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Twenty-five ring-stiffened circular cylinders were loaded to failure in bending. The results are presented in the form of design curves which are applicable to cylinders with heavy rings that fail as a result of local buckling.
Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.88, a Reynolds number of 129,000, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and a simple modification of the Newtonian flow theory. An evaluation of the crossflow theory is made through comparison of present results with available crossflow Mach number drag coefficients.
The proximity of the ground to an annular nozzle can result in thrust augmentation of up to 50 percent. The magnitude of thrust augmentation was a function of ground-to-nozzle distance, nozzle configuration, and nozzle pressure ratio. With a conventional circular nozzle a thrust reduction of 15 to 40 percent due to ground effect was obtained. The degree of thrust reduction was primarily a function of nozzle pressure ratio and ground-to-nozzle distance.
Thirty-four circular cylinders stiffened by heavy rings were loaded to failure in combined bending and shear with a transverse shear load. The results are presented in the form of an interaction curve which is applicable to the design of ring-stiffened cylinders that fail as a result of local buckling. (author)
Section data obtained from wind-tunnel tests at Mach number 4.04 and Reynolds numbers of about 5.0 times 10 to the 6th power and 8.4 times 10 to the 6th power are presented for 6- and 9-percent-thick symmetrical circular-arc airfoils with 30 percent chord trailing-edge flaps. Pressure distributions, Schlieren photographs, and force and moment coefficients are included.
Local heat-transfer coefficients, temperature recovery factors, and pressure distributions were measured on the front side of a circular cylinder at a nominal Mach number of 3.9 over a range of free-stream Reynolds numbers from 2.1 x 10 to the 3rd power to 6.7 x 10 to the 3rd power and yaw angles from zero degrees to 44 degrees. Yawing the cylinder reduced the heat-transfer coefficients and the pressure drag coefficients. The amount of reduction may be predicted by a theory presented herein.
The design of curved sheet panels to resist shear involves a consideration of several factors: the buckling resistance of the sheet, the stress at which buckling becomes permanent, and the strength which may be developed beyond the buckling limit by tension-field action. Although some experimental as well as theoretical work has been done on the buckling and tension-field phases of this problem, neither of these types of action appears to be very well understood. The problem is of sufficient importance from the standpoint of aircraft design, it is believed, to warrant further experimental investigation. This report presents the results of the first series of torsion tests of stiffened circular cylinders to be completed in connection with this study at Aluminum Research Laboratories. (author)
The properties of conical flow without axial symmetry are analyzed. The flow around cones of circular cross section at small angles of attack is determined by correctly considering the effect of the entropy gradients in the flow.
Contains theoretical and experimental analysis of circular inlets having a central body at Mach numbers of 3.30, 2.75, and 2.45. The inlets have been designed in order to have low drag and high pressure recovery. The pressure recoveries obtained are of the same order of magnitude as those previously obtained by inlets having very large external drag.
Report presents the results of a wind-tunnel investigation of the effect of wing-fuselage interference on lateral-stability characteristics made in the NACA 7 by 10-foot wind tunnel on four fuselages and two fins, representing high-wing, low-wing, and midwing monoplanes. The fuselages are of circular and elliptical cross section. The wings have rounded tips and, in plan form, one is rectangular and the three are tapered 3:1 with various amounts of sweep. The rate of change in the coefficients of rolling moment, yawing moment, and lateral force with angle of yaw is given in a form to show the increment caused by wing-fuselage interference for the model with no fin and the effect of wing-fuselage interference on fin effectiveness. Results for the fuselage-fin combination and the wing tested alone are also given.
The potential function for flow, with circulation, of a compressible fluid about a circular cylinder is obtained in series form including terms of the orders of m(4) where m is the Mach number of the free stream. The resulting equations are used to obtain pressure coefficients as a function of Mach number at a point on the surface of the cylinder for different values of circulation. The coefficients derived are compared with the Glauert-Prandtl and Karman-Tsien approximations which are functions of the pressure coefficients of an incompressible fluid. For the cases considered, the values of the pressure coefficients computed from the theory were found to be somewhere between the two approximations, the first underestimating and the second overestimating it.
A method is presented for determining the jet-boundary and plan-form corrections necessary for application to test data for a partial-span model with a reflection plane, an end plate, or no end plate in a closed circular wind tunnel. Examples are worked out for a partial-span model with each of the three end conditions in the Langley 19-foot pressure tunnel and the corrections are applied to measured values of lift, drag, pitching-moment, rolling-moment, and yawing-moment coefficients.
The results of a theoretical and experimental investigation of wall interference for an airfoil spanning a closed-throat circular wind tunnel are presented. Analytical equations are derived which relate the characteristics of an airfoil in the tunnel at subsonic speeds with the characteristics in free air. The analysis takes into consideration the effect of fluid compressibility and is based upon the assumption that the chord of the airfoil is small as compared with the diameter of the tunnel. The development is restricted to an untwisted, constant-chord airfoil spanning the middle of the tunnel. Brief theoretical consideration is also given to the problem of choking at high speeds. Results are then presented of tests to determine the low-speed characteristics of an NACA 4412 airfoil for two chord-diameter ratios. While, on the basis of these experiments, no appraisal is possible of the accuracy of the corrections at high speeds, the data indicate that at low Mach numbers the analytical results are valid, even for relatively large values of the chord-diameter ratio.
Tests have been made at high speeds to determine the drag of models, simulating propeller shanks, in the form of a circular cylinder and three airfoils, the NACA 16-025, the NACA 16-040, and the NACA 16-040 with the rear 25 percent chord cut off. All the models had a maximum thickness of 4 1/2 inches to conform with average propeller-shank dimensions and a span of 20 1/4 inches. For the tests the models were supported perpendicular to the lower surface of the wing of an XP-51 airplane. A wake-survey rake mounted below the wing directly behind the models was used to determine profile drag of Mach numbers of 0.3 to 0.8 over a small range of angle of attack. The drag of the cylinder was also determined from pressure-distribution and force measurements.
An investigation has been made in the Langley two-dimensional low-turbulence tunnel and in the Langley two-dimensional low-pressure tunnel of 6- and 10-percent-thick symmetrical circular-arc airfoil sections at low Mach numbers and several Reynolds numbers. The airfoils were equipped with 0.15-chord plain leading-edge flaps and 0.20-chord plan trailing-edge flaps. The section lift and pitching-moment characteristics were determined for both airfoils with the flaps deflected individually and in combination. The section drag characteristics were obtained for the 6-percent-thick airfoil with the flaps partly deflected as low-drag-control flaps and for airfoils with the flaps neutral. Surface pressures were measured on the 6-percent-thick airfoil section with the flaps deflected either individually or in appropriate combination to furnish flap load and hinge-moment data applicable to the structural design of the airfoil. A generalized method is developed that permits the determination of the chordwise pressure distribution over sharp-edge airfoils with plain leading-edge flaps and plain trailing-edge flaps of arbitrary size and deflection.
A method is presented for analyzing the stresses about cutouts in circular semimonocoque cylinders with flexible rings. The method involves the use of so-called perturbation stress distributions which are superposed on the stress distribution that would exist in the structure with no cutout in such a way as to give the effects of a cutout. The method can be used for any loading case for which the structure without the cutout can be analyzed and is sufficiently versatile to account for stringer and shear reinforcement about the cutout.
A preliminary theoretical investigation of the panel flutter and divergence of infinitely long, unstiffened and ring-stiffened thin-walled circular cylinders is described. Linearized unsteady potential-flow theory is utilized in conjunction with Donnell's cylinder theory to obtain equilibrium equations for panel flutter. Where necessary, a simplified version of Flugge's cylinder theory is used to obtain greater accuracy. By applying Nyquist diagram techniques, analytical criteria for the location of stability boundaries are derived. A limited number of computed results are presented.
Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86, a Reynolds number of 129,000 based on diameter, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and with a simple modification of the Newtonian flow theory. The comparison of experimental results shows that either theory gives adequate general aerodynamic characteristics but that the modified Newtonian theory gives a more accurate prediction of the pressure distribution. The calculated crossflow drag coefficients plotted as a function of crossflow Mach number were found to be in reasonable agreement with similar results obtained from other investigations at lower supersonic Mach numbers. Comparison of the results of this investigation with data obtained at a lower Mach number indicates that the drag coefficient of a cylinder normal to the flow is relatively constant for Mach numbers above about 4.