Shantanu Mhetras

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.

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

Impingement Heat Transfer From Jet Arrays on Turbulated Target Walls at Large Reynolds Numbers

Experiments to investigate heat transfer and pressure loss from jet array impingement are performed on the target wall at high Reynolds numbers. Reynolds numbers up to 450,000 are tested. The presence of a turbulated target wall and its effect on heat transfer enhancement against a smooth surface is investigated. Two different jet plate configurations are used with closely spaced holes and with angled as well as normal impingement holes. The test section cross-section is designed to expand along the streamwise direction maintaining the jet plate to target wall distance in order to reduce cross-flow effects. The jet plate holes are chamfered or filleted to minimize pressure loss through the jet plate. Heat transfer and pressure loss measurements are performed on a smooth target wall as well as turbulated target walls. Three turbulators configurations are used with streamwise riblets, short pins and spherical dimples. A steady state heat transfer measurement method is used to obtain the heat transfer coefficients while pressure taps located in the plenum and at several streamwise locations are used to record the pressure losses across the jet plate. Experiments are performed for a range of Reynolds numbers from 50,000 to 450,000 based on average jet hole diameters to cover the incompressible as well as compressible flow regimes. A target wall with short pins provides the best heat transfer with the dimpled target wall giving the lowest heat transfer among the three turbulators geometries studied. Addition of turbulators though does not significantly increase the pressure losses in the test section over the smooth target wall.Copyright

Heat Transfer and Pressure Loss Measurements in a Turbulated High Aspect Ratio Channel With Large Reynolds Number Flows

Experiments to investigate heat transfer and pressure loss are performed in a rectangular channel with an aspect ratio of 6 at very high Reynolds numbers under compressible flow conditions. Reynolds numbers up to 1.3 million are tested. The presence of a turbulated wall and the resultant heat transfer enhancement against a smooth surface is investigated. Three dimpled configurations including spherical and cylindrical dimples are studied on one wide wall of the channel. The presence of discrete ribs on the same wide wall is also investigated. A steady state heat transfer measurement method is used to obtain the heat transfer coefficients while pressure taps located at several streamwise locations in the channel walls are used to record the static pressures on the surface. Experiments are performed for a range of Reynolds numbers from 100,000 to 1,300,000 to cover the incompressible as well as compressible flow regimes. Results for low Reynolds numbers are compared to existing heat transfer data available in open literature for similar configurations. Heat transfer enhancement is found to decrease at high Re with the discrete rib configurations providing the best enhancement but highest pressure losses. Local measurements using the steady state, hue-detection based liquid crystal technique are also performed in the fully developed region for case 1 with dimples. Good comparison is obtained between averaged local heat transfer coefficient measurements and from thermocouple measurements.Copyright