A Numerical Investigation of Stator-Rotor Interaction Effects on Flow Field and Film Cooling Effectiveness in a 3D Transonic Turbine Stage With Highly Twisted Rotors

2013 ◽  
Author(s):  
Huazhao Xu ◽  
Jianhua Wang ◽  
Ting Wang
Keyword(s):  
Shock Waves ◽  
Film Cooling ◽  
Static Pressure ◽  
Pressure Fluctuations ◽  
Suction Side ◽  
Turbine Stage ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Air

To reduce aerodynamic losses and optimize turbine blade cooling designs, a comprehensive understanding of rotor-stator interaction effects on the blade aerodynamics and film cooling performance is essential. This paper focuses on the numerical analysis of the interactions between shock waves and unsteady wakes and their effects on cooling effectiveness of a highly twisted rotor within a transonic turbine stage. The parameters of the turbine stage are from the Pratt & Whitney Energy Efficient Engine (E3) program. The Realizable k-ε turbulence model was selected as the suitable turbulence model by our previous study. The investigation is conducted first by analyzing mean static pressure and the Root Mean Square (RMS) of the static pressure, followed by a detailed study of the flow field in the rotor passage at blowing ratios (Br) of 0.5, 1.0 and 1.5. Effects of the complicated interactions among shock waves, trailing edge wake shedding, and blockage of moving rotors are separated and identified individually through shock strength, vortices, and entropy production. The results show that: 1) For the stator, the shock waves emanating from the trailing edge of the neighboring stator impinging on the later part of the stator’s suction side, creating static pressure fluctuations as large as 20%. 2) For the rotor, the variation of static pressure is synchronized with the rotor passing frequency, but out of phase between the suction and pressure sides. 3) A high entropy region generated by the wake flow from the upstream trailing edge in the rotor passage intensifies and moves towards the rotor hub during the rotor passing periods. 4) Most of the cooling air injected from the rotor leading edge bends towards the suction side, and the cooling air injected from the pressure side turns towards the rotor hub. 5) An increase in the blowing ratio from Br = 0.5 to Br = 1.5 does not affect the pressure fluctuations, but does significantly increase film cooling effectiveness on the rotor pressure side. 6) The mean static pressure on the suction side of the twist blade is lower than a straight blade, indicating the benefit of producing larger torque by using twist rotors.

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.

Improved Trench Film Cooling With Shaped Trench Outlets

2011 ◽  
Author(s):  
Habeeb Idowu Oguntade ◽  
Gordon E. Andrews ◽  
Alan Burns ◽  
Derek B. Ingham ◽  
Mohammed Pourkashanian
Keyword(s):  
Mass Flow ◽  
Film Cooling ◽  
Cross Flow ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Vertical Slot ◽  
Superior Surface ◽  
Cooling Air ◽  
Computational Predictions

This paper presents the influence of the shaped trailing edge of trench outlets on film cooling effectiveness and aerodynamics. A 90° outlet wall to a trench will give a vertical slot jet into the cross flow and it was considered that improvements in the cooling effectiveness would occur if the trailing edge of the trench outlet was bevelled or filleted. CFD approach was used for these investigations which started with the predictions of the conventional sharp edged trench outlet for two experimental geometries. The computational predictions for the conventional sharp edged trench outlet were shown to have good agreement with the experimental data for two experimental geometries. The shaped trailing edge of the trench outlet was predicted to improve the film cooling effectiveness. The bevelled and filleted trench outlets were predicted to further suppress vertical jet momentum and give a Coanda effect that allowed the cooling air to attach to the downstream wall surface with a better transverse spread of the coolant film. The new trench outlet geometries would allow a reduction in film cooling mass flow rate for the same cooling effectiveness. Also, it was predicted that reducing the coolant mass flow per hole and increasing the number of holes gave, for the same total coolant mass flow, a much superior surface averaged cooling effectiveness for the same cooled surface area.

The Integrated Gas/Steam Nozzle With Steam Cooling: Part II — Design Considerations

1984 ◽  
Author(s):  
Ivan G. Rice
Keyword(s):  
Shock Wave ◽  
Film Cooling ◽  
Fat Body ◽  
Suction Side ◽  
Body Type ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Wave Development ◽  
Gas Nozzle

The integration of multiple steam nozzles with the first-stage annular-gas nozzle to form a binary-flow system in a reheat-gas turbine is presented whereby steam is first used as an internal vane coolant before being expanded and accelerated for work extraction. Steam nozzles are located in “fat-body” type vanes. Trailing-edge impingement followed by reverse-serpentine-flow cooling takes place. Internal trailing-edge-steam nozzles produce either diffusion or shock-wave boundary-layer disturbance inside the trailing edge to enhance heat transfer. Externally, steam blanketing reduces nozzle-profile loss and improves film cooling effectiveness by reducing the surface viscosity and secondly by controlling suction-side aft-shock-wave development. A new vane shape coupled with a gas-turning-combustor system is suggested to improve vane-film cooling effectiveness further.

Simulation of Air/Mist Cooling Among Shock Waves and Passing Wakes Interactions in a Transonic Gas Turbine Stage

2021 ◽  
Author(s):  
Ting Wang ◽  
Ramy Abdelmaksoud
Keyword(s):  
Shock Waves ◽  
Gas Turbine ◽  
Film Cooling ◽  
Suction Side ◽  
Water Droplets ◽  
Discrete Phase Model ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Enhancement ◽  
Mist Cooling

Abstract This paper presents a 2-D numerical investigation of the effect of interactions of moving wakes and shock waves on mist cooling performance over airfoils in the first stator-rotor stage of a transonic gas turbine. The discrete phase model (DPM) is used to simulate and track the evaporation and movement of the tiny water droplets. Breakup and coalescence sub-models are used to simulate the interaction between the droplets themselves. A linear sliding mesh technique is used to study the transient stator-rotor interaction. The results show that the passing unsteady wakes caused by the blade rotation press the mist on the blade suction side flowing near the blade surface, providing more enhanced film cooling effectiveness. The weak oblique shock waves do not exert a significant effect on the air/mist cooling effectiveness. Injecting a 10% mist ratio noticeably improved the cooling enhancement by reducing the wall temperature values up to 200 K in some locations. Injecting the tiny water droplets does not cause a noticeable pressure loss compared to the air-only cooling case. Injecting mist doesn’t alter the effect of shocks.

Tripod Hole Geometry Performance for a Vane Suction Surface Near Throat Location

2013 ◽  
Author(s):  
Sridharan Ramesh ◽  
Chris LeBlanc ◽  
Srinath V. Ekkad ◽  
Mary Anne Alvin
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Throat Region ◽  
Blowing Ratio ◽  
Hole Geometry ◽  
Film Cooling Holes ◽  
Cooling Air

Film cooling performance depends strongly on the hole exit geometry, blowing ratio, and hole location. The goal of this study is to evaluate film cooling geometries that can provide better protection over the suction surface of the airfoil beyond the throat region. This study compares the performance of standard cylindrical; fan-shaped (10° flare/laidback); tripod hole geometry (15° breakout angle); and tripod holes with shaped exits (5° flare on 15° tripod). Film cooling holes are located just upstream of the throat region on the suction side of an airfoil. The airfoil is a scaled up first stage vane from GE E3 engine and is mounted on a low speed linear cascade wind tunnel. A range of blowing ratios from 0.5 to 2.0 was covered for a cylindrical hole, while ensuring all other hole geometries run under similar mass flow rate conditions. Steady state IR (Infra-Red) technique was employed to measure adiabatic film cooling effectiveness. Results show that the tripod holes with and without shaped exits provide much higher film effectiveness than cylindrical and slightly higher effectiveness than shaped exit holes using 50% lesser cooling air while operating at the same blowing ratios. Effectiveness values up to 0.2–0.25 are seen 40-hole diameters downstream for the tripod hole configurations thus providing cooling in the important trailing edge portion of the airfoil.

Numerical Investigations of Wake and Shock Wave Effects on Film Cooling Performance in a Transonic Turbine Stage: Part 1 — Methodology Development and Qualification Over Stationary Stators and Rotors

2013 ◽  
Author(s):  
Huazhao Xu ◽  
Jianhua Wang ◽  
Ting Wang
Keyword(s):  
Experimental Data ◽  
Mach Number ◽  
Film Cooling ◽  
Suction Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Numerical Investigations ◽  
Blowing Ratio

To understand the unsteady shock wave and wake effects on the film cooling performance over a transonic 3-D rotating stage, a series of numerical investigations have been conducted and are presented in this two-part paper. Part 1 is focused on the development of the computational model and methodology of the system setup and model qualification; Part 2 is to investigate the unsteady effects of shock waves and wakes on film cooling performance in a transonic rotating stage. In Part 1, the film cooling experimental conditions (non-rotating) and test sections of Kopper et. al. and Hunter are selected for model qualification. The numerical computation is carried out by the commercial software Ansys/Fluent using the pressure based compressible flow governing equations. The effects of four turbulence models are carefully compared with the experimental data. The Realizable k-ε turbulence model is found to match the experimental data better than the other models and is thus used for the rest of the study, including Part 2. The results show that 1) the weak shock emanating from the neighboring stator’s trailing edge results in a temperature rise and a reduction of film cooling effectiveness on the suction side near the trailing edge, 2) cooling ejection from the trailing edge reduces the shock strength in the stator passage, 3) an increase in Mach number from 0.84 to 1.50 can reduce the total pressure losses of fluid flow near the end-walls, 4) the film cooling effectiveness increases with increasing blowing ratio and becomes more even on the stator with a higher blowing ratio, and 5) an increase in Mach number from 0.84 to 1.50 gives rise to a higher cooling effectiveness in the region from the cooling holes to 80% of the chord length of the stator on the pressure side, but becomes lower after this up to the trailing edge. However, on the stator’s suction side, higher Mach number results in a lower cooling effectiveness region around the film holes from 30% to 55% of the chord length, but cooling effectiveness increases downstream.

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.

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.

Experimental and Numerical Investigation of Trailing Edge Film Cooling by Circular Coolant Wall Jets Ejected From a Slot With Internal Rib Arrays

2003 ◽  
Author(s):  
P. Martini ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Film Cooling ◽  
Three Dimensional ◽  
Numerical Data ◽  
Trailing Edge ◽  
Adiabatic Wall ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Edge Features ◽  
Cooling Air

The present study concentrates on the experimental and computational investigation of a cooled trailing edge in a modern turbine blade. The trailing edge features a pressure side cutback and a slot, stiffened by two rows of evenly spaced ribs in an inline configuration. Cooling air is ejected through the slot and forms a cooling film on the trailing edge cutback region. In the present configuration the lateral spacing of the ribs equals two times their width. The height of the ribs, i.e. the height of the slot equals their width. Since the ribs are provided with fillet radii of half the slot height in size, circular coolant jets are exiting the slot tangentially to the trailing edge cutback. The adiabatic wall temperature mappings on the trailing edge cutback indicate that strong three-dimensional flow interaction between the coolant jets and the hot main flow takes place in such a way that two or more coolant jets coalesce depending on the blowing ratio. Experimental and numerical data to be presented in the present study include adiabatic film cooling effectiveness on the trailing edge cutback, the pressure distribution along the internal ribbed passage as well as slot discharge coefficients for different blowing ratios ranging from M = 0.35 to 1.1.

Effect of a Cutback Squealer and Cavity Depth on Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Shantanu Mhetras ◽  
Diganta Narzary ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cavity Depth ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blade Tip ◽  
Film Cooling Holes

Film-cooling effectiveness from shaped holes on the near tip pressure side and cylindrical holes on the squealer cavity floor is investigated. 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. Effects of varying blowing ratios and squealer cavity depth are also examined on film-cooling effectiveness. The film-cooling effectiveness distributions are measured on the blade tip, near tip pressure side and the inner pressure side and suction side rim walls using pressure sensitive paint technique. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade with two separate serpentine loops supplying coolant to the film-cooling holes. Two rows of cylindrical film-cooling holes are arranged offset to the suction side profile and along the camber line on the tip. Another row of shaped film-cooling holes is arranged along the pressure side just below the tip. The average blowing ratio of the cooling gas is controlled to be 0.5, 1.0, 1.5, and 2.0. A five-bladed linear cascade in a blow down facility with a tip gap clearance of 1.5% is used to perform the experiments. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,480,000 and the inlet and exit Mach numbers were 0.23 and 0.65, respectively. A blowing ratio of 1.0 is found to give best results on the pressure side, whereas the tip surfaces forming the squealer cavity give best results for M=2. Results show high film-cooling effectiveness magnitudes near the trailing edge of the blade tip due to coolant accumulation from upstream holes in the tip cavity. A squealer depth with a recess of 2.1mm causes the average effectiveness magnitudes to decrease slightly as compared to a squealer depth of 4.2mm.

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