scholarly journals Numerical Analysis of High-Altitude Inlet Air on Boundary Layer Flow Loss in an Aero-Engine Compressor

Energies ◽  
2020 ◽  
Vol 13 (16) ◽  
pp. 4155
Author(s):  
Feng Wu ◽  
Limin Gao ◽  
Lu Yang ◽  
Aqiang Lin ◽  
Hai Zhang
Keyword(s):  
Boundary Layer ◽  
Numerical Analysis ◽  
Mach Number ◽  
Pressure Gradient ◽  
High Altitude ◽  
Similarity Criterion ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Entropy Increase ◽  
Wall Boundary

A numerical analysis is performed to explore the high altitude and high Mach flight on the effect of wall boundary layer loss in the compressor. The accuracy for solution results by the application of the similarity criterion and parameter definition of the air inlet is compared with the existing experimental test result. The results indicate that the radial adverse pressure gradient in the rotor domain gradually increases along the span direction and decreases as flight Mach number increases; meanwhile, the circumferential adverse pressure gradient on the pressure side of the rotor blade is correspondingly larger and less than that on the suction side. In particular, the entropy increase along the streamwise shows a decreasing trend and an increasing trend inside the hub and shroud wall boundary layers, respectively. At 2.1 Ma, the entropy increase in the rotor domains enhances by 24.36–27.80% inside the shroud boundary layer, relative to the hub boundary layer; however, it decreases by 0.97–8.54% in the stator domain. With the increase in flight Mach number from 2.1 to 3.4, the average entropy increase reductions in the rotor domain decrease by 18.99–24.97% within the hub boundary layer and 5.71–8.1% within the shroud boundary layer. In the stator domain, it drops by 18.45–9.03% inside the hub boundary layer and 6.88–8.67% inside the shroud boundary layer. It was therefore found that, as Mach number increases from 2.1 to 3.4, the entropy increase reduction is larger inside the hub boundary layer than inside the shroud boundary layer.

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.

Direct Simulations of Wake-Perturbed Separated Boundary Layers

2011 ◽  
Author(s):  
Ayse G. Gungor ◽  
Mark P. Simens ◽  
Javier Jime´nez
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Plate Boundary ◽  
Maximum Velocity ◽  
Adverse Pressure Gradient ◽  
Suction Side ◽  
Single Experiment ◽  
The Stability

A wake-perturbed flat plate boundary layer with a stream-wise pressure distribution similar to those encountered on the suction side of typical low-pressure turbine blades is investigated by direct numerical simulation. The laminar boundary layer separates due to a strong adverse pressure gradient induced by suction along the upper simulation boundary, transitions and reattaches while still subject to the adverse pressure gradient. Various simulations are performed with different wake passing frequencies, corresponding to the Strouhal number 0.0043 < fθb/ΔU < 0.0496 and wake profiles. The wake profile is changed by varying its maximum velocity defect and its symmetry. Results indicate that the separation and reattachment points, as well as the subsequent boundary layer development, are mainly affected by the frequency, but that the wake shape and intensity have little effect. Moreover, the effect of the different frequencies can be predicted from a single experiment in which the separation bubble is allowed to reform after having been reduced by wake perturbations. The stability characteristics of the mean flows resulting from the forcing at different frequencies are evaluated in terms of local linear stability analysis based on the Orr-Sommerfeld equation.

Experimental Investigation of a Turbulent Compressible Boundary Layer

2009 ◽  
Author(s):  
Todd Reedy
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Pressure Gradient ◽  
Static Pressure ◽  
Adverse Pressure Gradient ◽  
Velocity Profiles ◽  
Experimental Results ◽  
Compressible Boundary Layer ◽  
Law Of The Wall ◽  
Displacement Thickness

A turbulent compressible boundary layer in a nominally Mach 4.2 flow was investigated experimentally. Pitot, wall-static pressure, total pressure and temperature measurements were utilized to determine Mach number, temperature, and velocity profiles within the boundary layer. An adverse pressure gradient was observed, resulting in non-uniform flow in the streamwise direction of the test section during development. Alterations were made to the tunnel top and bottom walls to account for the growing boundary layer displacement thickness, resulting in a much improved, uniform Mach number in the freestream and boundary layer. The existence of a slight adverse pressure gradient remained. Flow visualization was conducted via the Schlieren imaging technique. Experimental results were compared against turbulent compressible flow theory and were found to be in excellent agreement, based on an extension of the law-of-the-wall and law-of-the-wake. Velocity profiles and boundary layer thicknesses of the theoretical and experimental results aligned satisfactorily.

Flow Separation Control in an Annular to Conical Diffuser Using Two-Dimensional and Three-Dimensional Wall Steps

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
Kin Pong Lo ◽  
Christopher J. Elkins ◽  
John K. Eaton
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Flow Control ◽  
Separation Bubble ◽  
Adverse Pressure Gradient ◽  
Boundary Layer Separation ◽  
Core Flow ◽  
Layer Separation ◽  
Wall Boundary Layer ◽  
Wall Boundary

Conical diffusers are often installed downstream of a turbomachine with a central hub. Previous studies showed that nonstreamlined hubs had extended separated wakes that reduced the adverse pressure gradient in the diffuser. Active flow control techniques can rapidly close the central separation bubble, but this restores the adverse pressure gradient, which can cause the outer wall boundary layer to separate. The present study focuses on the use of a step-wall diffuser to stabilize the wall boundary layer separation in the presence of core flow control. Three-component mean velocity data for a set of conical diffusers were acquired using magnetic resonance velocimetry. The results showed the step-wall diffuser stabilized the wall boundary layer separation by fixing its location. An axisymmetric step separation bubble was formed. A step with a periodically varying height reduced the reattachment length of the step separation and allowed the diffuser to be shortened. The step-wall diffuser was found to be robust in a range of core flow velocity profiles. The minimum distance between the core flow control mechanism and the step-wall diffuser as well as the minimum length of the step were determined.

scholarly journals Boundary Layer Effects on the Transonic Flow in a Straight Turbine Cascade

1992 ◽  
Author(s):  
M. Štastný ◽  
P. Šafařík
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Boundary Layers ◽  
Transonic Flow ◽  
Side Wall ◽  
Suction Side ◽  
Exit Angle ◽  
Turbine Cascade ◽  
Wall Boundary ◽  
2D Flow

The analysis presented here deals with two subjects related to the transonic flow in a straight turbine cascade: 1) turbulization of the suction side boundary layer and its subsequent development along the profile 2) influence of side wall boundary layers on cascade flows. The turbulization effect of supersonic compression on the boundary layers which accompany transonic expansion on the suction side of the profile is investigated by independent methods. Both the experimental results and the calculation confirm the loss of stability of the laminar boundary layer and its subsequent transition into turbulence in the region of an adverse pressure gradient. Also the possibility of reverse transition of turbulent boundary layer in a subsequent strong favourable pressure gradient is investigated. The contraction effect of the wind tunnel side wall boundary layers is expressed by means of the AVDR factor over a wide range of flow regimes determined by the incidence angle and the Mach number. A correction of the exit flow angle from the cascade is made to obtain a purely 2D flow although it is apparent that the measured exit angle values and the exit angle obtained for a 2D flow differ considerably.

The Effect of Pulsed Blowing on the Boundary Layer of a Highly Loaded Low Pressure Turbine Blade

2013 ◽  
Author(s):  
Marion Mack ◽  
Roland Brachmanski ◽  
Reinhard Niehuis
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Mach Number ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Highly Loaded

The performance of the low pressure turbine (LPT) can vary appreciably, because this component operates under a wide range of Reynolds numbers. At higher Reynolds numbers, mid and aft loaded profiles have the advantage that transition of suction side boundary layer happens further downstream than at front loaded profiles, resulting in lower profile loss. At lower Reynolds numbers, aft loading of the blade can mean that if a suction side separation exists, it may remain open up to the trailing edge. This is especially the case when blade lift is increased via increased pitch to chord ratio. There is a trend in research towards exploring the effect of coupling boundary layer control with highly loaded turbine blades, in order to maximize performance over the full relevant Reynolds number range. In an earlier work, pulsed blowing with fluidic oscillators was shown to be effective in reducing the extent of the separated flow region and to significantly decrease the profile losses caused by separation over a wide range of Reynolds numbers. These experiments were carried out in the High-Speed Cascade Wind Tunnel of the German Federal Armed Forces University Munich, Germany, which allows to capture the effects of pulsed blowing at engine relevant conditions. The assumed control mechanism was the triggering of boundary layer transition by excitation of the Tollmien-Schlichting waves. The current work aims to gain further insight into the effects of pulsed blowing. It investigates the effect of a highly efficient configuration of pulsed blowing at a frequency of 9.5 kHz on the boundary layer at a Reynolds number of 70000 and exit Mach number of 0.6. The boundary layer profiles were measured at five positions between peak Mach number and the trailing edge with hot wire anemometry and pneumatic probes. Experiments were conducted with and without actuation under steady as well as periodically unsteady inflow conditions. The results show the development of the boundary layer and its interaction with incoming wakes. It is shown that pulsed blowing accelerates transition over the separation bubble and drastically reduces the boundary layer thickness.

Boundary Layer Flashback Model for Hydrogen Flames in Confined Geometries Including the Effect of Adverse Pressure Gradient

2020 ◽  
Author(s):  
Ólafur H. Björnsson ◽  
Sikke A. Klein ◽  
Joeri Tober
Keyword(s):  
Boundary Layer ◽  
Pressure Gradient ◽  
Flow Separation ◽  
Gradient Flow ◽  
Lewis Number ◽  
Adverse Pressure Gradient ◽  
Future Research ◽  
Diffuser Flow ◽  
Combustion Properties ◽  
Quenching Distance

Abstract The combustion properties of hydrogen make premixed hydrogen-air flames very prone to boundary layer flashback. This paper describes the improvement and extension of a boundary layer flashback model from Hoferichter [1] for flames confined in burner ducts. The original model did not perform well at higher preheat temperatures and overpredicted the backpressure of the flame at flashback by 4–5x. By simplifying the Lewis number dependent flame speed computation and by applying a generalized version of Stratford’s flow separation criterion [2], the prediction accuracy is improved significantly. The effect of adverse pressure gradient flow on the flashback limits in 2° and 4° diffusers is also captured adequately by coupling the model to flow simulations and taking into account the increased flow separation tendency in diffuser flow. Future research will focus on further experimental validation and direct numerical simulations to gain better insight into the role of the quenching distance and turbulence statistics.

A Superposition Analysis of the Turbulent Boundary Layer in an Adverse Pressure Gradient

1951 ◽  
Vol 18 (1) ◽  
pp. 95-100
Author(s):  
Donald Ross ◽  
J. M. Robertson
Keyword(s):  
Boundary Layer ◽  
Shear Stress ◽  
Wall Shear Stress ◽  
Turbulent Boundary Layer ◽  
Velocity Profile ◽  
Pressure Gradient ◽  
Adverse Pressure Gradient ◽  
Pressure Recovery ◽  
Stress Coefficient ◽  
Wall Shear

Abstract As an interim solution to the problem of the turbulent boundary layer in an adverse pressure gradient, a super-position method of analysis has been developed. In this method, the velocity profile is considered to be the result of two effects: the wall shear stress and the pressure recovery. These are superimposed, yielding an expression for the velocity profiles which approximate measured distributions. The theory also leads to a more reasonable expression for the wall shear-stress coefficient.

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