Numerical simulation of circulation control turbine cascade with Coanda jet and counter-flow blowing at high Mach numbers

2017 ◽  
Vol 121 (1243) ◽  
pp. 1239-1260 ◽  
Author(s):  
Y. Feng ◽  
Y. Song ◽  
F. Chen
Keyword(s):  
Guide Vane ◽  
Inlet Guide Vane ◽  
Circulation Control ◽  
Turbine Cascade ◽  
Flow Conditions ◽  
Suction Surface ◽  
Counter Flow ◽  
Flow Angle ◽  
High Mach ◽  
Coanda Jet

ABSTRACTThe performance of a circulation-control inlet guide vane that makes use of the Coanda effect was studied numerically in a high Mach number turbine cascade. The effect of different shapes (elliptic and circular) of the Coanda surface at the blade trailing edge was investigated by implementing both a Coanda jet and a counter-flow blowing. Under high subsonic flow conditions, with a total blowing ratio of 3% of the mainstream, the circulation control cascade can reach the same performance as the reference stator with a 13.5% reduction in the axial chord length, with minimal increase of the energy loss coefficient. The Coanda surfaces with small curvature are more efficient in entraining the mainstream flow, and they achieve better aerodynamic performance. The wall attachment of the Coanda jet is improved by employing counter-flow blowing, resulting in a slight increase of both the exit flow angle and the expansion ratio. Under supersonic flow conditions at the cascade exit, it is more difficult for the circulation control cascade to reach the appropriate flow turning due to a premature shock wave, which is absent in the original cascade until the very end of the suction surface.

scholarly journals Flow Measurements Behind the Inlet Guide Vane of a Centrifugal Compressor

1998 ◽  
Author(s):  
I. Kassens ◽  
M. Rautenberg
Keyword(s):  
Centrifugal Compressor ◽  
Total Pressure ◽  
Guide Vane ◽  
Incidence Angle ◽  
Pressure Ratio ◽  
Inlet Guide Vane ◽  
Flow Measurements ◽  
Flow Angle ◽  
Partial Load ◽  
Measurement Location

In a centrifugal compressor adjustable inlet guide vanes (IGV) in front of the impeller are used to regulate the pressure ratio and the mass flow. The stationary measurement of the velocity profile in front of the impeller with different angles of the IGV displays shock losses at the inlet edge of blade of the impeller. In the partial-load region (e.g. partial-load efficiency) the radial distribution of the flow influences considerably the performance of the impeller. The tested compressor consists of an adjustable IGV with straight vanes, a shrouded impeller and a vaneless, parallel diffuser. In the first measurement location, behind the IGV, total pressure, static pressure and flow angle were measured with a 5-hole cylinder probe. In the second measurement location, in front of the impeller, the measurement of the total pressure was carried out with a Kiel probe and the flow angle with a Cobra probe accordingly the static wall pressure was measured. Taking into consideration the fundamental thermodynamical equations it was possible to determine the velocity profiles because of the measured distributions of the flow angle in these two measurement locations. For different angles of the IGV and with various mass flows the distributions of the deflection defect behind the IGV are described. Starting with the measured distributions of the flow in front of the impeller the flow angles at the impeller inlet are calculated and the distributions of the incidence angle at the impeller inlet are figured out.

Design and Test of an Aspirated Counter-Rotating Fan

2006 ◽  
Author(s):  
Jack L. Kerrebrock ◽  
Alan H. Epstein ◽  
Ali A. Merchant ◽  
Gerald R. Guenette ◽  
David Parker ◽  
...  
Keyword(s):  
Guide Vane ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Inlet Guide Vane ◽  
Suction Surface ◽  
Viscous Boundary Layer ◽  
Blade Loading ◽  
Design Speed ◽  
Viscous Boundary ◽  
Shock Loss

The design and test of a two-stage, vaneless, aspirated counter-rotating fan is presented in this paper. The fan nominal design objectives were a pressure ratio of 3:1 and adiabatic efficiency of 87%. A pressure ratio of 2.9 at 89% efficiency was measured in the tests. The configuration consists of a counter-swirl-producing inlet guide vane, followed by a high tip speed (1450 feet/sec) non-aspirated rotor, and a counter-rotating low speed (1150 feet/sec) aspirated rotor. The lower tip speed and lower solidity of the second rotor results in a blade loading above conventional limits, but enables a balance between the shock loss and viscous boundary layer loss, the latter of which can be controlled by aspiration. The aspiration slot on the second rotor suction surface extends from the hub up to 80% span, with a conventional tip clearance, and the bleed flow is discharged at the hub. The fan was tested in a short duration blowdown facility. Particular attention was given to the design of the instrumentation to obtain efficiency measurements within 0.5 percentage points. High response static pressure measurements were taken between the rotors and downstream of the fan to determine the stall behavior. Pressure ratio, mass flow, and efficiency on speedlines from 90% to 102% of the design speed are presented and discussed along with comparison to CFD predictions and design intent. The results presented here complement those presented earlier for two aspirated fan stages with tip shrouds, extending the validated design space for aspirated compressors to include designs with conventional unshrouded rotors and with inward removal of the aspirated flow.

Effects of Axial Row-Spacing for Double-Jet Film-Cooling on the Cooling Effectiveness

2011 ◽  
Author(s):  
Zhan Wang ◽  
Jian-Jun Liu ◽  
Bai-tao An ◽  
Chao Zhang
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Guide Vane ◽  
Row Spacing ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Cylindrical Holes ◽  
Compound Angle

The effects of axial row-spacing for double jet film-cooling (DJFC) with compound angle on the cooling characteristics under different blowing ratios were investigated numerically. First, the flow fields and cooling effectiveness of DJFC on flat plate with different axial row-spacing were calculated. Film-cooling with fan-shaped or cylindrical holes was also calculated for the comparison. The results indicate that a larger axial row-spacing is helpful to form the anti-kidney vortex and to improve the cooling effectiveness. The DJFC was then applied to the suction and pressure surface of a real turbine inlet guide vane. Comparisons of film-cooling effectiveness with the cylindrical and fan-shaped holes were also conducted. The results for the guide vane show that on the suction surface the DJFC with a larger axial row-spacing leads to better film coverage and better film-cooling effectiveness than the cylindrical or fan-shaped holes. On the pressure surface, however, the film-cooling with fan-shaped holes is superior to the others.

The Effects of Inlet Guide Vane-Wake Impingement on the Boundary Layer and the Near-Wake of a Rotor Blade

2010 ◽  
Vol 132 (4) ◽  
Author(s):  
Francesco Soranna ◽  
Yi-Chih Chow ◽  
Oguz Uzol ◽  
Joseph Katz
Keyword(s):  
Boundary Layer ◽  
Rotor Blade ◽  
Guide Vane ◽  
Shear Velocity ◽  
Inlet Guide Vane ◽  
Surface Boundary ◽  
Momentum Thickness ◽  
Suction Surface ◽  
Near Wake ◽  
Trailing Edge

This paper examines the response of a rotor blade boundary layer and a rotor near-wake to an impinging wake of an inlet guide vane (IGV) located upstream of the rotor blade. Two-dimensional particle image velocimetry (PIV) measurements are performed in a refractive index matched turbomachinery facility that provides unobstructed view of the entire flow field. Data obtained at several rotor phases enable us to examine the IGV-wake-induced changes to the structure of the boundary layer and how these changes affect the flow and turbulence within the rotor near-wake. We focus on the suction surface boundary layer, near the blade trailing edge, but analyze the evolution of both the pressure and suction sides of the near-wake. During the IGV-wake impingement, the boundary layer becomes significantly thinner, with lower momentum thickness and more stable profile compared with other phases at the same location. Analysis of available terms in the integral momentum equation indicates that the phase-averaged unsteady term is the main contributor to the decrease in momentum thickness within the impinging wake. Thinning of the boundary/shear layer extends into the rotor near-wake, making it narrower and increasing the phase-averaged shear velocity gradients and associated turbulent kinetic energy (TKE) production rate. Consequently, the TKE increases during wake thinning, with as much as 75% phase-dependent variations in its peak magnitude. This paper introduces a new way of looking at the PIV data by defining a wake-oriented coordinate system, which enables to study the structure of turbulence around the trailing edge in great detail.

The Effect of Reduced Frequency on Transition and Separation at the Leading Edge of a Compressor Stator

2012 ◽  
Author(s):  
Samuel C. T. Perkins ◽  
Alan D. Henderson
Keyword(s):  
Pressure Gradient ◽  
Steady Flow ◽  
Separation Bubble ◽  
Leading Edge ◽  
Flow Conditions ◽  
Suction Surface ◽  
Flow Angle ◽  
Pressure Surface ◽  
Reduced Frequency ◽  
Wake Passing

Studies on the effects of stator reduced frequency in low pressure turbines have shown that periodic wake-induced unsteadiness can increase steady flow circulation by as much as 15% and reduce losses compared to a steady flow datum. A large separation bubble downstream of peak suction that formed under steady flow conditions was periodically suppressed by wake passing events, resulting in significantly reduced losses within the boundary layer. This research extends this concept to a controlled diffusion compressor stator blade with a circular arc leading edge. The blade was placed inside a large scale, two-dimensional, cascade with a rotating bar mechanism used to simulate an upstream rotor blade row. The blade profile has been shown to experience leading edge separations and subsequent transition on both the pressure and suction surfaces due to a velocity overspeed caused by discontinuities in surface curvature. Testing was carried out at reduced frequencies of 0.47, 0.94 and 1.88 at the design inlet flow angle 45.5° and Reynolds number based on chord of 230,000. The freestream turbulence intensity was 4.0%. A range of experimental measurements were used to look at the blade’s performance: high resolution time-averaged blade surface static pressure measurements, inlet and exit 3-hole probe traverses and instantaneous, ensemble averaged and time average surface mounted hot-film measurements for the calculation of turbulent intermittency and quasi wall-shear stress. Results showed that increasing the stator reduced frequency from, 0–1.88, increased the overall blade pressure loss. The losses generated by the pressure surface and suction surface differed significantly and are affected very differently. The pressure surface demonstrated a clear reduction in loss with an increase in reduced frequency whereas the opposite trend was seen on the suction surface. Wake-induced turbulent strips suppressed the formation of leading edge separation bubbles that formed under steady flow conditions and in between wake passing events. Wake-induced turbulent strips reduced in width and level of turbulent intermittency through the favorable pressure gradients leading to peak suction and grew in the adverse pressure gradient of the velocity overspeed. The flow between wake-induced turbulent strips partially relaminarised through the favorable pressure gradient leading to peak suction.

Aeroelasticity at Reversed Flow Conditions — Part II: Application to Compressor Surge

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Harald Schoenenborn ◽  
Thomas Breuer
Keyword(s):  
Guide Vane ◽  
Navier Stokes ◽  
Inlet Guide Vane ◽  
Flow Conditions ◽  
Complex Flow ◽  
Pressure Shock ◽  
Blade Vibration ◽  
Blow Down ◽  
Torsional Mode ◽  
Blade Loading

The prediction of blade loads during surge is still a challenging task. In the literature, the blade loading during surge is often referred to as “surge load,” which suggests that there is a single source of blade loading. In the second part of our paper it is shown that, in reality, the “surge load” may consist of two physically different mechanisms: the pressure shock when the pressure breaks down and aeroelastic excitation (flutter) during the blow-down phase in certain cases. This leads to a new understanding of blade loading during surge. The front block of a multistage compressor is investigated. For some points of the backflow characteristic, the quasi steady-state flow conditions are calculated using a Reynolds averaged Navier-Stokes (RANS)-solver. The flow enters at the last blade row, goes backwards through the compressor and leaves the compressor in front of the inlet guide vane. The results show a very complex flow field characterized by large recirculation regions on the suction sides of the airfoils and stagnation regions close to the trailing edges of the airfoils. Based on these steady solutions, unsteady calculations are performed with a linearized aeroelasticity code. It can be shown that some of the rotor stages are aerodynamically unstable in the first torsional mode. Thus, in addition to the pressure shock, the blades may be excited by flutter during the surge blow-down phase. In spite of the short blow-down phase typical for aero-engine high pressure compressors, this may lead to very high blade stresses due to high aeroelastic excitation at these special flow conditions. The analytical results compare very well with the observations during rig testing. The correct nodal diameter of the blade vibration is reproduced and the growth rate of the blade vibration is predicted quite well, as a comparison with tip-timing measurements shows. A new flutter region in the compressor map was experimentally and analytically detected.

Study of a Surge-Free Centrifugal Compressor With Automatically Variable Inlet and Diffuser Vanes

1996 ◽  
Author(s):  
Hideomi Harada
Keyword(s):  
Centrifugal Compressor ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Low Flow ◽  
Inlet Guide Vane ◽  
Analysis Method ◽  
Unstable Flow ◽  
Flow Angle ◽  
Operating Range

In order to improve the operating range of a centrifugal compressor, computer-controlled variable inlet and diffuser vanes were attached to a compressor with a pressure ratio of 2.5. Low-solidity cascade vanes capable of controlling the vane angle up to 0 degrees from the tangential direction were used for the vaned diffuser. The compressor’s overall performance was then tested using a closed-loop test stand. By automatically adjusting the diffuser vanes to the most suitable flow angle, pressure fluctuations caused by the unstable flow in the diffuser during low-flow operation of the centrifugal compressor could be suppressed, and the compressor could be operated nearly up to the shut-off flow rate without any surge. The author experimentally confirmed the critical operating range of both the impeller and diffuser at two different tip speeds and five inlet guide vane angles. Furthermore, a three-dimensional viscous flow-analysis method was applied to the impeller, and a three-dimensional momentum integral analysis method was applied to the diffuser. Then the critical operating ranges obtained in the experiments were qualitatively validated. The operating range of a centrifugal compressor under low-flow conditions, which has until now been limited because of surge, dramatically improved in this study, thereby demonstrating that it may be possible to develop a surge-free centrifugal compressor.

Design and Test of an Aspirated Counter-Rotating Fan

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Jack L. Kerrebrock ◽  
Alan H. Epstein ◽  
Ali A. Merchant ◽  
Gerald R. Guenette ◽  
David Parker ◽  
...  
Keyword(s):  
Static Pressure ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Inlet Guide Vane ◽  
Suction Surface ◽  
Viscous Boundary Layer ◽  
Blade Loading ◽  
Design Speed ◽  
Viscous Boundary ◽  
Shock Loss

The design and test of a two-stage, vaneless, aspirated counter-rotating fan is presented in this paper. The fan nominal design objectives were a pressure ratio of 3:1 and adiabatic efficiency of 87%. A pressure ratio of 2.9 at 89% efficiency was measured at the design speed. The configuration consists of a counter-swirl-producing inlet guide vane, followed by a high tip speed (1450ft∕s) nonaspirated rotor and a counter-rotating low speed (1150ft∕s) aspirated rotor. The lower tip speed and lower solidity of the second rotor result in a blade loading above conventional limits, but enable a balance between the shock loss and viscous boundary layer loss; the latter of which can be controlled by aspiration. The aspiration slot on the second rotor suction surface extends from the hub up to 80% span. The bleed flow is discharged inward through the blade hub. This fan was tested in a short duration blowdown facility. Particular attention was given to the design of the instrumentation to measure efficiency to 0.5% accuracy. High response static pressure measurements were taken between the rotors and downstream of the fan to determine the stall behavior. Pressure ratio, mass flow, and efficiency on speed lines from 90% to 102% of the design speed are presented and discussed along with comparison to computational fluid dynamics predictions and design intent. The results presented here complement those presented earlier for two aspirated fan stages with tip shrouds, extending the validated design space for aspirated compressors to include designs with conventional unshrouded rotors and with inward removal of the aspirated flow.

Inlet Flow Angle Determination of Transonic Compressor Cascades

1992 ◽  
Vol 114 (3) ◽  
pp. 487-493 ◽  
Author(s):  
W. Steinert ◽  
R. Fuchs ◽  
H. Starken
Keyword(s):  
Test Results ◽  
Flow Conditions ◽  
Suction Surface ◽  
Inlet Flow ◽  
Flow Angle ◽  
Transonic Compressor ◽  
Measuring Techniques ◽  
Special Measuring

Tests of transonic compressor cascades require special measuring techniques to determine the inlet flow angle around sonic inlet flow conditions. One of the main requirements for these methods is the ability to adjust the inlet flow angle during the test to a prescribed value. A method has been successfully applied that relies on theoretically determined suction surface velocities. The described method was applied in testing cascades at inlet Mack numbers between M1 = 0.75−1.18. The test results confirmed the practicability of this method.

scholarly journals Experimental Investigation on Spatial Development of Streamwise Vortices in a Turbine Inlet Guide Vane Cascade

1985 ◽  
Author(s):  
Toyotaka Sonoda
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Spatial Development ◽  
Inlet Guide Vane ◽  
Surface Boundary ◽  
Streamwise Vortices ◽  
Suction Surface ◽  
Turbine Inlet

In order to obtain a better understanding of secondary flow in a turbine cascade, spatial development of a leading-edge horseshoe vortex has been investigated experimentally in a large-scale, low-speed, high-accelerated, plane turbine inlet guide vane cascade. Flow has been visualized by issuing kerosene vapor into the inlet boundary layer and the vane suction surface boundary layer, respectively. Based on many cross-sectional photographs normal to the flow and supplemental measurements of the wall static pressure on the vane and the endwall, the evolution of a leading-edge horseshoe vortex into streamwise vortices and the generation of a new type streamwise vortex pair on the suction surface near the endwall are discussed.

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