Effects of the cooling configurations layout near the first-stage vane leading edge on the endwall cooling and phantom cooling of the vane suction side surface

2018 ◽  
Vol 123 ◽  
pp. 1021-1034
Author(s):  
Kun Du ◽  
Jun Li ◽  
Bengt Sunden
Keyword(s):  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Endwall Cooling

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.

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]

Effects of Reynolds Number and Surface Roughness Magnitude and Location on Compressor Cascade Performance

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Seung Chul Back ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Surface Roughness ◽  
Reynolds Number ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Reynolds Numbers ◽  
Pressure Side ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Loading

An experimental investigation has been conducted to characterize the influence of Reynolds number and surface roughness magnitude and location on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and loss have been measured via a five-hole probe, pitot probe, and pressure taps on the blades. Four different roughness magnitudes, Ra values of 0.38 μm (polished), 1.70 μm (baseline), 2.03 μm (rough 1), and 2.89 μm (rough 2), have been tested. Furthermore, various roughness locations have been examined. In addition to the as manufactured (baseline) and entirely rough blade cases, blades with roughness covering the leading edge, pressure side, and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been carried out for Reynolds numbers ranging from 300,000 to 640,000. For Reynolds numbers under 500,000, the tested roughnesses do not significantly degrade compressor blade loading or loss. However, loss and blade loading become sensitive to roughness at Reynolds numbers above 550,000. Cascade performance is more sensitive to roughness on the suction side than pressure side. Furthermore, roughness on the aft 2/3 of suction side surface has a greater influence on loss. For a given roughness location, there exists a Reynolds number at which loss begins to significantly increase. Finally, increasing the roughness area on the suction surface from the leading edge reduces the Reynolds number at which the loss begins to increase.

Heat Transfer Measurements in a Rotating Equilateral Triangular Channel With Various Rib Arrangements

2006 ◽  
Author(s):  
Dong Hyun Lee ◽  
Dong-Ho Rhee ◽  
Hyung Hee Cho ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Stationary Case ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Ribbed Channel ◽  
Triangular Channel ◽  
Rotating Channel

The present research investigates the heat transfer characteristics in an equilateral triangular channel to simulate the leading edge cooling passage of a gas turbine blade. The experiments are conducted for the stationary and rotating ribbed channel with three different attack angles (45°, 90° and 135°). Square ribs are installed in a staggered manner on the pressure and suction side surfaces of the channel. The rib height to channel hydraulic diameter ratio (e/Dh) is 0.079 and the rib-to-rib pitch (p) is 8 times of the rib height. To measure regional-averaged heat transfer coefficients in the channel, two rows of copper blocks with heaters are installed on each surface. The rotation number ranges from 0.0 to 0.1 for the fixed Reynolds number of 10,000. Inlet coolant-to-surface density ratio is about 0.2. For the channel with 90° ribs, the heat transfer rates of all regions have similar values for stationary case. However, for the rotating channel, heat transfer coefficients on the pressure side surface are significantly increased while the suction side surface has quite low heat transfer coefficients due to a single rotating secondary flow induced by Coriolis force. For the channel with angled rib arrangements, a pair of counter-rotating vortices is induced by the angled rib arrangements. High heat transfer coefficients are obtained on the regions near the inner wall for 45° angled ribbed channel and near the leading edge for the 135° angled ribbed channel. The heat transfer coefficients in rotating channel with angled ribs are almost the same as those of stationary case for the tested conditions because the secondary flow dominates the heat transfer. The channel with angled ribs consistently yields better thermal performance than the transverse ribbed channel for the test conditions of the present study.

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

2003 ◽  
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 analysed 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.

scholarly journals Investigation of Separation Phenomena in a Radial Pump at Reduced Flow Rate by Large-Eddy Simulation

2016 ◽  
Vol 138 (12) ◽  
Author(s):  
Antonio Posa ◽  
Antonio Lippolis ◽  
Elias Balaras
Keyword(s):  
Reverse Flow ◽  
Leading Edge ◽  
Suction Side ◽  
Secondary Flows ◽  
Vaned Diffuser ◽  
Flow Rates ◽  
Turbulent Statistics ◽  
Large Eddy ◽  
Radial Pump ◽  
Reduced Flow

Turbopumps operating at reduced flow rates experience significant separation and backflow phenomena. Although Reynolds-Averaged Navier–Stokes (RANS) approaches proved to be usually able to capture the main flow features at design working conditions, previous numerical studies in the literature verified that eddy-resolving techniques are required in order to simulate the strong secondary flows generated at reduced loads. Here, highly resolved large-eddy simulations (LES) of a radial pump with a vaned diffuser are reported. The results are compared to particle image velocimetry (PIV) experiments in the literature. The main focus of the present work is to investigate the separation and backflow phenomena occurring at reduced flow rates. Our results indicate that the effect of these phenomena extends up to the impeller inflow: they involve the outer radii of the impeller vanes, influencing significantly the turbulent statistics of the flow. Also in the diffuser vanes, a strong spanwise evolution of the flow has been observed at the reduced load, with reverse flow, located mainly on the shroud side and on the suction side (SS) of the stationary channels, especially near the leading edge of the diffuser blades.

Numerical simulation of swirling flow effect on the first stage vane film cooling distribution

2016 ◽  
Vol 07 (03) ◽  
pp. 1650031
Author(s):  
Hong Yin
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Swirling Flow ◽  
Cooling System ◽  
Leading Edge ◽  
Suction Side ◽  
Local Area ◽  
Swirl Flow ◽  
Flow Effect ◽  
Recirculation Flow

In advanced gas turbine technology, lean premixed combustion is an effective strategy to reduce peak temperature and thus, NO[Formula: see text] emissions. The swirler is adopted to establish recirculation flow zone, enhancing mixing and stabilizing the flame. Therefore, the swirling flow is dominant in the combustor flow field and has impact on the vane. This paper mainly investigates the swirling flow effect on the turbine first stage vane cooling system by conducting a group of numerical simulations. Firstly, the numerical methods of turbulence modeling using RANS and LES are compared. The computational model of one single swirl flow field is considered. Both the RANS and LES results give reasonable recirculation zone shape. When comparing the velocity distribution, the RANS results generally match the experimental data but fail to at some local area. The LES modeling gives better results and more detailed unsteady flow field. In the second step, the RANS modeling is incorporated to investigate the vane film cooling performance under the swirling inflow boundary condition. According to the numerical results, the leading edge film cooling is largely altered by the swirling flow, especially for the swirl core-leading edge aligned case. Compared to the pressure side, the suction side film cooling is more sensitive to the swirling flow. Locally, the film cooling jet is lifted and turned by the strong swirling flow.

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