Latest content added for UNT Digital Library Collection: National Advisory Committee for Aeronautics (NACA)http://digital.library.unt.edu/explore/collections/NACA/browse/?fq=untl_institution:UNTGD&fq=str_title_serial:NACA+Technical+Notes2014-09-25T20:32:43-05:00UNT LibrariesThis is a custom feed for browsing UNT Digital Library Collection: National Advisory Committee for Aeronautics (NACA)Two-Dimensional Irrotational Transonic Flows of a Compressible Fluid2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc172480/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc172480/"><img alt="Two-Dimensional Irrotational Transonic Flows of a Compressible Fluid" title="Two-Dimensional Irrotational Transonic Flows of a Compressible Fluid" src="http://digital.library.unt.edu/ark:/67531/metadc172480/thumbnail/"/></a></p><p>The methods of NACA TN No. 995 have been slightly modified and extended in include flows with circulation by considering the alteration of the singularities of the incompressible solution due to the presence of the hypergeometric functions in the analytic continuation of the solution. It was found that for finite Mach numbers the only case in which the nature of the singularity can remain unchanged is for a ratio of specific heats equal to -1. From a study of two particular flows it seems that the effect of geometry cannot be neglected, and the conventional "pressure-correction" formulas are not valid, even in the subsonic region if the body is thick, especially if there is a supersonic region in the flow.</p>Bending Tests of Metal Monocoque Fuselage Construction2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc172433/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc172433/"><img alt="Bending Tests of Metal Monocoque Fuselage Construction" title="Bending Tests of Metal Monocoque Fuselage Construction" src="http://digital.library.unt.edu/ark:/67531/metadc172433/thumbnail/"/></a></p><p>Study of the bending stress in smooth skin, aluminum alloy, true monocoque fuselage sections of varying ratio of diameter to thickness.</p>Aerodynamic Investigation of a Cup Anemometer2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc172432/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc172432/"><img alt="Aerodynamic Investigation of a Cup Anemometer" title="Aerodynamic Investigation of a Cup Anemometer" src="http://digital.library.unt.edu/ark:/67531/metadc172432/thumbnail/"/></a></p><p>Results of an investigation wherein the change of the normal force coefficient with Reynolds Number was obtained statically for a 15.5-centimeter hemispherical cup.</p>Tables and Charts of Flow Parameters Across Oblique Shocks2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc172508/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc172508/"><img alt="Tables and Charts of Flow Parameters Across Oblique Shocks" title="Tables and Charts of Flow Parameters Across Oblique Shocks" src="http://digital.library.unt.edu/ark:/67531/metadc172508/thumbnail/"/></a></p><p>Shock-wave equations have been evaluated for a range of Mach number in front of the shock from 1.05 to 4.0. Mach number behind the shock, pressure ratio, derivation of flow, and angle of shock are presented on charts. Values are also included for density ratio and change in entropy.</p>Two-Dimensional Subsonic Compressible Flows Past Arbitrary Bodies by the Variational Method2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc172521/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc172521/"><img alt="Two-Dimensional Subsonic Compressible Flows Past Arbitrary Bodies by the Variational Method" title="Two-Dimensional Subsonic Compressible Flows Past Arbitrary Bodies by the Variational Method" src="http://digital.library.unt.edu/ark:/67531/metadc172521/thumbnail/"/></a></p><p>Instead of solving the nonlinear differential equation which governs the compressible flow, an approximate method of solution by means of the variational method is used. The general problem of steady irrotational flow past an arbitrary body is formulated. Two examples were carried out, namely, the flow past a circular cylinder and the flow past a thin curved surface. The variational method yields results of velocity and pressure distributions which compare excellently with those found by existing methods. These results indicate that the variational method will yield good approximate solution for flow past both thick and thin bodies at both high and low Mach numbers.</p>Charts for the Computation of Equilibrium Composition of Chemical Reactions in the Carbon-Hydrogen-Nitrogen System at Temperatures from 2000 to 5000 Degrees K2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc100821/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc100821/"><img alt="Charts for the Computation of Equilibrium Composition of Chemical Reactions in the Carbon-Hydrogen-Nitrogen System at Temperatures from 2000 to 5000 Degrees K" title="Charts for the Computation of Equilibrium Composition of Chemical Reactions in the Carbon-Hydrogen-Nitrogen System at Temperatures from 2000 to 5000 Degrees K" src="http://digital.library.unt.edu/ark:/67531/metadc100821/thumbnail/"/></a></p><p>Charts are provided for the estimation and progressive adjustment of two independent variables on which the calculations are based. Additional charts are provided for the graphical calculation of the composition.</p>Tables for the Computation of Wave Drag of Arrow Wings of Arbitrary Airfoil Section2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc100822/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc100822/"><img alt="Tables for the Computation of Wave Drag of Arrow Wings of Arbitrary Airfoil Section" title="Tables for the Computation of Wave Drag of Arrow Wings of Arbitrary Airfoil Section" src="http://digital.library.unt.edu/ark:/67531/metadc100822/thumbnail/"/></a></p><p>Tables and computing instructions for the rapid evaluation of the wave drag of delta wings and of arrow wings having a ration of the tangent of the trailing-edge sweep angle to the tangent of the leading-edge sweep angle in the range from -1.0 to 0.8. The tables cover a range of both subsonic and supersonic leading edges.</p>Summary of 65-Series Compressor-Blade Low-Speed Cascade Data by Use of the Carpet-Plotting Technique2014-09-25T20:32:43-05:00http://digital.library.unt.edu/ark:/67531/metadc100823/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc100823/"><img alt="Summary of 65-Series Compressor-Blade Low-Speed Cascade Data by Use of the Carpet-Plotting Technique" title="Summary of 65-Series Compressor-Blade Low-Speed Cascade Data by Use of the Carpet-Plotting Technique" src="http://digital.library.unt.edu/ark:/67531/metadc100823/thumbnail/"/></a></p><p>Carpet plots included permit the selection of the blade camber and the design angle of attack required to fulfill a design vector diagram. Other carpet plots provide means for the prediction of off-design turning angles.</p>Spin tests of two models of a low-wing monoplane to investigate scale effect in the model test range2014-03-30T18:00:15-05:00http://digital.library.unt.edu/ark:/67531/metadc279675/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc279675/"><img alt="Spin tests of two models of a low-wing monoplane to investigate scale effect in the model test range" title="Spin tests of two models of a low-wing monoplane to investigate scale effect in the model test range" src="http://digital.library.unt.edu/ark:/67531/metadc279675/thumbnail/"/></a></p><p>None</p>The Unsteady Lift of a Finite Wing2014-03-30T18:00:15-05:00http://digital.library.unt.edu/ark:/67531/metadc279608/<p><a href="http://digital.library.unt.edu/ark:/67531/metadc279608/"><img alt="The Unsteady Lift of a Finite Wing" title="The Unsteady Lift of a Finite Wing" src="http://digital.library.unt.edu/ark:/67531/metadc279608/thumbnail/"/></a></p><p>Unsteady lift function for wings of finite aspect ratio have been calculated by approximate methods involving corrections of the aerodynamic inertia and of the angle of the infinite wing. The starting lift of the finite wing is found to be only slightly less than that of the infinite wing; whereas the final lift may be considerably less. The calculations indicate that the distribution of lift near the start is similar to the final distribution. Both the indicia and the oscillating lift functions are given. Approximate operational equivalents of the functions have been devised to facilitate the calculation of lift under various conditions of motion.</p>