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At least 1,225 records · Page 68
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Analysis of sound propagation in ducts using the wave envelope concept

A finite difference formulation is presented for sound propagation in a rectangular two-dimensional duct without steady flow for plane wave input. Before the difference equations are formulated, the governing Helmholtz equation is first transformed to a form whose solution does not oscillate along the length of the duct. This transformation reduces the required number of grid points by an order of magnitude, and the number of grid points becomes independent of the sound frequency. Physically, the transformed pressure represents the amplitude of the conventional sound wave. Example solutions are presented for sound propagation in a one-dimensional straight hard-wall duct and in a two-dimensional straight soft-wall duct without steady flow. The numerical solutions show evidence of the existence along the duct wall of a developing acoustic pressure diffusion boundary layer which is similar in nature to the conventional viscous flow boundary layer. In order to better illustrate this concept, the wave equation and boundary conditions are written such that the frequency no longer appears explicitly in them. The frequency effects in duct propagation can be visualized solely as an expansion and stretching of the suppressor duct.

Baumeister, K. J.↗
NASA NTRS2003

Effect of Sub-Boundary Layer Vortex Generations on Incident Turbulence

Sub-boundary layer vortex generators were tested in a wind tunnel to assess their effect on the velocity field within the wake region of a turbulent boundary layer. Both mean flow quantities and turbulence statistics were measured. Although very small relative to the boundary layer thickness, these so-called micro vortex generators were found to have a measurable effect on the power spectra and integral length scales of the turbulence at a distance many times the height of the devices themselves. In addition, the potential acoustic impact of these devices is also discussed. Measured turbulence spectra are used as input to an acoustic formulation in a manner that compares predicted sound pressure levels that result from the incident boundary-layer turbulence, with and without the vortex generators in the flow.

Casper, J.↗
NASA NTRS1984

F-111 natural laminar flow glove flight test data analysis and boundary layer stability analysis

An analysis of 34 selected flight test data cases from a NASA flight program incorporating a natural laminar flow airfoil into partial wing gloves on the F-111 TACT airplane is given. This analysis determined the measured location of transition from laminar to turbulent flow. The report also contains the results of a boundary layer stability analysis of 25 of the selected cases in which the crossflow (C-F) and Tollmien-Schlichting (T-S) disturbance amplification factors are correlated with the measured transition location. The chord Reynolds numbers for these cases ranges from about 23 million to 29 million, the Mach numbers ranged from 0.80 to 0.85, and the glove leading-edge sweep angles ranged from 9 deg to 25 deg. Results indicate that the maximum extent of laminar flow varies from 56% chord to 9-deg sweep on the upper surface, and from 51% chord at 16-deg sweep to 6% chord at 25-deg sweep on the lower. The results of the boundary layer stability analysis indicate that when both C-F and T-S disturbances are amplified, an interaction takes place which reduces the maximum amplification factor of either type of disturbance that can be tolerated without causing transition.

Runyan, L. J.↗
NASA NTRS2012

Numerical Study of Pressure Fluctuations due to High-Speed Turbulent Boundary Layers

Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by fully developed turbulence in supersonic turbulent boundary layers with an emphasis on both pressure fluctuations at the wall and the acoustic fluctuations radiated into the freestream. The wall and freestream pressure fields are first analyzed for a zero pressure gradient boundary layer with Mach 2.5 and Reynolds number based on momentum thickness of approximately 2835. The single and multi-point statistics reported include the wall pressure fluctuation intensities, frequency spectra, space-time correlations, and convection velocities. Single and multi-point statistics of surface pressure fluctuations show good agreement with measured data and previously published simulations of turbulent boundary layers under similar flow conditions. Spectral analysis shows that the acoustic fluctuations outside the boundary layer region have much lower energy content within the high-frequency region. The space-time correlations reflect the convective nature of the pressure field both at the wall and in the freestream, which is characterized by the downstream propagation of pressure-carrying eddies. Relative to those at the wall, the pressure-carrying eddies associated with the freestream signal are larger and convect at a significantly lower speed. The preliminary DNS results of a Mach 6 boundary layer show that the pressure rms in the freestream region is significantly higher than that of the lower Mach number case.

Duan, Lian↗
NASA NTRS1991

Solar rocket plume/mirror interactions

The extent to which the plume from a solar thermal rocket will impinge on the solar collector is studied by flow field analysis. Such interaction can adversely affect collector performance through fouling, excessive heat loading, or pressure loads that deform the delicate structures. The geometrical shape of the collector is such that only the flow from the nozzle boundary layer can reach it, but the thrust levels of interest lead to very viscous nozzle flows with thick boundary layers. Reasonable accuracy in solving these flows requires a fully coupled viscous-inviscid procedure. Results show that the fraction of the plume that hits the collector can be well estimated by continuum theory, but that transitional and rarefied phenomena will have some impact on how it is distributed over the surface. Initial results for one representative condition show that approx. 4 percent of the total flow in the jet makes its way to the collector. The pressures on the collector, however, remain quite low because of its distance from the engine. Additional work is needed to document the effect of thrust scaling and wall cooling on impingement.

Yu, Sheng-Tao↗
NASA NTRS1967

Boundary Shock Waves

Gas flow at high speed out of solid or liquid surface accompanied by heat transfer investigated for boundary shock wave occurrence

SHOCK WAVE↗
NASA NTRS1969

Aerodynamic forces of fluttering cylindrical and/or planar structures

Complexity of the phenomena of panel flutter instability has resulted in the necessity of developing separate design criteria for a variety of flow conditions and panel configurations. Vehicle panel configurations with low aspect ratios are of interest in low supersonic flow, where boundary layer effects are important.

Yates, J. E.↗