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At least 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Boundary-Layer-Ingesting Inlet Flow Control

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP(sub avg)) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

Owens, Lewis R.↗
NASA NTRS1984

On the stability of the decelerating laminar boundary layer

The stability of a decelerating boundary-layer flow is investigated experimentally and numerically. Experimentally, a flat plate having a Blasius boundary layer is decelerated in an 18 m towing tank. The boundary layer becomes unstable to two-dimensional waves, which break down into three-dimensional patterns, hairpin vortices, and finally turbulent bursts when the vortices lift off the wall. The unsteady boundary-layer equations are solved numerically to generate instantaneous velocity profiles for a range of boundary and initial conditions. A quasi-steady approximation is invoked and the stability of local velocity profiles is determined by solving the Orr-Sommerfeld equation using Chebyshev matrix methods. Comparisons are made between the numerical predictions and the experimentally observed instabilities.

Gad-El-hak, M.↗
NASA NTRS1984

Heat and mass transfer in combustion - Fundamental concepts and analytical techniques

Fundamental combustion phenomena and the associated flame structures in laminar gaseous flows are discussed on physical bases within the framework of the three nondimensional parameters of interest to heat and mass transfer in chemically-reacting flows, namely the Damkoehler number, the Lewis number, and the Arrhenius number which is the ratio of the reaction activation energy to the characteristic thermal energy. The model problems selected for illustration are droplet combustion, boundary layer combustion, and the propagation, flammability, and stability of premixed flames. Fundamental concepts discussed include the flame structures for large activation energy reactions, S-curve interpretation of the ignition and extinctin states, reaction-induced local-similarity and non-similarity in boundary layer flows, the origin and removal of the cold boundary difficulty in modeling flame propagation, and effects of flame stretch and preferential diffusion on flame extinction and stability. Analytical techniques introduced include the Shvab-Zeldovich formulation, the local Shvab-Zeldovich formulation, flame-sheet approximation and the associated jump formulation, and large activation energy matched asymptotic analysis. Potentially promising research areas are suggested.

Law, C. K.↗
NASA NTRS1986

Turbulent Reynolds analogy factors of stacked large-eddy breakup devices

Direct measurements are made of turbulent Reynolds analogy factors, referenced to a flat plate, for turbulent boundary layer flows altered by stacked arrays of large eddy breakup devices (LEBUs). These are of interest as drag reducers when inserted into a boundary layer transverse to the flow. The data thus obtained furnish evidence that heat transfer, skin friction drag, and LEBU performance factors in low Reynolds number flows are sensitive to flow history. Attention is given to the apparatus and measurement procedures used.

Lindemann, A. M.↗
NASA NTRS2013

A Comparative Study of Four Impedance Eduction Methodologies Using Several Test Liners

A comparative study of four commonly used impedance eduction methods is presented for a range of liner structures and test conditions. Two of the methods are restricted to uniform flow while the other two accommodate both uniform and boundary layer flows. Measurements on five liner structures (a rigid-wall insert, a ceramic tubular liner, a wire mesh liner, a low porosity conventional liner, and a high porosity conventional liner) are obtained using the NASA Langley Grazing Flow Impedance Tube. The educed impedance of each liner is presented for forty-two test conditions (three Mach numbers and fourteen frequencies). In addition, the effects of moving the acoustic source from upstream to downstream and the refractive effects of the mean boundary layer on the wire mesh liner are investigated. The primary conclusions of the study are that: (1) more accurate results are obtained for the upstream source, (2) the uniform flow methods produce nearly identical impedance spectra at and below Mach 0.3 but significant scatter in the educed impedance occurs at the higher Mach number, (3) there is better agreement in educed impedance among the methods for the conventional liners than for the rigid-wall insert, ceramic, or wire mesh liner, and (4) the refractive effects of the mean boundary layer on the educed impedance of the wire mesh liner are generally small except at Mach 0.5.

Watson, Willie R.↗
NASA NTRS1980

Effect of pressure gradients on Gortler instability

Gortler instability for boundary-layer flows over generally curved walls is considered. The full linearized disturbance equations are obtained in an orthogonal curvilinear coordinate system. A perturbation procedure to account for second-order effects is used to determine the effects of the displacement thickness and the variation of the streamline curvature on the neutral stability of the Blasius flow. The pressure gradient in the mean flow is accounted for by solving the nonsimilar boundary-layer equations. Growth rates are obtained for the actual mean flow and the Falkner-Skan flows. The results demonstrate the strong influence of the pressure gradient and the nonsimilarity of the basic flow on the stability characteristics.

Ragab, S. A.↗
NASA NTRS1981

Goertler instability

Goertler instability for boundary-layer flows over generally curved walls is considered. The full-linearized disturbance equations are obtained in an orthogonal curvilinear coordinate system. A perturbation procedure to account for second-order effects is used to determine the effects of the displacement thickness and the variation of the streamline curvature on the neutral stability of the Blasius flow. The streamwise pressure gradient in the mean flow is accounted for by solving the nonsimilar boundary-layer equations. Growth rates are obtained for the actual mean flow and compared with those for the Blasius flow and the Falkner-Skan flows. The results demonstrate the strong influence of the streamwise pressure gradient and the nonsimilarity of the basic flow on the stability characteristics.

Ragab, S. A.↗
NASA NTRS1996

Experimental and Computational Boundary-Layer Studies in a Supersonic Two-Dimensional Nozzle

Calculations have been carried out on the adiabatic laminar boundary layer developing on the surface of a two-dimensional supersonic nozzle, consisting of contoured nozzle blocks and flat sidewalls. Two- and three-dimensional Navier-Stokes codes, as well as two-dimensional boundary-layer codes have been employed. Thee codes have been adapted to the characteristics of a specific wind tunnel nozzle, so that their numerical results could be directly compared with experimental data obtained in the same nozzle. Such comparisons have been made for the boundary-layer growth on the contoured nozzle, and for the boundary-layer growth, surface streamlines and surface shear on the sidewalls. The three-dimensional Navier-Stokes code was found to be the only one to correctly predict the mean boundary-layer flow on both the sidewalls and the contoured nozzle. Theory and experiment both indicate that the sidewall flow is highly three-dimensional, with non-uniform shear, comer vortices and a boundary layer strongly distorted by cross flows induced by lateral pressure gradients.

Brogan, Torence P.↗
NASA NTRS1998

Experimental and Computational Boundary-Layer Studies In a Supersonic Two-Dimensional Nozzle

Experiments and computations were carried out on the adiabatic laminar boundary layer developing along the surfaces of a two-dimensional supersonic nozzle, consisting of upper and lower contoured nozzle blocks and flat sidewalls. Two- and three-dimensional Navier-Stokes codes, as well as two-dimensional boundary-layer codes were employed. These codes were adapted to the characteristics of a specific wind tunnel nozzle, so that their numerical results could be directly compared with experimental data obtained in the same nozzle. Such comparisons were made for the boundary-layer growth on the nozzle contoured surfaces, and for the boundary-layer growth, surface streamlines and surface shear on the sidewalls. The three-dimensional Navier-Stokes code was found to be the only one to correctly predict the mean laminar boundary-layer flow on both the sidewalls and the contoured surfaces. Theory and experiment both indicated that the sidewall flow is highly three-dimensional, with non-uniform shear, corner vortices and a boundary layer strongly distorted by cross flows induced by lateral pressure gradients.

Demetriades, Anthony↗
NASA NTRS2021

Hypersonic Second-Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗
NASA NTRS2021

Hypersonic Second-Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗
NASA NTRS1989

Computation of three dimensional turbulent boundary layers in internal flows, including turbomachinery rotor blades

A method is developed for predicting three-dimensional turbulent boundary layers occurring in internal flows, including rotor blades of turbomachinery. These boundary layers are complex, turbulent, and are subject to Coriolis and centrifugal forces. The algebraic Reynolds stress model (ARSM) developed in this paper satisfies the realizability conditions exactly and captures the changes in turbulenet structure arising from curvature and rotation. The prediction of pressure driven secondary flow agrees well with the data and all the three turbulent models (k-e, algebraic eddy viscosity, and ARSM) show the same level of agreement. The prediction of boundary layer on rotor blades shows much better agreement with the ARSM. It is essential to employ the higher order models to capture the effects of rotation and curvature and three-dimensional boundary layers in turbomachinery.

Zhang, J.↗
NASA NTRS1987

Shear-layer structures in near-wall turbulence

The structure of internal shear layer observed in the near-wall region of turbulent flows is investigated by analyzing flow fields obtained from numerical simulations of channel and boundary-layer flows. It is found that the shear layer is an important contributor to the turbulence production. The conditionally averaged production at the center of the structure was almost twice as large as the long-time mean value. The shear-layer structure is also found to retain its coherence over streamwise distances on the order of a thousand viscous length units, and propagates with a constant velocity of about 10.6 u sub rho throughout the near wall region.

Johansson, A. V.↗
NASA NTRS1981

Nonparallel stability of three-dimensional flows

The linear stability of three-dimensional incompressible, isothermal, nonparallel boundary-layer flows has been investigated. The method of multiple scales is used to derive the partial-differential equations that describe the spatial modulations of the amplitude, phase and wavenumber of a disturbance. Group velocities are used to determine the disturbance growth direction. The envelope method is used to calculate the logarithmic amplitude growth rate N. The theory is applied to the flows over a swept-back tapered wing with boundary-layer suction. Results of such analysis for the X-21 wing are discussed. It is found that the nonparallel effects for this wing is substantial.

Padhye, A. R.↗
NASA NTRS1983

Effects of interstage diffuser flow distortion on the performance of a 15.41-centimeter tip diameter axial power turbine

The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.

Mclallin, K. L.↗
NASA NTRS1983

Effects of interstage diffuser flow distortion on the performance of a 15.41-centimeter tip diameter axial power turbine

The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.

Mclallin, K. L.↗
NASA NTRS1983

Growth of Goertler vortices in compressible boundary layers along curved surfaces

An investigation of the growth of the three-dimensional, counter-rotating, longitudinal type vortices is considered in two-dimensional laminar compressible boundary-layer flow. The basic approximation of the disturbance equations that includes the terms due to boundary layer growth is considered and solved numerically. These terms are shown to have large local effects near the neutral stability region. The study shows that the instability of the boundary layer with respect to the three-dimensional vortices sets in at higher Goertler number as Mach number increases. Also the maximum amplitude ratio of the vortices is reduced by about 20 percent as Mach number increases from 0 to 5.

El-Hady, N. M.↗