SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The results reviewed are divided into two categories: those that relate two-dimensional shell theory to three-dimensional elasticity theory and those concerned with shell theory per se. In the second category results for general elastic systems that carry over, by specialization or analogy, to shells and results that are unique to shell theory itself are considered.
The sensitivity of silica heat shield requirements to gap width, tile edge radius, and heat transfer distribution within tile gaps was investigated. A two-dimensional thermal model was modified and used to determine the effect of two dimensional heat transfer distributions at high temperature reusable surface insulation edges on shuttle thermal protection system (TPS) requirements. The sensitivity of TPS requirements to coating thickness, emissivity, substructure thickness, and changes in gap heating for several locations on shuttle was also studied. An inverse solution technique was applied to temperature data obtained in the Ames 20 MW turbulent duct in order to examine the effect of tile edge radius on TPS requirements. The derived heating values were then used to predict TPS requirements. Results show that increasing tile radius reduces TPS requirements.
An interactive procedure was developed for supersonic viscous flows that can be used for either two-dimensional or axisymmetric configurations. The procedure is directed to supersonic internal flows as well as those supersonic external flows that require consideration of mutual interaction between the outer flow and the boundary layer flow. The flow field is divided into two regions: an inner region which is highly viscous and mostly subsonic and an outer region where the flow is supersonic and in which viscous effects are small but not negligible. For the outer region a numerical solution is obtained by applying the method of characteristics to a system of equations which includes viscous and conduction transport terms only normal to the streamlines. The inner region is treated by a system of equations of the boundary layer type that includes higher order effects such as longitudinal and transverse curvature and normal pressure gradients. These equations are coupled and solved simultaneously in the physical coordinates by using an implicit finite difference scheme. This system can also be used to calculate laminar and turbulent boundary layers using a scalar eddy viscosity concept.
Tests were conducted to demonstrate a rapid scan two dimensional laser velocimeter (LV) measurement technique for aerial applications research. The LV system is capable of simultaneously measuring both vertical and axial flow velocity components in a near or far field vortex system. Velocity profiles were successfully measured in the wake vortex of a representative agricultural aircraft model, with the vortex system rapidly transporting in ground effect. Results indicate that the laser velocimetry technique can provide quantitative information of wake vortex characteristics in ground effect.
An interactive numerical procedure has been developed for supersonic viscous flows (either two-dimensional or axisymmetric configurations). The flow field is divided into two regions: (1) an inner region which is highly viscous and mostly subsonic, and (2) an outer region where the flow is supersonic and in which viscous effects are small, but not negligible. This paper presents a detailed description of: I. Outer Region - numerical solution obtained by applying the method of characteristics to a system of equations which includes viscous and conduction transport terms only normal to the streamlines; II. Inner Region - treated by a system of equations of the boundary-layer type that includes higher order effects, such as longitudinal and transverse curvature and normal pressure gradients (equations are coupled and solved simultaneously in physical coordinates, using an implicit finite-difference scheme); III. Interactive Procedure - in the interaction mode, the two regions are coupled iteratively along a matching line, where the Mach number is of the order of 1.2.
A ceramic/metallic aircraft gas turbine outer gas path seal designed for improved engine performance was studied. Transient temperature and stress profiles in a test seal geometry were determined by numerical analysis. During a simulated engine deceleration cycle from sea-level takeoff to idle conditions, the maximum seal temperature occurred below the seal surface, therefore the top layer of the seal was probably subjected to tensile stresses exceeding the modulus of rupture. In the stress analysis both two- and three-dimensional finite element computer programs were used. Predicted trends of the simpler and more easily usable two-dimensional element programs were borne out by the three-dimensional finite element program results.
A computational scheme and a standard program is proposed for solving systems of nonstationary spatially one-dimensional nonlinear differential equations using Newton's method. The proposed scheme is universal in its applicability and its reduces to a minimum the work of programming. The program is written in the FORTRAN language and can be used without change on electronic computers of type YeS and BESM-6. The standard program described permits the identification of nonstationary (or stationary) solutions to systems of spatially one-dimensional nonlinear (or linear) partial differential equations. The proposed method may be used to solve a series of geophysical problems which take chemical reactions, diffusion, and heat conductivity into account, to evaluate nonstationary thermal fields in two-dimensional structures when in one of the geometrical directions it can take a small number of discrete levels, and to solve problems in nonstationary gas dynamics.
An apparatus for experimental research into unsteady transonic flows is described. The apparatus, as installed in the NASA-Ames 11 by 11 Foot Transonic Wind Tunnel, can impart full two-degree-of-freedom motions at reduced frequencies to 0.3, oscillatory amplitudes to tn-2 degs, mean angles to 12 degs, Mach numbers to 1.4 and Reynolds numbers to 12x10. The test wing is fully instrumented for dynamic waveform measurements and the data can be acquired, processed, and displayed in real-time with a new computational data acquisition system. Following a description of the apparatus, sample data from a recently completed test program is presented.
The development of a numerical simulation of time dependent, turbulent, compressible flow about two dimensional multi-element airfoils of arbitrary shape is described. The basis of this simulation is a technique of automatic numerical generation of coordinate systems fitted to the multiple bodies regardless of their number or shape. Procedures developed whereby the coordinate lines are automatically concentrated in the boundary layer at any Reynolds number are discussed. The compressible turbulent solution involves an algebraic eddy viscosity turbulence model. The laminar version was run for transonic flow at free stream Mach numbers up to 0.9.
A two dimensional oscillating airfoil test apparatus is presented as a method of measuring unsteady aerodynamic forces on an airfoil or rotor blade section. The oscillating airfoil test rig, which is being built for use in an 11 X 11-foot transonic wind tunnel (speed range M = 0.4 - 1.4), will allow determination of unsteady loadings and detailed pressure distributions on representative airfoil sections undergoing simulated pitching and flapping motions. The design details of the motion generating system and supporting structure are presented. This apparatus is now in the construction phase.
Three-dimensional fully self-consistent computer models were used to determine the evolution of galaxies consisting of 100 000 simulation stars. Comparison of two-dimensional simulations with three-dimensional simulations showed only a very slight stabilizing effect due to the additional degree of freedom. The addition of a fully self-consistent, nonrotating, exponential core/halo component resulted in considerable stabilization. A second series of computer experiments was performed to determine the collapse and relaxation of initially spherical, uniform density and uniform velocity dispersion stellar systems. The evolution of the system was followed for various amounts of angular momentum in solid body rotation. For initally low values of the angular momentum satisfying the Ostriker-Peebles stability criterion, the systems quickly relax to an axisymmetric shape and resemble elliptical galaxies in appearance. For larger values of the initial angular momentum bars develop and the systems undergo a much more drastic evolution.
An analysis is presented for calculating the flow field in supersonic mixed-compression aircraft inlets operating at angle of attack. The flow field is computed by a steady three-dimensional bicharacteristic method. The bow shock wave and the reflected internal shock wave system are computed by a three-dimensional discrete shock wave fitting procedure. Viscous and thermal diffusion may be included as source terms in the bicharacteristic method. A production type computer program capable of determining the flow field in a variety of axisymmetric mixed-compression supersonic inlets is available. The results of the present analysis agree well with those produced by the two-dimensional method of characteristics when axisymmetric flow fields are computed. For three-dimensional flow fields, the results of the present analysis agree well with experimental data except in regions of high viscous interaction and boundary layer removal. The present analysis does not compute the boundary layer, nor does it account for boundary layer bleed.
A theoretical analysis is presented yielding sets of partial differential equations for determination of turbulent aerodynamic flowfields in the vicinity of an airfoil trailing edge. A four phase interaction algorithm is derived to complete the analysis. Following input, the first computational phase is an elementary viscous corrected two dimensional potential flow solution yielding an estimate of the inviscid-flow induced pressure distribution. Phase C involves solution of the turbulent two dimensional boundary layer equations over the trailing edge, with transition to a two dimensional parabolic Navier-Stokes equation system describing the near-wake merging of the upper and lower surface boundary layers. An iteration provides refinement of the potential flow induced pressure coupling to the viscous flow solutions. The final phase is a complete two dimensional Navier-Stokes analysis of the wake flow in the vicinity of a blunt-bases airfoil. A finite element numerical algorithm is presented which is applicable to solution of all partial differential equation sets of inviscid-viscous aerodynamic interaction algorithm. Numerical results are discussed.
A second-order time-accurate and spatially factored algorithm was used in a finite difference scheme for the numerical solution of the time-dependent, incompressible, two dimensional Navier-Stokes equations in conservation-law form using vorticity and stream function variables. The systems of equations are solved at each time step by an iterative technique. Numerical results were obtained for a circular cylinder at a Reynolds number of 15, and an NACA 0012 airfoil at zero angle of attack at Reynolds numbers of 10 to the third and 10 to the fourth powers. The results are in agreement with another numerical technique, and the computing time required to obtain the steady state solution at the Reynolds number of 10 to the 4th power was 49.7 sec on CDC 7600 computer using a 65 x 84 computational grind.
A description is given of three programs catalogued in the form of object modules in the library of a system for processing photographic images computer. PFT is the subprogram of the multi-dimensional BPF of real-valued information, in the operative computer memory. INRECO is a subprogram-interface between the real and complex formats for representing two-dimensional spectra and images. FFT2 is a subprogram for calculating the correlation functions of the image using the previous subprograms.
The inlet was investigated with both a two-oblique-shock and an isentropic compression ramp. Provisions for a variety of bypass systems were incorporated and none decreased pressure recovery by more than about 1 percent. By utilizing throat boundary-layer control, critical pressure recoveries of 64 and 70 percent were obtained with the two-shock and isentropic ramps, respectively. Diffuser-exit air distortion was about 5 percent at critical operation Theoretical cowl pressure drag was equal to about 10 percent of the net thrust of an assumed engine.
The extension to a compressible fluid of Lagally's theorem on the moment on a body in an incompressible fluid and Poggi's method of treating the flow of compressible fluids are employed for the determination of the effect of compressibility on the moment of an arbitrary body. Only the case of the two-dimensional subsonic flow of an ideal compressible fluid is considered. As examples of the application of the general theory, two well-known systems of profiles are treated; namely the elliptic profile and the symmetrical Joukowski profiles with sharp trailing edges. The effect of compressibility on the position of the center of pressure is also discussed. Several numerical examples of both the elliptic and the Joukowski profiles are given.
For the purpose of testing the theory of an oscillating airfoil of two degrees of freedom, a wing was mounted in the wind tunnel between two walls in such a manner that it could execute vertical (flexural) oscillations as well as torsional oscillations about an arbitrary axis parallel to the span. It was possible to vary the inertia and elasticity parameters and also to increase artificially the negligibly small natural damping of the system. The oscillations were recorded to a strongly magnified scale. The experimentally determined critical (or flutter) velocities fully agree with the theoretical ones of Wagner and Kussner within the limits of computational and measuring accuracy. An extremely narrow wing without end walls (three dimensional problem) showed the same oscillations as one with end walls (two-dimensional problem).