Spanwise Differences in Static and Dynamic Stall on a Pitching Rotor Blade Tip Model

A high-Reynolds-number pump (HIREP) facility has been used to acquire flow measurements in the rotor blade tip clearance region, with blade chord Reynolds numbers of 3,900,000 and 5,500,000. The initial experiment involved rotor blades with varying tip clearances, while a second experiment involved a more detailed investigation of a rotor blade row with a single tip clearance. The flow visualization on the blade surface and within the flow field indicate the existence of a trailing-edge separation vortex, a vortex that migrates radially upward along the trailing edge and then turns in the circumferential direction near the casing, moving in the opposite direction of blade rotation. Flow visualization also helps in establishing the trajectory of the tip leakage vortex core and shows the unsteadiness of the vortex. Detailed measurements show the effects of tip clearance size and downstream distance on the structure of the rotor tip leakage vortex. The character of the velocity profile along the vortex core changes from a jetlike profile to a wakelike profile as the tip clearance becomes smaller. Also, for small clearances, the presence and proximity of the casing endwall affects the roll-up, shape, dissipation, and unsteadiness of the tip leakage vortex. Measurements also show how much circulation is retained by the blade tip and how much is shed into the vortex, a vortex associated with high losses.

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Heat Transfer and Flow on the Squealer Tip of a Gas Turbine Blade

Volume 3: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/2000-gt-0195 ◽

2000 ◽

Cited By ~ 12

Author(s):

Gm S. Azad ◽

Je-Chin Han ◽

Robert J. Boyle

Keyword(s):

Heat Transfer ◽

Heat Transfer Coefficient ◽

Transfer Coefficient ◽

Gas Turbine ◽

Rotor Blade ◽

Static Pressure ◽

Cavity Surface ◽

Edge Region ◽

Blade Tip ◽

The Heat Transfer Coefficient

Experimental investigations are performed to measure the detailed heat transfer coefficient and static pressure distributions on the squealer tip of a gas turbine blade in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a modern first stage gas turbine rotor blade with a blade tip profile of a GE-E3 aircraft gas turbine engine rotor blade. A squealer (recessed) tip with a 3.77% recess is considered here. The data on the squealer tip are also compared with a flat tip case. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Two different turbulence intensities of 6.1% and 9.7% at the cascade inlet are also considered for heat transfer measurements. Static pressure measurements are made in the mid-span and near-tip regions, as well as on the shroud surface opposite to the blade tip surface. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and an exit Reynolds number based on the axial chord of 1.1×106. A transient liquid crystal technique is used to measure the heat transfer coefficients. Results show that the heat transfer coefficient on the cavity surface and rim increases with an increase in tip clearance. The heat transfer coefficient on the rim is higher than the cavity surface. The cavity surface has a higher heat transfer coefficient near the leading edge region than the trailing edge region. The heat transfer coefficient on the pressure side rim and trailing edge region is higher at a higher turbulence intensity level of 9.7% over 6.1% case. However, no significant difference in local heat transfer coefficient is observed inside the cavity and the suction side rim for the two turbulence intensities. The squealer tip blade provides a lower overall heat transfer coefficient when compared to the flat tip blade.

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Investigation of Helicopter Rotor-Blade-Tip-Vortex Alleviation Using a Slotted Tip

AIAA Journal ◽

10.2514/1.3254 ◽

2004 ◽

Vol 42 (3) ◽

pp. 524-535 ◽

Cited By ~ 21

Author(s):

Yong Oun Han ◽

J. Gordon Leishman

Keyword(s):

Rotor Blade ◽

Tip Vortex ◽

Helicopter Rotor ◽

Helicopter Rotor Blade ◽

Blade Tip

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Back Flow in Turbines of Nuclear Closed-Cycle Gas Turbine Plants in Case of Circuit Pipe Rupture

Journal of Engineering for Power ◽

10.1115/1.3445957 ◽

1975 ◽

Vol 97 (2) ◽

pp. 189-194 ◽

Cited By ~ 6

Author(s):

K. Bammert ◽

P. Zehner

Keyword(s):

Gas Turbine ◽

Rotor Blade ◽

Back Flow ◽

Blade Tip ◽

Stator Blade ◽

Flow Through ◽

High Temperature Reactor ◽

Turbine Cascades ◽

Characteristic Features ◽

Safety Considerations

For operation of a gas turbine in single-cycle arrangement with a high-temperature reactor, rupture of a main circuit pipe has to be included in the safety considerations. In the event of such an accident there may be a back flow through the turbo machines or a forward flow up to the choking limit. This paper is a report on tests carried out in a two-dimensional cascade wind tunnel on turbine cascades under back flow conditions. By the example of three selected representative cascades the characteristic features in turbine cascades with back flow are discussed. These cascades are a rotor blade tip section with aerofoil-like profiles and a wide pitch, a stator blade or rotor blade mean section with an usual deflection and a rotor blade root section with a narrow pitch and a large deflection.

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Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

Journal of Turbomachinery ◽

10.1115/1.1308567 ◽

2000 ◽

Vol 122 (4) ◽

pp. 717-724 ◽

Cited By ~ 98

Author(s):

Gm. S. Azad ◽

Je-Chin Han ◽

Shuye Teng ◽

Robert J. Boyle

Keyword(s):

Heat Transfer ◽

Heat Transfer Coefficient ◽

Transfer Coefficient ◽

Gas Turbine ◽

Turbine Blade ◽

Turbulence Intensity ◽

Rotor Blade ◽

Static Pressure ◽

Pressure Distributions ◽

Blade Tip

Heat transfer coefficient and static pressure distributions are experimentally investigated on a gas turbine blade tip in a five-bladed stationary linear cascade. The blade is a two-dimensional model of a first-stage gas turbine rotor blade with a blade tip profile of a GE-E3 aircraft gas turbine engine rotor blade. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1×106. The middle 3-blade has a variable tip gap clearance. All measurements are made at three different tip gap clearances of about 1, 1.5, and 2.5 percent of the blade span. Heat transfer measurements are also made at two different turbulence intensity levels of 6.1 and 9.7 percent at the cascade inlet. Static pressure measurements are made in the midspan and the near-tip regions as well as on the shroud surface, opposite the blade tip surface. Detailed heat transfer coefficient distributions on the plane tip surface are measured using a transient liquid crystal technique. Results show various regions of high and low heat transfer coefficient on the tip surface. Tip clearance has a significant influence on local tip heat transfer coefficient distribution. Heat transfer coefficient also increases about 15–20 percent along the leakage flow path at higher turbulence intensity level of 9.7 over 6.1 percent. [S0889-504X(00)00404-9]

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Experimental Characterization of the Evolution of Global Flow Structure in the Passage of an Axial Compressor

10.1115/gt2021-60325 ◽

2021 ◽

Author(s):

Ayush Saraswat ◽

Subhra Shankha Koley ◽

Joseph Katz

Keyword(s):

Flow Structure ◽

Rotor Blade ◽

Incidence Angle ◽

Substantial Reduction ◽

Axial Compressor ◽

Suction Side ◽

Rotor Wake ◽

Tip Leakage ◽

Blade Tip

Abstract Ongoing experiments conducted in a one-and-half stages axial compressor installed in the JHU refractive index-matched facility investigate the evolution of flow structure across blade rows. After previously focusing only on the rotor tip region, the present stereo-PIV (SPIV) measurements are performed in a series of axial planes covering an entire passage across the machine, including upstream of the IGV, IGV-rotor gap, rotor-stator gap, and downstream of the stator. The measurements are performed at flow rates corresponding to pre-stall condition and best efficiency point (BEP). Data are acquired for various rotor-blade orientations relative to the IGV and stator blades. The results show that at BEP, the wakes of IGV and rotor are much more distinct and the wake signatures of one row persists downstream of the next, e.g., the flow downstream of the stator is strongly affected by the rotor orientation. In contrast, under pre-stall conditions, the rotor orientation has minimal effect on the flow structure downstream of the stator. However, the wakes of the stator blades, where the axial momentum is low, are now wider. For both conditions, the flow downstream of the rotor is characterized by two regions of axial momentum deficit in addition to the rotor wake. A deficit on the pressure side of the rotor wake tip is associated with the tip leakage vortex (TLV) of the previous rotor blade, and is much broader at pre-stall condition. A deficit on the suction side of the rotor wake near the hub appears to be associated with the hub vortex generated by the neighboring blade, and is broader at BEP. At pre-stall, while the axial momentum upstream of the rotor decreases over the entire tip region, it is particularly evident near the rotor blade tip, where the instantaneous axial velocity becomes intermittently negative. Downstream of the rotor, there is a substantial reduction in mean axial momentum in the upper half of the passage, concurrently with an increase in the circumferential velocity. Consequently, the incidence angle upstream of the stator increases in certain regions by up to 30 degrees. These observations suggest that while the onset of the stall originates from the rotor tip flow, one must examine its impact on the flow structure in the stator passage as well.

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Prediction of Unshrouded Rotor Blade Tip Heat Transfer

Volume 4: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/95-gt-142 ◽

1995 ◽

Cited By ~ 25

Author(s):

A. A. Ameri ◽

E. Steinthorsson

Keyword(s):

Heat Transfer ◽

Rotor Blade ◽

Stokes Equations ◽

Space Shuttle ◽

Suction Side ◽

Computer Code ◽

Turbine Rotor ◽

High Rate ◽

Blade Surface ◽

Blade Tip

The rate of heat transfer on the tip of a turbine rotor blade and on the blade surface in the vicinity of the tip, was successfully predicted. The computations were performed with a multiblock computer code which solves the Reynolds Averaged Navier-Stokes equations using an efficient multigrid method. The case considered for the present calculations was the SSME (Space Shuttle Main Engine) high pressure fuel side turbine. The predictions of the blade tip heat transfer agreed reasonably well with the experimental measurements using the present level of grid refinement. On the tip surface, regions with high rate of heat transfer was found to exist close to the pressure side and suction side edges. Enhancement of the heat transfer was also observed on the blade surface near the tip. Further comparison of the predictions was performed with results obtained from correlations based on fully developed channel flow.

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