**digital.library.unt.edu/ark:/67531/metadc57186/

### Effect of ice formations on section drag of swept NACA 63A-009 airfoil with partical-span leading-edge slat for various modes of thermal ice protection

**Date:**March 15, 1954

**Creator:**Von Glahn, Uwe H & Gray, Vernon H

**Description:**Studies were made to determine the effect of ice formations on the section drag of a 6.9-foot-chord 36 degree swept NACA 63A-009 airfoil with partial-span leading-edge slat. In general, the icing of a thin swept airfoil will result in greater aerodynamic penalties than for a thick unswept airfoil. Glaze-ice formations at the leading edge of the airfoil caused large increases in section drag even at liquid-water content of 0.39 gram per cubic meter. The use of an ice-free parting strip in the stagnation region caused a negligible change in drag compared with a completely unheated airfoil. Cyclic de-icing when properly applied caused the drag to decrease almost to the bare-airfoil drag value.

**Contributing Partner:**UNT Libraries Government Documents Department

**Permallink:**digital.library.unt.edu/ark:/67531/metadc59121/

### Investigation of turbines for driving supersonic compressors II : performance of first configuration with 2.2 percent reduction in nozzle flow area / Warner L. Stewart, Harold J. Schum, Robert Y. Wong

**Date:**July 22, 1952

**Creator:**Stewart, Warner L; Schum, Harold J & Wong, Robert Y

**Description:**The experimental performance of a modified turbine for driving a supersonic compressor is presented and compared with the performance of the original configuration to illustrate the effect of small changes in the ratio of nozzle-throat area to rotor-throat area. Performance is based on the performance of turbines designed to operate with both blade rows close to choking. On the basis of the results of this investigation, the ratio of areas is concluded to become especially critical in the design of turbines such as those designed to drive high-speed, high-specific weight-flow compressors where the turbine nozzles and rotor are both very close to choking.

**Contributing Partner:**UNT Libraries Government Documents Department

**Permallink:**digital.library.unt.edu/ark:/67531/metadc59457/

### On the Instability of Methods for the Integration of Ordinary Differential Equations

**Date:**April 1, 1956

**Creator:**Rutishauser, Heinz

**Description:**Examples and a criterion for stability of integration methods is provided. The criterion is applied to well-known integration formulas.

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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63924/

### Free Convection Under the Conditions of the Internal Problem

**Date:**April 1, 1958

**Creator:**Ostroumov, G. A.

**Description:**Convection is called free is the stresses (including the normal pressure) to which the fluid is subjected at its boundaries do not perform mechanical work, that is, if all the boundaries of the fluid are stationary. The case where this is not true is termed forced convection. It corresponds to the action on the fluid of some mechanical suction pumping the fluid.

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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63997/

### Evaporation, Heat Transfer, and Velocity Distribution in Two-Dimensional and Rotationally Symmetrical Laminar Boundary-Layer Flow

**Date:**February 1, 1958

**Creator:**Froessling, Nils

**Description:**The fundamental boundary layer equations for the flow, temperature and concentration fields are presented. Two dimensional symmetrical and unsymmetrical and rotationally symmetrical steady boundary layer flows are treated as well as the transfer boundary layer. Approximation methods for the calculation of the transfer layer are discussed and a brief survey of an investigation into the validity of the law that the Nusselt number is proportional to the cube root of the Prandtl number is presented.

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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63998/

### The Turbulent Boundary Layer on a Rough Curvilinear Surface

**Date:**September 1, 1958

**Creator:**Droblenkov, V. F.

**Description:**A number of semiempirical approximate methods exist for determining the characteristics of the turbulent boundary layer on a curvilinear surface. At present, among these methods, the one proposed by L. G. Loitsianskii is given frequent practical application. This method is sufficiently effective and permits, in the case of wing profiles with technically smooth surfaces, calculating the basic characteristics of the boundary layer and the values of the overall drag with an accuracy which suffices for practical purposes. The idea of making use of the basic integral momentum equation ((d delta(sup xx))/dx) + ((V' delta(sup xx))/V) (2 + H) = (tau(sub 0))/(rho V(exp 2)) proves to be fruitful also for the solution of the problems in the determination of the characteristics of the turbulent boundary layer on a rough surface.

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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63908/

### Stability of Cylindrical and Conical Shells of Circular Cross Section, with Simultaneous Action of Axial Compression and External Normal Pressure

**Date:**April 1, 1958

**Creator:**Mushtari, K. M. & Sachenkov, A. V.

**Description:**We consider in this report the determination of the upper limit of critical loads in the case of simultaneous action of a compressive force, uniformly distributed over plane cross sections, and of isotropic external normal pressure on cylindrical or conical shells of circular cross section. As a starting point we use the differential equations for neutral equilibrium of conical shells which have been used for the solution of the problem of stability of conical shells under torsion and under axial compression; upon solution of the problem it is possible to satisfy all boundary conditions, in contrast to the report where no attention is paid to the fulfillment of the boundary conditions, and to the report where only part of the boundary conditions are satisfied by solution of the problem according to Galerkin's method. Approximate formulas are used for the determination of the critical external normal pressure with simultaneous action of longituninal compression. Let us note that the formulas suggested in reference 5 are not well founded and may lead, in a number of cases, to a substantial mistake in the magnitude of the critical load.

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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63878/

### 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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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63919/

### 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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**Permallink:**digital.library.unt.edu/ark:/67531/metadc63891/