Turbine Nozzle Endwall Phantom Cooling With Compound Angled Pressure Side Injection

2018 ◽  
Author(s):  
Kevin Liu ◽  
Hongzhou Xu ◽  
Michael Fox
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
Axial Direction ◽  
Secondary Flows ◽  
Test Case ◽  
Pressure Side ◽  
Complex Flow ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Extended Area

Cooling of the turbine nozzle endwall is challenging due to its complex flow field involving strong secondary flows. Increasingly-effective cooling schemes are required to meet the higher turbine inlet temperatures required by today’s gas turbine applications. Therefore, in order to cool the endwall surface near the pressure side of the airfoil and the trailing edge extended area, the spent cooling air from the airfoil film cooling and pressure side discharge slots, referred to as “phantom cooling” is utilized. This paper studies the effect of compound angled pressure side injection on nozzle endwall surface. The measurements were conducted in a high speed linear cascade, which consists of three nozzle vanes and four flow passages. Two nozzle test models with a similar film cooling design were investigated, one with an axial pressure side film cooling row and trailing edge slots; the other with the same cooling features but with compound angled injection, aiming at the test endwall. Phantom cooling effectiveness on the endwall was measured using a Pressure Sensitive Paint (PSP) technique through the mass transfer analogy. Two-dimensional phantom cooling effectiveness distributions on the endwall surface are presented for four MFR (Mass Flow Ratio) values in each test case. Then the phantom cooling effectiveness distributions are pitchwise-averaged along the axial direction and comparisons were made to show the effect of the compound angled injection. The results indicated that the endwall phantom cooling effectiveness increases with the MFR significantly. A compound angle of the pressure side slots also enhanced the endwall phantom cooling significantly. For combined injections, the phantom cooling effectiveness is much higher than the pressure side slots injection only in the endwall downstream extended area.

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.

Effect of Inlet Flow Angle on Gas Turbine Blade Tip Film Cooling

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Intensity Level ◽  
Pressure Side ◽  
Trailing Edge ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The influence of incidence angle on film-cooling effectiveness is studied for a cutback squealer blade tip. Three incidence angles are investigated −0 deg at design condition and ±5 deg at off-design conditions. Based on mass transfer analogy, the film-cooling effectiveness is measured with pressure sensitive paint techniques. The film-cooling effectiveness distribution on the pressure side near tip region, squealer cavity floor, and squealer rim tip is presented for the three incidence angles at varying blowing ratios. The average blowing ratio is controlled to be 0.5, 1.0, 1.5, and 2.0. One row of shaped holes is provided along the pressure side just below the tip; two rows of cylindrical film-cooling holes are arranged on the blade tip in such a way that one row is offset to the suction side profile and the other row is along the camber line. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade. Test is done in a five-blade linear cascade in a blow-down facility with a tip gap clearance of 1.5% of the blade span. The Mach number and turbulence intensity level at the cascade inlet were 0.23 and 9.7%, respectively. It is observed that the incidence angle affects the coolant jet direction on the pressure side near tip region and the blade tip. The film-cooling effectiveness distribution is also altered. The peak of laterally averaged effectiveness is shifted upstream or downstream depending on the off-design incidence angle. The film cooling effectiveness inside the tip cavity can increase by 25% with the positive incidence angle. However, in general, the overall area-averaged film-cooling effectiveness is not significantly changed by the incidence angles in the range of study.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer, and film cooling effectiveness of advanced high pressure turbine blade tips and endwalls. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with leading edge and trailing edge cutouts. Both blade tip configurations have pressure side film cooling and cooling air extraction through dust holes, which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9×105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aerothermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the midchord region. However, on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

3D Velocity and Scalar Field Measurements of an Airfoil Trailing Edge With Slot Film Cooling: The Effect of an Internal Structure in the Slot

2012 ◽  
Author(s):  
Julia Ling ◽  
Sayuri D. Yapa ◽  
Michael J. Benson ◽  
Christopher J. Elkins ◽  
John K. Eaton
Keyword(s):  
Film Cooling ◽  
Field Measurements ◽  
Pressure Side ◽  
Resonance Imaging ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Horseshoe Vortices ◽  
Edge Span ◽  
Slot Film Cooling

Measurements of the 3D velocity and concentration fields were obtained using magnetic resonance imaging for a pressure side cutback film cooling experiment. The cutback geometry consisted of rectangular slots separated by straight lands; inside each of the slots was an airfoil-shaped blockage. The results from this trailing edge configuration, the “island airfoil,” are compared to the results obtained with the “generic airfoil,” a geometry with narrower slots, wider, tapered lands, and no blockages. The objective was to determine how the narrower lands and internal blockages affected the average film cooling effectiveness and the spanwise uniformity. Velocimetry data revealed that strong horseshoe vortices formed around the blockages in the slots, which resulted in greater coolant non-uniformity on the airfoil breakout surface and in the wake. The thinner lands of the island airfoil allowed the coolant to cover a larger fraction of the trailing edge span, giving a much higher spanwise-averaged surface effectiveness, especially near the slot exit where the generic airfoil lands are widest.

Multi-Row Film Cooling Performances of a High Lift Blade and Vane

2012 ◽  
Author(s):  
S. Naik ◽  
J. Krueckels ◽  
M. Gritsch ◽  
M. Schnieder
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Operating Conditions ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
High Lift ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

This paper investigates the aerodynamic and film cooling effectiveness characteristics of a first stage turbine high lift guide vane and its corresponding downstream blade. The vane and blade geometrical profiles and operating conditions are representative of that normally found in a heavy-duty gas turbine. Both the vane and the blade airfoils consist of multi-row film cooling holes located at various axial positions along the airfoil chord. The film cooling holes are geometrically three-dimensional in shape and depending on the location on the airfoil; they can be either symmetrically fan shaped or non-symmetrically fan shaped. Additionally the film cooling holes can be either compounded or in-line with the external flow direction. Numerical studies and experimental investigations in a linear cascade have been conducted at vane and blade exit isentropic Mach number of 0.8. The influence of the coolant flow ejected from the film cooling holes has been investigated for both the vane and the blade profiles. For the nozzle guide vane, the measured film cooling effectiveness compared well with the predictions, especially on the pressure side. The suction side film cooling effectiveness, which consisted of two pre-throat film rows, proved very effective up-to the suction side trailing edge. For the blade, there was a reasonable comparison between the measured and predicted film cooling effectiveness. Again the blade pre-throat fan shaped cooling holes proved very effective up-to the suction side trailing edge. For the vane, the impact of varying the blowing ratios showed a strong variation in the film cooling effectiveness on the pressure side. However, on the blade, the effect of varying the blowing ratio had a greater impact on the suction side film effectiveness compared to the pressure side.

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.

Effect of Inlet Flow Angle on Gas Turbine Blade Tip Film Cooling

2007 ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Intensity Level ◽  
Pressure Side ◽  
Trailing Edge ◽  
Blow Down ◽  
Cooling Effectiveness ◽  
Flow Angle ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The influence of incidence angle on film cooling effectiveness is studied for a cutback squealer blade tip. Three incidence angles are investigated — 0° at design condition and ± 5° at off-design conditions. Based on mass transfer analogy, the film cooling effectiveness is measured with pressure sensitive paint (PSP) techniques. The film cooling effectiveness distribution on pressure side near tip region, squealer cavity floor and squealer rim tip is presented for the three incidence angles at varying blowing ratios. The average blowing ratio is controlled to be 0.5, 1.0, 1.5 and 2.0. One row of shaped holes are provide along the pressure side just below the tip; two rows of cylindrical film cooling holes are arranged on the blade tip in such a way that one row offset to the suction side profile and the other row is along the camber line. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade. Test is done in a five-blade linear cascade in a blow down facility with a tip gap clearance of 1.5% of the blade span. The Mach number and turbulence intensity level at the cascade inlet was 0.23 and 9.7%, respectively. It is observed that the incidence angle affects the coolant jet direction on the pressure side near tip region and blade tip. The film cooling effectiveness distribution is altered. The peak of laterally averaged effectiveness is shifted to upstream or downstream depending on the off-design incidence angles. The film cooling effective inside the tip cavity can be increased up to 25% with the positive incidence angle. However, in general, the overall area averaged film cooling effectiveness is not significantly changed by the incidence angles in the range of study.

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.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2010 ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer and film cooling effectiveness of advanced high-pressure turbine blade tips and endwall. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with a leading edge and trailing edge cut-out. Both blade tip configurations have pressure side film cooling, and cooling air extraction through dust holes which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9 × 105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aero-thermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although, the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the mid-chord region. However on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

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.

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