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The paper investigates a two-dimensional oscillating cascade with a subsonic leading edge locus in a supersonic flow. The blades are assumed to be of small thickness and camber, and are undergoing small amplitude-harmonic oscillations. The problem is reduced to the solution of a functional integral equation, and an expression is given for the kernel function.
An experimental film cooling study was conducted on a 3x size model turbine vane. Injection at the leading edge was from a single row of holes angled in a spanwise direction for two configurations of holes at 18 or 35 deg to the surface. The reduction in the local Stanton number for injection at a coolant-to-mainstream density ratio of 2.18 was calculated from heat flux measurements downstream of injection. Results indicate that optimum cooling occurs near a coolant-to-mainstream velocity ratio of 0.5. Shallow injection angles appear to be most beneficial when injecting into a highly accelerated mainstream.
A vortex-lattice method for predicting the aerodynamics of wings having separation at the sharp edges in incompressible flows is extended to compressible subsonic flows using a modified Prandtl-Glauert transformation. Numerical results showing the effect of freestream Mach number on the aerodynamic coefficients are compared with available experimental data for several planforms. It is shown that the proposed method is suitable for predicting the aerodynamic loads on low-aspect wings at moderate angles of attack for high subsonic freestream Mach number. The method is limited to angles of attack up to 12 deg for high subsonic freestream Mach number and to angles of attack up to 20 deg for Mach number not exceeding 0.5.
An experimental investigation was conducted to model the film-cooling performance for a turbine-vane leading edge using the stagnation region of a cylinder in cross flow. Experiments were conducted with a single row of spanwise-angled coolant holes for a range of the coolant blowing ratio with a freestream-to-wall temperature ratio of about 2.1 and a Reynolds number of 170,000, characteristic of the gas-turbine environment. Data from local heat-flux measurements are presented for coolant-hole injection angles of 25, 35, and 45 deg with the row of holes located at three positions relative to the stagnation line on the cylinder. Results show the spanwise (hole-to-hole) variation of heat-flux reduction due to film cooling and indicate conditions for the optimum film-cooling performance.
A flow-visualization study has shown that strong Karman vortices develop behind the blunt trailing edge of a plate when the free-stream velocities over both surfaces are equal and that the vortices tend to disappear when the surface velocities are unequal. This observation provides an explanation for the occurence and disappearance of certain discrete tones often found to be present in the noise spectra of coaxial jets. Both the vortex formation and the tones occur at a Strouhal number based on the lip thickness and the average of the external steady-state velocities of about 0.2. Results from theoretical calculations of the vortex formation, based on an inviscid incompressible analysis of the motion of point vortices, were in good agreement with the experimental observations.
Sequential edge-tracking algorithm employs circular scanning to point permit effective real-time tracking of coastlines and rivers from earth resources satellites. Technique eliminates expensive high-resolution cameras. System might also be adaptable for application in monitoring automated assembly lines, inspecting conveyor belts, or analyzing thermographs, or x ray images.
A nonlinear procedure for computing the pressure distribution on an airfoil with a trailing edge jet flap is described. The method is not restricted to thin airfoils or shallow jet-deflection angles. Correlation with experiment indicates that the characteristics of the pressure distribution are predicted by the theory, but the effect of entrainment is overpredicted with the entrainment coefficient used.
The plane elastostatic problem of internal and edge cracks in an infinite orthotropic strip is considered. The problems for the material types 1 and 2 are formulated in terms of singular integral equations. For the symmetric case the stress intensity factors are calculated and are compared with the isotropic results. The results show that because of the dependence of the Fredholm kernels on the elastic constants, unlike the crack problem for an infinite plane, in the strip the stress intensity factors are dependent on the elastic constants and are generally different than the corresponding isotropic results.
The plane elastostatic problem of internal and edge cracks in an infinite orthotropic strip is considered. The problems for the material types I and II are formulated in terms of singular integral equations. For the symmetric case the stress intensity factors are calculated and are compared with the isotropic results. The results show that because of the dependence of the Fredholm kernels on the elastic constants, unlike the crack problem for an infinite plane, in the strip the stress intensity factors are dependent on the elastic constants and are generally different from the corresponding isotropic results.
Potentiostatic determinations in various mixtures of FeCl2-FeCl3 with excess HCl show rest potentials that are 0.1 V less electropositive than the theoretical values from the formulated ratios of FeII to FeIII (probably as a result of complexing). The standard rate constant sub s. ranges between 0.0001 and 0.000 cm/sec. Tafel slopes b of roughly 0.12V per decade indicate single-electron exchange kinetics. No significant trend in either b or sub s was attributed to mixture composition. The higher sub s. values occurred with and edge-on pyrolytic graphite that had undergone a permanent surface change.
The effects of wing trailing-edge control surfaces on the static transonic aerodynamic characteristics of a transport configuration with a supercritical wing were studied. The configuration was tested with both an area-ruled fuselage and a cylindrical fuselage. The Mach number range was from 0.80 to 0.96 and the angle of attack range was from -1 deg to 12 deg. The Reynolds number was 1,580,000 based on the mean aerodynamic chord. Tabular data are presented.
The problem of determining the effect of laminar boundary layers on the lift of thin wings in subsonic flow at high Reynolds numbers is considered. The boundary value problem is formulated in the framework of the triple-deck theory of Brown and Stewartson. The resulting fourth-order boundary value was solved by an iterative finite-difference technique. An inverse iteration procedure provides proper treatment of the trailing-edge singularity, and asymptotic far-field expansions and coordinate stretchings are used to deal with the problem of the slow algebraic decay of the solution.
A first order theory of the edge-illuminated p(+)-n-n(+) silicon solar cell under very high injection levels has been derived. The very high injection level illuminated J-V characteristic is derived for any general base width to diffusion length (W/L) ratio and it includes the minority carrier reflection by the n-n(+) high-low junction. The beneficial effects of the high-low junction are shown to be significant until extremely high injection levels are reached. The theoretical dependencies of Jsc and Voc on temperature, incident intensity, and base resistivity are derived and discussed in detail. Some experimental results are given and these are discussed in relation to the theory.
The plane elastostatic problem of internal and edge cracks in an infinite orthotropic strip is considered. The problems for the material types I and II are formulated in terms of singular integral equations. For the symmetric case the stress-intensity factors are calculated and are compared with the isotropic results. The results show that because of the dependence of the Fredholm kernels on the elastic constants in the strip (unlike the crack problem for an infinite plane) the stress-intensity factors are dependent on the elastic constants and are generally different from the corresponding isotropic results
The problem of a semi infinite strip containing an edge crack is considered. It is assumed that the strip is loaded by a frictionless rigid wedge pressed into the crack. The resulting crack contact problem is formulated in terms of a system of singular integral equations. The behavior of the solution near the singular points is studied in detail. A series of numerical examples are given and the results are compared with those obtained by the method of boundary collocation and by the simple beam theory.
Stress-intensity factors, calculated by a three-dimensional finite-element analysis, for shallow or deep quarter-elliptical corner cracks at the edge of a hole in a finite-thickness plate are presented. The plate was subjected to remote uniform tension, remote bending, or simulated pin loading in the hole. The crack depth-to-plate thickness ranged from 0.2 to 0.8, while the ratio of crack depth-to-plate crack length ranged from 0.2 to 2. The ratio of hole radius-to-plate thickness was held at 0.5. To verify the accuracy of the three-dimensional finite-element models empolyed, convergence studies were conducted (number of degrees of freedom ranged from 4400 to 9300). The stress-intensity factor variations along the crack front are presented and compared with other solutions from the literature.
A method for rapidly estimating the overall forces and moments at supercritical speeds, below drag divergence, of transport configurations with supercritical wings is presented. The method was also used for estimating the rolling moments due to the deflection of wing trailing-edge controls. This analysis was based on a vortex-lattice technique modified to approximate the effects of wing thickness and boundary-layer induced camber. Comparisons between the results of this method and experiment indicate reasonably good correlation of the lift, pitching moment, and rolling moment. The method required much less storage and run time to compute solutions over an angle-of-attack range than presently available transonic nonlinear methods require for a single angle-of-attack solution.
A numerical method for the estimation of leading edge thrust for supersonic wings of arbitrary planform was developed and was programmed as an extension to an existing high speed digital computer method for prediction of wing pressure distributions. The accuracy of the method was assessed by comparison with linearized theory results for a series of flat delta wings. Application of the method to wings of arbitrary planform, both flat and cambered, is illustrated in several examples.