Influence of the upstream slot geometry on the endwall cooling and phantom cooling of vane suction side surface

2017 ◽  
Vol 121 ◽  
pp. 688-700 ◽  
Author(s):  
Kun Du ◽  
Zhigang Li ◽  
Jun Li ◽  
Bengt Sunden
Keyword(s):  
Side Surface ◽  
Suction Side ◽  
Endwall Cooling

Numerical Investigation on the Effect of Vortex Generator on Axial Compressor Performance

2014 ◽  
Author(s):  
Ruchika Agarwal ◽  
Anand Dhamarla ◽  
Sridharan R. Narayanan ◽  
Shraman N. Goswami ◽  
Balamurugan Srinivasan
Keyword(s):  
Cross Flow ◽  
Axial Compressor ◽  
Vortex Generator ◽  
Side Surface ◽  
Suction Side ◽  
Test Case ◽  
Stall Margin ◽  
Layer Flow ◽  
Flow Effects ◽  
End Wall

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. The present work is focused on the studying the effects of Vortex Generator (VG) on NASA Rotor 37 test case using Computational Fluid Dynamics (CFD). VG helps in controlling the inception of the stall by generating vortices and energizes the low momentum boundary layer flow which enhances the rotor performance. Three design configuration namely, Counter-rotating, Co-rotating and Plow configuration VG are selected based on the improved aerodynamic performance discussed in reference [1]. These VG are located at 90% span and 42% chord on suction side surface of the blade. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 5.4% from baseline. The reasons for this numerical stall margin improvement are discussed in detail.

Predictions of the Effect of Roughness on Heat Transfer From Turbine Airfoils

1998 ◽  
Author(s):  
Anil K. Tolpadi ◽  
Michael E. Crawford
Keyword(s):  
Heat Transfer ◽  
Side Surface ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Surface Finishing ◽  
Transfer Coefficients ◽  
Average Roughness ◽  
Wide Range ◽  
Turbine Airfoils ◽  
Good Agreement

The heat transfer and aerodynamic performance of turbine airfoils are greatly influenced by the gas side surface finish. In order to operate at higher efficiencies and to have reduced cooling requirements, airfoil designs require better surface finishing processes to create smoother surfaces. In this paper, three different cast airfoils were analyzed: the first airfoil was grit blasted and codep coated, the second airfoil was tumbled and aluminide coated, and the third airfoil was polished further. Each of these airfoils had different levels of roughness. The TEXSTAN boundary layer code was used to make predictions of the heat transfer along both the pressure and suction sides of all three airfoils. These predictions have been compared to corresponding heat transfer data reported earlier by Abuaf et al. (1997). The data were obtained over a wide range of Reynolds numbers simulating typical aircraft engine conditions. A three-parameter full-cone based roughness model was implemented in TEXSTAN and used for the predictions. The three parameters were the centerline average roughness, the cone height and the cone-to-cone pitch. The heat transfer coefficient predictions indicated good agreement with the data over most Reynolds numbers and for all airfoils-both pressure and suction sides. The transition location on the pressure side was well predicted for all airfoils; on the suction side, transition was well predicted at the higher Reynolds numbers but was computed to be somewhat early at the lower Reynolds numbers. Also, at lower Reynolds numbers, the heat transfer coefficients were not in very good agreement with the data on the suction side.

scholarly journals High-Fidelity Simulations of a Linear HPT Vane Cascade Subject to Varying Inlet Turbulence

2017 ◽  
Author(s):  
Richard Pichler ◽  
Richard D. Sandberg ◽  
Gregory Laskowski ◽  
Vittorio Michelassi
Keyword(s):  
Heat Flux ◽  
Turbulence Intensity ◽  
Side Surface ◽  
Suction Side ◽  
Transition Process ◽  
Length Scale ◽  
Length Scales ◽  
Inflow Turbulence ◽  
High Turbulence ◽  
Turbulence Length

The effect of inflow turbulence intensity and turbulence length scales have been studied for a linear high-pressure turbine vane cascade at Reis = 590,000 and Mis = 0.93, using highly resolved compressible large-eddy simulations employing the WALE turbulence model. The turbulence intensity was varied between 6% and 20% while values of the turbulence length scales were prescribed between 5% and 20% of axial chord. The analysis focused on characterizing the inlet turbulence and quantifying the effect of the inlet turbulence variations on the vane boundary layers, in particular on the heat flux to the blade. The transition location on the suction side of the vane was found to be highly sensitive to both turbulence intensity and length scale, with the case with turbulence intensity 20% and 20% length scale showing by far the earliest onset of transition and much higher levels of heat flux over the entire vane. It was also found that the transition process was highly intermittent and local, with spanwise parts of the suction side surface of the vane remaining laminar all the way to the trailing edge even for high turbulence intensity cases.

The Effects of Vane Showerhead Injection Angle and Film Compound Angle on Nozzle Endwall Cooling (Phantom Cooling)

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Nitrogen Gas ◽  
Injection Angle ◽  
Test Models ◽  
Cooling Hole ◽  
Cooling Effects ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Endwall Cooling

The effects of airfoil showerhead (SH) injection angle and film-cooling hole compound angle on nozzle endwall cooling (second order film-cooling effects, also called "phantom cooling") were experimentally investigated in a scaled linear cascade. The test cascade was built based on a typical industrial gas turbine nozzle vane. Endwall surface phantom cooling film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Two separate nozzle test models were fabricated, which have the same number and size of film-cooling holes but different configurations. One had a SH angle of 45 deg and no compound angles on the pressure and suction side (SS) film holes. The other had a 30 deg SH angle and 30 deg compound angles on the pressure and SS film-cooling holes. Nitrogen gas (cooling air) was fed through nozzle vanes, and measurements were conducted on the endwall surface between the two airfoils where no direct film cooling was applied. Six cooling mass flow ratios (MFRs, blowing ratios) were studied, and local (phantom) film effectiveness distributions were measured. Film effectiveness distributions were pitchwise averaged for comparison. Phantom cooling on the endwall by the SS film injections was found to be insignificant, but phantom cooling on the endwall by the pressure side (PS) airfoil film injections noticeably helped the endwall cooling (phantom cooling) and was a strong function of the MFR. It was concluded that reducing the SH angle and introducing a compound angle on the PS injections would enhance the endwall surface phantom cooling, particularly for a higher MFR.

Aerodynamic Blade Row Interactions in an Axial Compressor—Part II: Unsteady Profile Pressure Distribution and Blade Forces

2004 ◽  
Vol 126 (1) ◽  
pp. 45-51 ◽  
Author(s):  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Boundary Layer ◽  
Pressure Distribution ◽  
Pressure Sensors ◽  
Stability Limit ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Experimental Investigations ◽  
Unsteady Pressure ◽  
Blade Row

This two-part paper presents experimental investigations of unsteady aerodynamic blade row interactions in the first stage of the four-stage low-speed research compressor of Dresden. Both the unsteady boundary layer development and the unsteady pressure distribution of the stator blades are investigated for several operating points. The measurements were carried out on pressure side and suction side at midspan. In Part II of the paper the investigations of the unsteady pressure distribution on the stator blades are presented. The experiments were carried out using piezoresistive miniature pressure sensors, which are embedded into the pressure and suction side surface of a single blade. The unsteady pressure distribution on the blade is analyzed for the design point and an operating point near the stability limit. The investigations show that it is strongly influenced by both the incoming wakes and the potential flow field of the downstream rotor blade row. If a disturbance arrives the leading edge or the trailing edge of the blade the pressure changes nearly simultaneously along the blade chord. Thus the unsteady profile pressure distribution is independent of the wake propagation within the blade passage. A phase shift of the reaction on pressure and suction side is observed. The unsteady response of the boundary layer and the profile pressure distribution is compared. Based on the unsteady pressure distribution the unsteady pressure forces of the blades are calculated and discussed.

The Effects of Vane Showerhead Injection Angle and Film Compound Angle on Nozzle Endwall Cooling (Phantom Cooling)

2014 ◽  
Author(s):  
Luzeng Zhang ◽  
Juan Yin ◽  
Hee Koo Moon
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Nitrogen Gas ◽  
Injection Angle ◽  
Test Models ◽  
Cooling Effects ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Endwall Cooling

The effects of airfoil showerhead injection angle and film cooling hole compound angle on nozzle endwall cooling (second order film cooling effects, also called “phantom cooling”) was experimentally investigated in a scaled linear cascade. The test cascade was built based on a typical industrial gas turbine nozzle vane. Endwall surface phantom cooling film effectiveness measurements were made using a computerized pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness can be obtained by the mass transfer analogy. Two separate nozzle test models were fabricated, which have the same number and size of film cooling holes but different configurations. One had a showerhead angle of 45° and no compound angles on the pressure and suction side film holes. The other had a 30° showerhead angle and 30° compound angles on the pressure and suction side film cooling holes. Nitrogen gas (cooling air) was fed through nozzle vanes, and measurements were conducted on the endwall surface between the two airfoils where no direct film cooling was applied. Six cooling mass flow ratios (MFRs, blowing ratios) were studied, and local (phantom) film effectiveness distributions were measured. Film effectiveness distributions were pitchwise averaged for comparison. Phantom cooling on the endwall by the suction side film injections was found to be insignificant, but the pressure side airfoil film injections noticeably helped the endwall cooling (phantom cooling) and was a strong function of the MFR. It was concluded that reducing the showerhead angle and introducing a compound angle on the pressure side injections would enhance the endwall surface phantom cooling, particularly for a higher MFR.

Effect of Wake Disturbance Frequency on the Secondary Flow Vortex Structure in a Turbine Blade Cascade

2000 ◽  
Vol 122 (3) ◽  
pp. 606-613 ◽  
Author(s):  
Christopher G. Murawski ◽  
Kambiz Vafai
Keyword(s):  
Flow Visualization ◽  
Secondary Flow ◽  
Turbine Blade ◽  
Vortex Structure ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Disturbance Frequency ◽  
Flow Frequency ◽  
Flow Vortex

An experimental study of the effect of wake disturbance frequency on the secondary flow vortices in a two-dimensional linear cascade is presented. The flow Reynolds numbers, based on exit velocity and suction side surface length were 25,000, 50,000 and 85,000. Secondary flow was visualized by injecting smoke into the boundary layer and illuminating it with a laser light sheet located at the exit of the cascade. To simulate wakes from upstream blade rows, a set of spanwise cylinders were traversed across the front of the blade row. The flow visualization results with a single wake disturbance reveal that the recovery time of the secondary flow vortex structure decreases as the wake traverse velocity is increased. The results of flow visualization with multiple wakes showed that wake disturbance frequencies below the axial chord flow frequency allowed complete recovery of the secondary flow vortex structure before the next wake encounters the blade leading edge. Wake disturbance frequencies that exceeded the axial chord flow frequency resulted in no observable recovery of the secondary flow vortex structure. Axial chord flow frequency is defined as the axial velocity in the cascade divided by the axial chord length of the turbine blade. [S0098-2202(00)02203-3]

Sign in / Sign up

Export Citation Format

Share Document