On slot injection into a supersonic laminar boundary layer

1973 ◽  
Vol 332 (1588) ◽  
pp. 1-22 ◽  
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Boundary Conditions ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Extreme Conditions ◽  
Numerical Work ◽  
Plate Length ◽  
Injection Region ◽  
Supersonic Laminar

The structure of a supersonic laminar boundary layer near a flat plate is examined when fluid is injected into it, normal to the surface, with velocity O (∊ 3 U * ∞ ) over a distance O (∊ 3 L). Here U * ∞ is the undisturbed velocity of the fluid, L is the plate length and ∊ -8 is a representative Reynolds number of the flow. It is found that the pressure rises upstream of the injection region and that in all the cases fully computed the blowing takes place in a favourable pressure gradient. Afterwards the pressure rises to its undisturbed value. Further incomplete studies suggest that in more extreme conditions, e. g. longer slots, the pressure gradient can be adverse just downstream of the start of the blow and that separation can even occur there. The analytical discussion rests heavily on the notion of the triple-deck, a subdivision of the boundary layer suitable for investigating its response to sudden changes in boundary conditions. Extensive numerical work is also required and the methods devised fail when the boundary layer separates at the onset of the blow. The relation between this type of injection and weak plate injection where the blowing velocity is O (∊ 4 U * ∞ ) and extends over a distance O ( L ) is also considered.

Stability and transition of a supersonic laminar boundary layer on an insulated flat plate

1960 ◽  
Vol 9 (2) ◽  
pp. 257-299 ◽  
Author(s):  
John Laufer ◽  
Thomas Vrebalovich
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Flat Plate ◽  
Laminar Boundary Layer ◽  
Amplification Coefficient ◽  
Unit Distance ◽  
The Stability ◽  
Stability Limits ◽  
Supersonic Laminar

Self-excited oscillations have been discovered experimentally in a supersonic laminar boundary layer along a flat plate. By the use of appropriate measuring techniques, the damping and amplification of the oscillations are studied and the stability limits determined at free-stream Mach numbers 1·6 and 2·2. The wave-like nature of the oscillations is demonstrated and their wave velocities are measured using a specially designed ‘disturbance generator’. It is shown empirically that the stability limits expressed in terms of the boundary-layer-thickness Reynolds number are independent of the Mach number and dependent only on the oscillation frequency. The main effect of compressibility is an increase in wave velocity with Mach number. This has the consequence that the disturbances, although possessing the same dimensionless amplification coefficient as in the incompressible case, have less time (per unit distance) to grow in amplitude. Thus, the adiabatic compressible boundary layer is shown to be more stable than the incompressible one. In general, the experiments confirm the basic assumptions and predictions of the existing stability theory and also suggest the desirability of improvement in the theory in certain phases of the problem. Finally, on the basis of these results a rough estimate of the transition Reynolds number is made in the compressible flow range.

Pressure gradient effects on the large-scale structure of turbulent boundary layers

2013 ◽  
Vol 715 ◽  
pp. 477-498 ◽  
Author(s):  
Zambri Harun ◽  
Jason P. Monty ◽  
Romain Mathis ◽  
Ivan Marusic
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Pressure Gradient ◽  
Boundary Layers ◽  
Amplitude Modulation ◽  
Large Scale ◽  
Outer Region ◽  
Adverse Pressure Gradient ◽  
Turbulent Boundary Layers ◽  
Small Scale

AbstractResearch into high-Reynolds-number turbulent boundary layers in recent years has brought about a renewed interest in the larger-scale structures. It is now known that these structures emerge more prominently in the outer region not only due to increased Reynolds number (Metzger & Klewicki, Phys. Fluids, vol. 13(3), 2001, pp. 692–701; Hutchins & Marusic, J. Fluid Mech., vol. 579, 2007, pp. 1–28), but also when a boundary layer is exposed to an adverse pressure gradient (Bradshaw, J. Fluid Mech., vol. 29, 1967, pp. 625–645; Lee & Sung, J. Fluid Mech., vol. 639, 2009, pp. 101–131). The latter case has not received as much attention in the literature. As such, this work investigates the modification of the large-scale features of boundary layers subjected to zero, adverse and favourable pressure gradients. It is first shown that the mean velocities, turbulence intensities and turbulence production are significantly different in the outer region across the three cases. Spectral and scale decomposition analyses confirm that the large scales are more energized throughout the entire adverse pressure gradient boundary layer, especially in the outer region. Although more energetic, there is a similar spectral distribution of energy in the wake region, implying the geometrical structure of the outer layer remains universal in all cases. Comparisons are also made of the amplitude modulation of small scales by the large-scale motions for the three pressure gradient cases. The wall-normal location of the zero-crossing of small-scale amplitude modulation is found to increase with increasing pressure gradient, yet this location continues to coincide with the large-scale energetic peak wall-normal location (as has been observed in zero pressure gradient boundary layers). The amplitude modulation effect is found to increase as pressure gradient is increased from favourable to adverse.

Turbine Cascade Flow Control Using a “Wake Filling” Pulsed Plasma Actuator

2005 ◽  
Author(s):  
H. Perez-Blanco ◽  
Robert Van Dyken ◽  
Aaron Byerley ◽  
Tom McLaughlin
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Laminar Boundary Layer ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Plasma Actuator ◽  
Pulsed Plasma ◽  
Power Cost ◽  
Hot Film

Separation bubbles in high-camber blades under part-load conditions have been addressed via continuous and pulsed jets, and also via plasma actuators. Numerous passive techniques have been employed as well. In this type of blades, the laminar boundary layer cannot overcome the adverse pressure gradient arising along the suction side, resulting on a separation bubble. When separation is abated, a common explanation is that kinetic energy added to the laminar boundary layer speeds up its transition to turbulent. In the present study, a plasma actuator installed in the trailing edge (i.e. “wake filling configuration”) of a cascade blade is used to excite the flow in pulsed and continuous ways. The pulsed excitation can be directed to the frequencies of the large coherent structures (LCS) of the flow, as obtained via a hot-film anemometer, or to much higher frequencies present in the suction-side boundary layer, as given in the literature. It is found that pulsed frequencies much higher than that of LCS reduce losses and improve turning angles further than frequencies close to those of LCS. With the plasma actuator 50% on time, good loss abatement is obtained. Larger “on time” values yield improvements, but with decreasing returns. Continuous high-frequency activation results in the largest loss reduction, at increased power cost. The effectiveness of high frequencies may be due to separation abatement via boundary layer excitation into transition, or may simply be due to the creation of a favorable pressure gradient that averts separation as the actuator ejects fluid downstream. Both possibilities are discussed in light of the experimental evidence.

The Critical Height of Distributed Roughness to Cause Boundary Layer Transition

1959 ◽  
Vol 63 (588) ◽  
pp. 722-722
Author(s):  
R. L. Dommett
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Laminar Boundary Layer ◽  
Boundary Layer Transition ◽  
Critical Height ◽  
Layer Transition ◽  
Local Conditions ◽  
Additional Advantage ◽  
Boundary Layer Equations ◽  
General Method

It has been found that there is a critical height for “sandpaper” type roughness below which no measurable disturbances are introduced into a laminar boundary layer and above which transition is initiated at the roughness. Braslow and Knox have proposed a method of predicting this height, for flow over a flat plate or a cone, using exact solutions of the laminar boundary layer equations combined with a correlation of experimental results in terms of a Reynolds number based on roughness height, k, and local conditions at the top of the elements. A simpler, yet more general, method can be constructed by taking additional advantage of the linearity of the velocity profile near the wall in a laminar boundary layer.

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