Diganta Narzary

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

This paper is focused on the effect of film-hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film-cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film-cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film-cooling holes are divided into four rows and loosely distributed on the platform. Four film-cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete-hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film-cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.

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

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-E 3 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 2.1 mm causes the average effectiveness magnitudes to decrease slightly as compared to a squealer depth of 4.2 mm.

Full-Coverage Film Cooling for a Turbine Blade with Axial-Shaped Holes

[Abstract] The film cooling effectiveness on a fully film cooled high pressure turbine blade is investigated using the Pressure Sensitive Paint technique. Three rows of radial angled cylindrical holes are arranged in the leading edge region, while axial laidback fanshaped holes are provided on the blade surfaces with four rows on the pressure side and two rows on the suction side. The shaped holes are featured with 10° lateral expansion from the hole centerline and additional 10° forward expansion to the blade surface. The coolant ejects through all the film cooling holes at four average blowing ratios ranging from 0.3 to 1.2. The influence of wake from upstream vane is simulated by placing a periodic set of rods upstream of the test blade. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The Mach numbers at the inlet and the exit are 0.27 and 0.44, respectively, resulting in a blade pressure ratio of 1.14. Results show the film cooling effectiveness increases with increasing of average blowing ratio. The presence of upstream wake rods can be very detrimental to the film effectiveness on the blade surface depending on the wake rod phase location. Compared with the case of no showerhead injection, the spanwsie averaged film cooling effectiveness in the downstream region of the pressure side and suction side rows increases with the showerhead injection. The film effectiveness on the pressure side is comparable for compound angle shaped holes and axial shaped holes; while on the suction side, the compound angle shaped holes provide better effectiveness than the axial shaped holes due to jet deflection and expanded hole breakout area..

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

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-E 3 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 up-stream 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.

Effect of Unsteady Wake on Film-Cooling Effectiveness Distribution on a Gas Turbine Blade With Compound Shaped Holes

The effect of fan-shaped, laid-back compound angled cooling holes placed along the span of a fully-cooled high pressure turbine blade in a 5-blade linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Four rows of shaped film cooling holes are provided on the pressure side while two such rows are provided on the suction side of the blade. Three rows of cylindrical holes are drilled at 30° to the surface on the leading edge to capture the effect of showerhead film coolant injection. The coolant is injected at four different average blowing ratios of 0.3, 0.6, 0.9 and 1.2. Presence of wake due to upstream vanes is studied by placing a periodic set of rods upstream of the test blade. The wake is generated using 4.8mm diameter rods. The wake rods can be clocked by changing their stationary positions in front of the test blade to simulate a progressing wake. Effect of wake is recorded at four phase locations with equal intervals. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively resulting in a blade pressure ratio of 1.14. Turbulence intensity level at the cascade inlet is 6% with an integral length scale of around 5cm. Results show that the fan-shaped, laid-back compound angled holes produce uniform and wide coolant coverage on the suction side except for those regions affected by the passage and tip leakage vortices. The advantage of compound shaped hole design is seen from the higher effectiveness values on the suction side compared to that of the compound cylindrical holes. The presence of a stationary upstream wake can result in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.3 to 1.2 show more or less uniform increment in effectiveness increase on the pressure side, whereas on the suction side, the increment shows signs of saturation beyond M = 0.6.Copyright

Influence of Shock Wave on Turbine Vane Suction Side Film Cooling With Compound-Angle Shaped Holes

This paper studies the effect of shock wave on turbine vane suction side film cooling using a conduction-free Pressure Sensitive Paint (PSP) technique. Tests were performed in a five-vane annular cascade with a blow-down flow loop facility. The exit Mach numbers are controlled to be 0.7, 1.1, and 1.3, from subsonic to transonic flow conditions. Two foreign gases N2 and CO2 are selected to study the effects of two coolant-to-mainstream density ratios, 1.0 and 1.5, on film cooling. Four averaged coolant blowing ratios in the range, 0.4 to 1.6 are investigated. The test vane features 3 rows of radial-angle cylindrical holes around the leading edge, and 2 rows of compound-angle shaped holes on the suction side. Results suggest that the PSP is an accurate technique capable of producing clear and detailed film cooling effectiveness contours at transonic flow conditions. At lower blowing ratio, film cooling effectiveness decreases with increasing exit Mach number. On the other hand, an opposite trend is observed at high blowing ratio. In transonic flow, the rapid rise in pressure caused by shock benefits film-cooling by deflecting the coolant jet toward the vane surface at higher blowing ratio. Results show that denser coolant performs better, typically at higher blowing ratio in transonic flow. Results also show that the optimum momentum flux ratio decreases with density ratio at subsonic condition. In transonic flow, however, the trend is reversed and the peak effectiveness values plateau over a long range of momentum flux ratio.Copyright

Heat Transfer and Thermal Mechanical Stress Distributions in Gas Turbine Blades

In this paper the distributions of heat transfer and thermal mechanical stress in the metal blade surface are investigated. The stream that surrounds the blade was considered at the time that the cooling airflow runs through the blade interior. Cooling channel flow and gases were simulated using a finite volume program, Fluent. The conjugate problem was addressed using coupled domains solid-fluid. Beside the numerical approach, measurements of metal blade surface temperature distributions based on the temperature sensitive paint technique, TSP, were conducted. The cooling effectiveness was compared showing good agreement between computational/experimental results. Additionally to laboratory conditions, finite volume results were obtained for real engine operating conditions. These results were used to establish temperature boundary conditions into a second computational model programmed in ANSYS, based on finite elements. This second model allowed calculating the distribution of thermo-mechanical stress in the blade material. The results show the temperature distribution in the blade surface. Based on this, the heat transfer rate was calculated finding it as a strong function of position. The cooling effectiveness was also calculated, which in turn performs with less variation over the sections of the blade under investigation. Following, the thermal effects in the metal blade surface lead to calculate the stress distribution. Differences in stresses magnitude were also found, suggesting a strong correlation between heat transfer and stress in the metal blade surface.Copyright

Influence of Cooling Flow Rate Variation on Gas Turbine Blade Temperature Distributions

Temperature and flow rate of combustion gases and cooling stream are essential conditions for blade integrity in gas turbines. Since the combustion products pass directly to the first stage of blades high thermal stresses can develop, so the temperature field in the blade material must be controlled to avoid damage and/or reduction of blade useful life. This paper discusses an investigation on the influence of cooling airflow reduction on blade life. The flow rate reduction under consideration may be due to malfunctions of the compressor such like deposits or partial blockage in the blade ducts. It has been reported that air discharge from the compressor can be reduced up to 15% of the nominal rate due to deposits related with impurities contained in the environment. In this work an evaluation of the effect of reducing the cooling airflow rate on the temperature distribution on the blades surface is attempted. The flow stream that surrounds the blade together with the cooling airflow in the blade interior channels were characterized in the laboratory. Fields of temperature on the blade surface were obtained using the temperature sensitive paint technique, TSP. Thermocouple measurements were used for punctual temperatures as a reference. The results showed the regions of possible thermal stresses concentration as a function of cooling airflow rate variations. Additionally, the problem was resolved computationally in conjugate mode, considering both fluid streams external and internal plus heat conduction at the interior of the blade material. The computer model is used to simulate other conditions not addressed in the experiment. The paper discusses the comparison of numerical to experimental results and discusses the methodology for further work.Copyright

Turbine Blade Tip Film-Cooling and Heat Transfer Measurements at High Blowing Ratios

Film-cooling and heat transfer characteristics of a gas turbine blade tip with a suction side rail was investigated in a stationary 3-blade rectilinear cascade. Mounted at the end of a blow-down facility the cascade operated at inlet and exit Mach numbers of 0.29 and 0.75, respectively. The rail was marginally offset from the suction side edge of the tip and extended from the leading to the trailing edge. A total of 17 film-cooling holes were placed along the near-tip pressure side surface and 3 on the near-tip leading edge surface with the objective of providing coolant to the tip. The tip surface itself did not carry any film-cooling holes. Relatively high blowing ratios of 2.0, 3.0, 4.0, and 4.5 and three tip gaps of 0.87%, 1.6%, and 2.3% of blade span made up the test matrix. Pressure sensitive paint (PSP) and Thermo-Chromic Liquid Crystal (TLC) were the experimental techniques employed to measure film-cooling effectiveness and heat transfer coefficient, respectively. Results indicated that when the tip gap was increased, film-cooling effectiveness on the tip surface decreased and heat transfer to the tip surface increased. On the other hand, when the blowing ratio was increased, film effectiveness increased but the effect on heat transfer coefficient was relatively small. The highest heat transfer coefficient levels were found atop the suction side rail, especially in the downstream two-thirds of its length whereas the lowest levels were found on the tip floor in the widest section of the blade.Copyright