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Considerations of the Total Drag of Supersonic Airfoil Sections

Description: The results of calculations of the viscous and pressure drags of some two-dimensional supersonic airfoils at zero lift are presented. The results indicate that inclusion of viscous drag alters many previous results regarding the desirability of certain airfoil shapes for securing low drags at supersonic speeds. At certain Reynolds and Mach numbers, for instance, a circular-arc airfoil may theoretically have less drag than the previously advocated symmetrical wedge-shape profile; although under different conditions, the circular-arc airfoil may have a higher drag.
Date: July 1947
Creator: Ivey, H. Reese & Klunker, E. Bernard
Item Type: Report
Partner: UNT Libraries Government Documents Department

Tables and Charts of Flow Parameters Across Oblique Shocks

Description: 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.
Date: August 1948
Creator: Neice, Mary M.
Item Type: Report
Partner: UNT Libraries Government Documents Department

Two-Dimensional Irrotational Transonic Flows of a Compressible Fluid

Description: 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.
Date: June 1948
Creator: Kuo, Yung-Huai
Item Type: Report
Partner: UNT Libraries Government Documents Department

Charts for the Computation of Equilibrium Composition of Chemical Reactions in the Carbon-Hydrogen-Nitrogen System at Temperatures from 2000 to 5000 Degrees K

Description: 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.
Date: July 1948
Creator: Huff, Vearl N. & Calvert, Clyde S.
Item Type: Report
Partner: UNT Libraries Government Documents Department

The calculation of span LCAD distributions of swept-back wings

Description: Span load distributions of swept-back wings have been calculated. The method used was to replace the wing with a bound vortex at the quarter-chord line and to calculate the downwash due to the system of bound and trailing vortices to conform at the three-quarter-chord line to the slope of the flat-plate wing surface. Results are given for constant-chord and 5:1 tapered plan forms, for sweep-back angles of 0 degrees, 30 degrees, and 45 degrees, and for aspect ratios of 3, 6, and 9. Some comments on the stalling of swept-back wings are included.
Date: December 1, 1941
Creator: Mutterperl, William
Item Type: Report
Partner: UNT Libraries Government Documents Department

Comparison of Vee-Type and Conventional Tail Surfaces in Combination with Fuselage and Wing in the Variable-Density Tunnel

Description: The pitching and the yawing moments of a vee-type and a conventional type of tail surface were measured. The tests were made in the presence of a fuselage and a wing-fuselage combination in such a way as to determine the moments contributed by the tail surfaces. The results showed that the vee-type tail tested, with a dihedral angle of 35.3 deg, was about 71 percent as effective in pitch as the conventional tail and had a yawing-moment to pitching-moment ratio of 0.3. The conventional tail, the panels of which were all congruent to those of the vee-type tail, had a yawing-moment to pitching-moment ratio of 0.48. These ratios are in fair agreement with values calculated by methods shown in this and previous reports. The values of the measured moments were reduced from 15 to 25 percent of the calculated value by fuselage interference.
Date: July 1, 1941
Creator: Greenberg, Harry
Item Type: Report
Partner: UNT Libraries Government Documents Department

Methods of analyzing wind-tunnel data for dynamic flight conditions

Description: The effects of power on the stability and the control characteristics of an airplane are discussed and methods of analysis are given for evaluating certain dynamic characteristics of the airplane that are not directly discernible from wind-tunnel tests alone. Data are presented to show how the characteristics of a model tested in a wind tunnel are affected by power. The response of an airplane to a rolling and a yawing disturbance is discussed, particularly in regard to changes in wing dihedral and fin area. Solutions of the lateral equations of motion are given in a form suitable for direct computations. An approximate formula is developed that permits the rapid estimation of the accelerations produced during pull-up maneuvers involving abrupt elevator deflections.
Date: October 1, 1941
Creator: Donlan, C J & Recant, I G
Item Type: Report
Partner: UNT Libraries Government Documents Department

The Lagrangian Multiplier Method of Finding Upper and Lower Limits to Critical Stresses of Clamped Plates

Description: The theory of Lagrangian multipliers is applied to the problem of finding both upper and lower limits to the true compressive buckling stress of a clamped rectangular plate. The upper and lower limits thus bracket the truss, which cannot be exactly found by the differential-equation approach. The procedure for obtaining the upper limit, which is believed to be new, presents certain advantages over the classical Raleigh-Rite method of finding upper limits. The theory of the lower-limit procedure has been given by Trefftz but, in the present application, the method differs from that of Trefftz in a way that makes it inherently more quickly convergent. It is expected that in other buckling problems and in some vibration problems problems the Lagrangian multiplier method finding upper and lower limits may be advantageously applied to the calculation of buckling stresses and natural frequencies.
Date: January 1, 1946
Creator: Hu, Pai C. & Budiansky, Bernard
Item Type: Report
Partner: UNT Libraries Government Documents Department

A Theoretical Investigation of Longitudinal Stability of Airplanes with Free Controls Including Effect of Friction in Control System

Description: The relation between the elevator hinge moment parameters and the control forces for changes in forward speed and in maneuvers is shown for several values of static stability and elevator mass balance. The stability of the short period oscillations is shown as a series of boundaries giving the limits of the stable regions in terms of the elevator hinge moment parameters. The effects of static stability, elevator moment of inertia, elevator mass unbalance, and airplane density are also considered. Dynamic instability is likely to occur if there is mass unbalance of the elevator control system combined with a small restoring tendency (high aerodynamic balance). This instability can be prevented by a rearrangement of the unbalancing weights which, however, involves an increase of the amount of weight necessary. It can also be prevented by the addition of viscous friction to the elevator control system provided the airplane center of gravity is not behind a certain critical position. For high values of the density parameter, which correspond to high altitudes of flight, the addition of moderate amounts of viscous friction may be destabilizing even when the airplane is statically stable. In this case, increasing the viscous friction makes the oscillation stable again. The condition in which viscous friction causes dynamic instability of a statically stable airplane is limited to a definite range of hinge moment parameters. It is shown that, when viscous friction causes increasing oscillations, solid friction will produce steady oscillations having an amplitude proportional to the amount of friction.
Date: January 1, 1944
Creator: Greenberg, Harry & Sternfield, Leonard
Item Type: Report
Partner: UNT Libraries Government Documents Department