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132 records · Page 8

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NASA NTRSDate not supplied

Dynamic Analysis of Reynolds Number Effects on Trailing Edge Transonic Vortex Shedding and Its Impact on Turbine Blade Aerodynamic Performance

In this work, we will discuss observations from images acquired from a time-resolved, high-speed self-aligned focusing Schlieren campaign that was performed at the CW-22 facility at NASA Glenn Research Center to understand the dynamic behavior of thick trailing-edge high-pressure turbine blades simulating a ceramic matrix composite (CMC) construction at high inlet turbulence conditions. For the CMC-9 blade, which has a trailing edge thickness of 9% of the axial chord, we identified a regime where an excessive total pressure loss (loss anomaly) is observed only for a narrow range of Reynolds numbers at a fixed exit Mach number of 0.74. The loss anomaly is qualitatively explained by our images, which were taken at 10 distinct blade Reynolds numbers spanning a factor of 6. The images show a significant increase in energy related to the oscillations due to transonic vortex shedding at the Reynolds numbers related to the high loss conditions. This increased energy leads to the formation of strong acoustic waves that turn into shock waves at the highest loss conditions. From our observations stemming from Spectral POD analysis of the high-speed images, we see the acoustic waves produced by the trailing edge vortex shedding exist in all conditions tested; but the shedding frequency has a very slight trend upwards as the Reynolds number is increased, varying about 6% in the range tested. Considering this variation of shedding frequency as a function of Reynolds number, which is well-established for other bluff-body flows, we stipulate there may be a potential feedback mechanism involving an acoustic information transfer path across neighboring blades that may explain why only a narrow range of Reynolds numbers displays strong, shock-forming vortex shedding. We consider a few feedback paths and examine the timing based on the mean flow field from a high-resolution LES simulation. It appears that all feedback mechanisms are viable, presenting an integer number of delay cycles with respect to disturbances generated at the trailing edge. Most noteworthy, however, is the acoustic/shock-boundary layer interaction mechanism at the blade geometric throat. Based on our analysis, this mechanism likely plays an important role in realistic, curved turbine blade passages.

Aerodynamics↗

NASA NTRS1978

Thermoelastic effects in frictionally heated line contact of slablike bodies

Many sealing configurations incorporate the geometry of line contact on bluff, slablike bodies. When sliding speed is sufficiently high along or across the line of contact, instabilities may arise from interactions of thermal expansion, elastic deformation and frictional heating. These lead to concentrated contact with elevated temperatures and pressures. Previous studies have been largely restricted to two-dimensional models of such contact. The present study shows how those earlier results must be modified to apply to the more realistic geometry, and show also that the principal features of the interactions are the same in both geometries.

Heckmann, S. R.↗

NASA NTRS1962

HEAT TRANSFER AND PRESSURE DISTRIBUTIONS ON A HEMISPHERECYLINDER AND A BLUFF-AFTERBODY MODEL IN METHANE-AIR COMBUSTION PRODUCTS AND IN AIR

Heat transfer and pressure distributions on a hemisphere-cylinder and a bluff-afterbody model in methane-air combustion products and in air

AERODYNAMIC HEATING↗

NASA NTRS1995

Laser Light Scattering by Shock Waves

Scattering of coherent light as it propagates parallel to a shock wave, formed in front of a bluff cylindrical body placed in a supersonic stream, is studied experimentally and numerically. Two incident optical fields are considered. First, a large diameter collimated beam is allowed to pass through the shock containing flow. The light intensity distribution in the resultant shadowgraph image, measured by a low light CCD camera, shows well-defined fringes upstream and downstream of the shadow cast by the shock. In the second situation, a narrow laser beam is brought to a grazing incidence on the shock and the scattered light, which appears as a diverging sheet from the point of interaction, is visualized and measured on a screen placed normal to the laser path. Experiments are conducted on shocks formed at various free-stream Mach numbers, M, and total pressures, P(sub 0). It is found that the widths of the shock shadows in a shadowgraph image become independent of M and P(sub 0) when plotted against the jump in the refractive index, (Delta)n, created across the shock. The total scattered light measured from the narrow laser beam and shock interaction also follows the same trend. In the numerical part of the study, the shock is assumed to be a 'phase object', which introduces phase difference between the upstream and downstream propagating parts of the light disturbances. For a given shape and (Delta)n of the bow shock the phase and amplitude modulations are first calculated by ray tracing. The wave front is then propagated to the screen using the Fresnet diffraction equation. The calculated intensity distribution, for both of the incident optical fields, shows good agreement with the experimental data.

Panda, J.↗

NASA NTRS1995

Stabilization Of Combustion Of Sprayed Fuel

Several modifications of nozzle spraying liquid propellant into combustion chamber proposed to stabilize combustion. Proposed changes in design of nozzle alter flow field in combustion chamber, according to fluid-mechanical principles, in such way as to suppress oscillations. Various alternative nozzle configurations include bluff or toroidal body to generate turbulence. Other features helping to suppress oscillations include down-stream recirculation zones, baffles, and damping cavities. Similar modifications help to suppress oscillations in industrial combustion chambers and in commercial and domestic oil-burning furnaces.

Voecks, Gerald E.↗

NASA NTRS1991

Local and overall aerodynamic coefficients for bodies in hypersonic, rarefied flow

A computational method is given for the prediction of local pressure and viscous shear stress on windward surfaces of bluff, convex, axisymmetric or quasi-axisymmetric, hypersonic bodies in the transitional, rarefied flow regime. Overall aerodynamic forces and moments are computed by integration of the local quantities. The method is based upon a correlation of local pressure and shear stress computed by the direct simulation Monte Carlo (DSMC) numerical technique for cold wall, real gas conditions and some supplemental data from low-density, hypersonic wind tunnels. The relative simplicity of the method makes it feasible to do the necessary calculations with a personal computer. Two-dimensional shapes and leeward surfaces are not included in the scope of the method as it is presented here. Results are compared with DSMC computations for both local and overall coefficients. The latter includes sphere and blunt cone drag as well as lift and pitching moment coefficients for the NASA AFE vehicle at various angles of attack. Very satisfactory agreement is shown.

Potter, J. Leith↗