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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.
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Elastic boundary conditions effect on natural frequencies and resonant displacement of thin isotropic circular cylindrical shell under concentrated load with harmonic time history
Blue edge of RR Lyrae instability strip, examining convection, radiation and surface boundary conditions effects
Cantilever column critical dynamic instability load under nonconservative follower force including thermomechanical coupling effect from boundary value problem formulation
Supersonic propulsion system inlet, engine and exhaust nozzle in wind tunnel and flight tests, discussing boundary layer effects on performance
Interaction regions created by the impingement of full span, externally generated, shock waves on a nozzle wall boundary layer were investigated. Incident shock strength was varied to produce unseparated, incipient, and fully separated flow fields. Significant departures from two-dimensionality were observed over the entire range of shock strengths tested and were identified with sidewall and corner boundary layer effects. However, comparisons of present centerline results with published two-dimensional data, obtained under similar test conditions and geometrical constraints, showed excellent agreement (e.g., incipient separation pressure levels, wall pressure distributions, free-interaction, and scale of the interaction region). This raises some question concerning the degree of two-dimensionality achieved in these previous investigations.
Rate coefficient data for the recombination reaction H + 02 + Ar yields H02 + Ar have been determined from studies of lean hydrogen-oxygen mixtures behind incident shock waves over the temperature range of 948 to 1125 K. Hydroxyl radical concentration profiles were measured by ultraviolet absorption spectroscopy, and rate data were obtained through analysis of induction time and exponential growth parameter data. Analysis of the data yielded a rate coefficient which was generally lower than most of the more recent values obtained from shock tube studies. The effect of boundary layer formation on the conditions behind the shock was also examined and found to be negligible.
The double-pulse, double-exposure holography technique was applied to visualize the flow field within a transonic compressor rotor with a tip speed of 1800 ft/sec. The principal objective was to visualize the shock waves created in the flow field which was supersonic relative to the rotating blade row. The upstream rotor blade bow shocks and, at high speed, the outermost portion of the leading edge passage shock were successfully observed in the holograms. Techniques were devised for locating these shocks in three dimensions, and the results were compared with theoretical predictions. Density changes between the two pulses due to motion of the shocks were large and, therefore, it was not possible to resolve the fringe systems in detail for the 100% speed conditions. However, gross features of the shocks were easily observed, and the upstream shocks were well displayed. In all cases the shock angles were somewhat larger than predicted by theory, and a distinct increase in angle near the outer wall was observed, which may be attributed to endwall boundary layer effects. The location and orientation of the observed leading edge passage shocks were in good agreement with static pressure contours obtained from measurements in the outer casing over the rotor tip.
Complete seasonal ERTS-1 coverage of Atlantic coastal wetlands from Delaware Bay to Georgia provides a basis for assessment of temporal data for wetland mapping, evaluation, and monitoring. Both MSS imagery and digital data have proved useful for gross wetland species delineation and determination of the upper wetland boundary. Tidal effects and (band to band or seasonal) spectral reflectance differences make it possible to type vegetatively coastal wetlands in salinity related categories. Management areas, spoil disposal sites, drainage ditches, lagoon-type developments and highway construction can be detected indicating a monitoring potential for the future. A northern test site (Maryland-Virginia) and a southern test site (Georgia-South Carolina), representing a range of coastal marshes from saline to fresh, were chosen for intensive study. Wetland maps were produced at various scales using both ERTS imagery (bands 5 and 7) and digital data (bands 4, 5 and 7).
The method of Fourier transforms is used to determine the kernel function which relates the pressure on a lifting surface to the prescribed downwash within the framework of Dowell's (1971) shear flow model. This model is intended to improve upon the potential flow aerodynamic model by allowing for the aerodynamic boundary layer effects neglected in the potential flow model. For simplicity, incompressible, steady flow is considered. The proposed method is illustrated by deriving known results from potential flow theory.
This paper describes an experiment in the application of holographic interferometry to the visualization of flow in a transonic compressor. The principal objective of the study was to visualize the shock waves created by the rotor blades and to determine their relative orientation in three dimensions. The technique employed double-pulse, double-exposure holography to record the flow field through a plastic window built into the outer case over the rotor blades. Illumination was by diffuse reflection from the interior surfaces of the compressor. No other optical components in or on the compressor were used. The resulting holograms successfully showed the upstream bow shocks and, at high speed, the outer-most portion of the leading edge passage shock. Techniques were devised for locating these shocks in three dimensions and the results were compared with theoretical predictions. In all cases, the shock angles were somewhat larger than predicted by theory, and a distinct increase in angle near the outer wall was observed, which may be attributed to end wall boundary layer effects.
Two computer programs determine effect of boundary layer on attenuation of sound in a circular duct lined with material used in noise suppresion in fan inlet and exhaust ducts of turbofan engines.
Preliminary analytical and experimental inlet forebody investigations have been conducted at Mach numbers of 6.0 and 8.5. The forebody design concept consisted of a sharp-nosed right circular cone followed by elliptical cross sections. This concept resulted in swept isentropic compression which would allow swept cowl leading edges. Measurements were made to define the condition of the inviscid flow field developed by the forebody, including flow profiles in the vicinity of cowl leading-edge stations, and the three-dimensional boundary-layer effects. The investigation verified some of the expected differences between the predicted and the experimental results.
A theoretical method is presented for the computation of the flow field about an axisymmetric body operating in a viscous, incompressible fluid. A potential flow method was used to determine the inviscid flow field and to yield the boundary conditions for the boundary layer solutions. Boundary layer effects in the forces of displacement thickness and empirically modeled separation streamlines are accounted for in subsequent potential flow solutions. This procedure is repeated until the solutions converge. An empirical method was used to determine base drag allowing configuration drag to be computed.
Technique analyzes viscosity-dependent aerodynamic characteristics of multielement infinite swept wings in incompressible flow. Use of source distributions rather than displacement thickness to represent boundary layer effect on potential flow and of iterative technique for matrix inversion reduces computer time for overall analysis.
Simultaneous time-resolved measurements of temperature, density, pitot pressure, and wall pressure in both air and O2 test gases were obtained in the Langley pilot model expansion tube. These tests show nonequilibrium chemical and vibrational relaxation significantly affect the test-flow condition. The use of an electromagnetic device to preopen the secondary diaphragm before the arrival of the primary shock wave resulted in an improvement in the agreement between the measured pitot pressure and the value inferred from measured density and interface velocity. Boundary-layer splitter plates used to reduce the wall boundary layer show that this disagreement in the measured and inferred pitot pressures is not a result of boundary-layer effects.
Results of aerodynamic heating tests conducted in October 1974 on a space shuttle orbiter model using the phase change paint technique are presented. The model was a 0.040 scale representation of the forward 50 percent of the orbiter. Surface roughness effects on boundary layer transition were investigated. Roughness was simulated by using steel balls varying in diameter from 0 (no balls) to 0.039 inch with 0.040 inch wide by 0.080 inch deep gaps. A nominal Mach number of 8 was tested with Reynolds number varying from 0.75 through 3.5 million per foot. Angle of attack was varied from 20 deg to 40 deg.
Numerical calculations have been made for the early stages of collapse of axisymmetric rotating protostars of 1, 2, and 5 solar masses. The calculations employ a range of values of total angular momentum, as well as two types of initial density distribution. The effects of boundary conditions are tested by using constant volume and constant surface pressure with identical initial conditions. The principal result of the calculations is that, in all cases tried, the collapse leads to the formation of a ring structure in the interior of the cloud, with a local density minimum at the center of the cloud. The rings approach equilibrium with a structure consistent with that of previous analytic determinations, after which they undergo further gravitational collapse. The collapse of a two-solar-mass cloud, similar to that assumed by Cameron and Pine (1973), does not appear to lead to the equilibrium nebula these authors construct.
Aerodynamic heating phase change paint tests for the space shuttle orbiter are reported. The model was a 0.040 scale representation of the forward 50% of the orbiter. Surface roughness effects on boundary layer transition were investigated. The roughness was simulated by steel balls 0.020 and 0.025 inch in diameter and a 0.25 in. diameter hole simulating the forward ET attach socket. A nominal Mach number of was tested with unit Reynolds number varying from 0.75 x 1 million ft through 3.5x 1 million ft. Angle of attack was varied from 20 degrees to 40 degrees.