Improvement of Cooled Turbine Airfoils by Special Cutback

2010 ◽  
Author(s):  
Boris I. Mamaev ◽  
Mikhail M. Petukhovskiy ◽  
Alexander V. Pozdnyakov ◽  
Marat R. Valeev
Keyword(s):  
Substantial Reduction ◽  
Reynolds Numbers ◽  
Pressure Side ◽  
Trailing Edge ◽  
Turbine Efficiency ◽  
Flow Parameters ◽  
Pressure Distributions ◽  
Turbine Airfoils ◽  
Bend Angles ◽  
Cooling Air

A substantial reduction in high temperature turbine efficiency due to a thickening trailing edge on the blades can be compensated by ejection of cooling air on the airfoil pressure side near the edge, which is made thinner at the expense of a pressure-side contour bend. A blade-row midspan section of the aircraft high-pressure turbine was chosen for investigations. Flow parameters of the section: inlet and outlet angles were 36° and 65°, respectively (axial reference), outlet isentropic Mach number was 0.94. Four linear cascades were examined. They differed mainly in the airfoil trailing edge geometry. Three airfoils had the same thin trailing edges and contour bend angles ε = 10, 15 and 20°; one airfoil with a thick round edge had no bend. Widths of the slot for cooling air ejection were the same for all airfoils tested. Measurements were made for exit Mach numbers from 0.6 to 0.95 and relative cooling mass flows from 0 to 1.5%. The respective Reynolds numbers varied from 7.5·106 to 9·106. The incidence value was 2°. Pressure distributions along profiles, outlet total and static pressures, back pressures for cooling air with gas-outlet angles were measured. The experiments showed streamlining of all cascades were favorable. For the airfoils with ε = 10 and 15° the profile losses were low and normal for uncooled cascades with thin trailing edge. Hence, for such bends losses due to a step on the airfoil pressure side were negligible. As expected, the losses in the cascade with the thick rounding edge were significantly higher. The losses in the cascade with ε = 20° were the greatest. The coolant exit had no distinct influence on streamlining airfoils. The back-pressure for cooling air was approximately equal to the outlet static pressure. For cascades with ε = 10 and 15° the ejection of coolant led to a small increase of losses due to additional mixing losses. Thus, the airfoil contour bend is a powerful tool for the aerodynamic improvement of cooled turbines. It may lead to gains in stage efficiency of 1…1.5%. It should be noted that this tool has already been used successfully for several aircraft and industrial turbines of recent design.

Heat Transfer and Pressure Drop in a Converging Pedestal Array With Exit Area Variation

2018 ◽  
Author(s):  
L. W. Soma ◽  
F. E. Ames ◽  
S. Acharya
Keyword(s):  
Heat Transfer ◽  
Pressure Drop ◽  
Flow Rate ◽  
Film Cooling ◽  
Reynolds Numbers ◽  
Flow Path ◽  
Internal Cooling ◽  
Channel Measurements ◽  
Trailing Edge ◽  
Cooling Air

The trailing edge of a vane is one of the most difficult areas to cool due to a narrowing flow path, high external heat transfer rates, and deteriorating external film cooling protection. Converging pedestal arrays are often used as a means to provide internal cooling in this region. The thermally induced stresses in the trailing edge region of these converging arrays have been known to cause failure in the pedestals of conventional solidity arrays. The present paper documents the heat transfer and pressure drop through two high solidity converging rounded diamond pedestal arrays. These arrays have a 45 percent pedestal solidity. One array which was tested has nine rows of pedestals with an exit area in the last row consistent with the convergence. The other array has eight rows with an expanded exit in the last row to enable a higher cooling air flow rate. The expanded exit of the eight row array allows a 30% increase in the coolant flow rate compared with the nine row array for the same pressure drop. Heat transfer levels correlate well based on local Reynolds numbers but fall slightly below non converging arrays. The pressure drop across the array naturally increases toward the trailing edge with the convergence of the flow passage. A portion of the cooling air pressure drop can be attributed to acceleration while a portion can be attributed to flow path losses. Detailed array static pressure measurements provide a means to develop a correlation for the prediction of pressure drop across the cooling channel. Measurements have been acquired over Reynolds numbers based on exit flow conditions and the characteristic pedestal length scale ranging from 5000 to over 70,000.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

Experiments With Three-Dimensional Passive Flow Control Devices on Low-Pressure Turbine Airfoils

2004 ◽  
Vol 128 (2) ◽  
pp. 251-260 ◽  
Author(s):  
Douglas G. Bohl ◽  
Ralph J. Volino
Keyword(s):  
Reynolds Number ◽  
Flow Control ◽  
Separation Bubble ◽  
Three Dimensional ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Turbine Airfoils

The effectiveness of three-dimensional passive devices for flow control on low pressure turbine airfoils was investigated experimentally. A row of small cylinders was placed at the pressure minimum on the suction side of a typical airfoil. Cases with Reynolds numbers ranging from 25,000 to 300,000 (based on suction surface length and exit velocity) were considered under low freestream turbulence conditions. Streamwise pressure profiles and velocity profiles near the trailing edge were documented. Without flow control a separation bubble was present, and at the lower Reynolds numbers the bubble did not close. Cylinders with two different heights and a wide range of spanwise spacings were considered. Reattachment moved upstream as the cylinder height was increased or the spacing was decreased. If the spanwise spacing was sufficiently small, the flow at the trailing edge was essentially uniform across the span. The cylinder size and spacing could be optimized to minimize losses at a given Reynolds number, but cylinders optimized for low Reynolds number conditions caused increased losses at high Reynolds numbers. The effectiveness of two-dimensional bars had been studied previously under the same flow conditions. The cylinders were not as effective for maintaining low losses over a range of Reynolds numbers as the bars.

Detached Eddy Simulation of Film Cooling Performance on the Trailing Edge Cut-Back of Gas Turbine Airfoils

2005 ◽  
Author(s):  
P. Martini ◽  
A. Schulz ◽  
H.-J. Bauer ◽  
C. F. Whitney
Keyword(s):  
Film Cooling ◽  
Back Surface ◽  
Pressure Side ◽  
Mixing Process ◽  
Transfer Coefficients ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Turbine Airfoils

The present study deals with the unsteady flow simulation of trailing edge film cooling on the pressure side cut-back of gas turbine airfoils. Before being ejected tangentially on the inclined cut-back surface, the coolant air passes a partly converging passage that is equipped with turbulators such as pin fins and ribs. The film mixing process on the cut-back is complicated. In the near slot region, due to the turbulators and the blunt pressure side lip, turbulence is expected to be anisotropic. Furthermore, unsteady flow phenomena like vortex shedding from the pressure side lip might influence the mixing process (i.e. the film cooling effectiveness on the cut-back surface). In the current study, three different internal cooling designs are numerically investigated starting from the steady RaNS solution, and ending with unsteady detached eddy simulations (DES). Blowing ratios M = 0.5; 0.8; 1.1 are considered. To obtain both, film cooling effectiveness as well as heat transfer coefficients on the cut-back surface, the simulations are performed using adiabatic and diabatic wall boundary conditions. The DES simulations give a detailed insight into the unsteady film mixing process on the trailing edge cut-back, which is indeed influenced by vortex shedding from the pressure side lip. Furthermore, the time averaged DES results show very good agreement with the experimental data in terms of film cooling effectiveness and heat transfer coefficients.

Evaluation of RANS Predictions on a Linear Nozzle Vane Cascade With Trailing Edge Cutback Film Cooling

2013 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Vane ◽  
Pressure Surface ◽  
Exit Mach Number ◽  
Cooling Air

The present study concentrates on the numerical investigation of a cooled trailing edge in a linear nozzle vane cascade typical of a high-pressure turbine. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is also ejected through two rows of cooling holes placed on the pressure side, upstream of the cutback. The main goal is to evaluate the reliability of RANS predictions in such a complex cooling system. Different coolant-to-mainstream mass flow ratio values up to MFR = 2.8% were simulated at exit Mach number of M2is = 0.2. The computed performance of the trailing edge cooling scheme was compared to available measurements of: holes and cutback exit velocity and discharge behavior; boundary layer along traverses located on the pressure side, downstream of each row of cooling holes and approaching the trailing edge; adiabatic film cooling effectiveness. Special emphasis was dedicated to coolant-mainstream interaction and film cooling effectiveness over the pressure surface of the vane. Despite the steady approach, the simulations provided a reliable overview of coolant and mainstream aerodynamic features. The limitations in predicting the measured drop in cooling effectiveness toward the trailing edge were highlighted as well.

Experiments With Three Dimensional Passive Flow Control Devices on Low-Pressure Turbine Airfoils

2005 ◽  
Author(s):  
Douglas G. Bohl ◽  
Ralph J. Volino
Keyword(s):  
Reynolds Number ◽  
Flow Control ◽  
Separation Bubble ◽  
Three Dimensional ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Trailing Edge ◽  
Low Pressure Turbine ◽  
Wide Range ◽  
Turbine Airfoils

The effectiveness of three dimensional passive devices for flow control on low pressure turbine airfoils was investigated experimentally. A row of small cylinders was placed at the pressure minimum on the suction side of a typical airfoil. Cases with Reynolds numbers ranging from 25,000 to 300,000 (based on suction surface length and exit velocity) were considered under low freestream turbulence conditions. Streamwise pressure profiles and velocity profiles near the trailing edge were documented. Without flow control a separation bubble was present, and at the lower Reynolds numbers the bubble did not close. Cylinders with two different heights and a wide range of spanwise spacings were considered. Reattachment moved upstream as the cylinder height was increased or the spacing was decreased. If the spanwise spacing was sufficiently small, the flow at the trailing edge was essentially uniform across the span. The cylinder size and spacing could be optimized to minimize losses at a given Reynolds number, but cylinders optimized for low Reynolds number conditions caused increased losses at high Reynolds numbers. The effectiveness of two-dimensional bars had been studied previously under the same flow conditions. The cylinders were not as effective for maintaining low losses over a range of Reynolds numbers as the bars.

Investigation of the Suction Side Boundary Layer Development on Low Pressure Turbine Airfoils With and Without Separation Using a Preston Probe

2014 ◽  
Author(s):  
Stephan Stotz ◽  
Christian T. Wakelam ◽  
Reinhard Niehuis ◽  
Yavuz Guendogdu
Keyword(s):  
Boundary Layer ◽  
Dynamic Pressure ◽  
Suction Side ◽  
Transition Process ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Low Pressure Turbine ◽  
Pressure Distributions ◽  
Turbine Airfoils ◽  
Probe Geometry

Characterizing the transition process of airfoils can be very challenging and requires often extensive measurement methods. Frequently at low Reynolds numbers the suction side separation often occurs close to the trailing edge so that asserting reattachment of the flow to form a closed separation bubble from the profile pressure distributions becomes uncertain. In the current work the suction side transition process is investigated more precisely with a convenient method to determine the dynamic pressure close to the suction surface using a Preston probe (flattened Pitot tube). Therefore four low pressure turbine airfoils, which show different characteristics of the transition process in the static pressure distribution have been investigated at the High-Speed Cascade Wind Tunnel at the Universität der Bundeswehr München at constant Mach number and under a wide range of Reynolds numbers (40 000 to 400 000). It is shown that this method is appropriate to determine transition start and end as well as the separation and reattachment point of a separated flow as long as the probe height is small enough compared to the boundary layer thickness. The measurement results are compared to profile pressure distributions and hot-wire boundary layer profiles. Also the influence of periodic unsteady inflow conditions on the dynamic pressure near the wall is revealed in the time average. Limitations due to the probe geometry are discussed and a method to estimate the influence of the probe geometry on the measured dynamic pressure coefficient is suggested.

Experimental and numerical investigations of hole injection on the suction side throat of transonic turbine vanes in a cascade with trailing edge injection

2017 ◽  
Vol 232 (8) ◽  
pp. 1454-1466 ◽  
Author(s):  
Jie Gao ◽  
Ming Wei ◽  
Yunning Liu ◽  
Qun Zheng ◽  
Ping Dong
Keyword(s):  
Shock Wave ◽  
Mass Flow ◽  
Suction Side ◽  
Hole Injection ◽  
Pressure Side ◽  
Trailing Edge ◽  
Transonic Flows ◽  
Pressure Distributions ◽  
Numerical Predictions ◽  
Mach Numbers

Trailing-edge mixing flows associated with coolant injection are complex, in particular at transonic flows, and result in significant aerodynamics losses. The objective of this paper is to evaluate the impacts of hole injection near the suction side throat on shock wave control and aerodynamic losses. A series of tests and calculations on effects of hole injection on the suction-side throat of a high-pressure turbine vane cascade with and without trailing-edge injection were conducted. Wake traverses with a five-hole probe and tests of pressure distributions on the turbine profile were taken for total injection mass flow ratios of 0% and 1.2% under test Mach numbers of 0.7, 0.78, and 0.87. Meantime, numerical predictions are carried out for exit isentropic Mach numbers of 0.7, 0.78, 0.87, and 1.1 and hole-injection mass flow ratios of 0%, 0.17%, 0.3%, and 0.89%. Numerical predictions show a reasonable agreement with the experimental data, and wake total pressure losses and flow angles as well as pressure distributions on the turbine profile were compared to calculations without hole injection, indicating a significant effect of hole injection on the profile wake development and its blockage effect on the shock-wave flow in the vane cascade passage. At subsonic flows, the hole injection on the suction side throat thickens the suction-side boundary layer, and increases the flow mixing, thus causing increased wake losses and flow angles. At transonic flows, while the trailing-edge injection reduces the strength of the shock wave at the trailing-edge pressure side, the hole injection on the suction side throat alters the local pressure fields, and then tends to enhance the shock-wave at the trailing-edge pressure-side; however, it seems to reduce the strength of the shock-wave at the trailing-edge suction side.

Film Cooling Effectiveness and Heat Transfer on the Trailing Edge Cut-Back of Gas Turbine Airfoils With Various Internal Cooling Designs

2005 ◽  
Author(s):  
P. Martini ◽  
A. Schulz ◽  
H.-J. Bauer
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Back Surface ◽  
Internal Cooling ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Airfoils

The present study deals with trailing edge film cooling on the pressure side cut-back of gas turbine airfoils. Before being ejected tangentially onto the inclined cut-back surface the coolant air passes a partly converging passage that is equipped with turbulators such as pin fins and ribs. The experiments are conducted in a generic set-up and cover a broad variety of internal cooling designs. A subsonic atmospheric open-loop wind tunnel is utilized for the tests. The test conditions are characterized by a constant Reynolds number of Rehg = 250,000, a turbulence intensity of Tuhg = 7%, and a hot gas temperature of Thg = 500K. Due to the ambient temperature of the coolant, engine realistic density ratios between coolant and gas can be realized. Blowing ratios cover a range of 0.20<M<1.25. The experimental data to be presented include discharge coefficients, adiabatic film cooling effectiveness and heat transfer coefficients in the near slot region (x/H<15). The results clearly demonstrate the strong influence of the internal cooling design and the relatively thick pressure side lip (t/H = 1) on film cooling performance downstream of the ejection slot.

Analysis of the Flow Around Supercavitating Hydrofoils with Midchord and Face Cavity Detachment

1991 ◽  
Vol 35 (03) ◽  
pp. 198-209
Author(s):  
Spyros A. Kinnas ◽  
Neal E. Finer
Keyword(s):  
Pressure Distribution ◽  
Integral Equations ◽  
Inviscid Flow ◽  
Panel Method ◽  
Pressure Side ◽  
Flow Conditions ◽  
Trailing Edge ◽  
Pressure Distributions ◽  
Work First ◽  
Flow Criterion

In this work, first the linearized supercavitating hydrofoil problem with arbitrary cavity detachment points is formulated in terms of unknown source and vorticity distributions. The corresponding integral equations are inverted analytically and the results are expressed in terms of integrals of quantities which depend only on the hydrofoil shape. These integrals are computed numerically, in an accurate and efficient way, to produce cavity shapes and pressure distributions on the foil and cavity. The effect of the cavity detachment points on the shape of the cavity and the foil pressure distribution is investigated. An inviscid flow criterion for the cavity detachment point is derived for the case where the cavity detaches in front of the trailing edge on the pressure side of the hydrofoil. Finally, the accuracy of the linearized cavity theory is assessed for different foils and flow conditions, by analyzing the produced cavity shapes with a nonlinear panel method.

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