Design Considerations for Tip Clearance Control and Measurement on a Turbine Rainbow Rotor With Multiple Blade Tip Geometries

2016 ◽  
Vol 139 (4) ◽  
Author(s):  
S. Lavagnoli ◽  
C. De Maesschalck ◽  
V. Andreoli
Keyword(s):  
Large Scale ◽  
Numerical Procedure ◽  
International Standards ◽  
Tip Clearance ◽  
Radial Deformation ◽  
Numerical Predictions ◽  
Blade Tip ◽  
Rotor Disk ◽  
Clearance Control

The accurate design, control, and monitoring of the running gaps between static and moving components are vital to preserve the mechanical integrity and ensure the correct functioning of any compact rotating machinery. Throughout engine service, the rotor tip clearance undergoes large variations due to installation tolerances or as the result of different thermal expansion rates of the blades, rotor disk, and casing during speed transients. Hence, active tip clearance control concepts and engine health-monitoring systems rely on precise real-time gap measurements. Moreover, this tip gap information is crucial for engine development programs to verify the mechanical and aerothermal designs and validate numerical predictions. This paper presents an overview of the critical design requirements for testing engine-representative blade tip flows in a rotating turbine facility. This paper specifically focuses on the challenges related with the design, verification, and monitoring of the running tip clearance during a turbine experiment. In the large-scale turbine facility of the von Karman Institute, a rainbow rotor was mounted for simultaneous aerothermal testing of multiple blade tip geometries. The tip shapes are a selection of high-performance squealer-like and contoured blade tip designs. On the rotor disk, the blades are arranged in seven sectors operating at different clearance levels from 0.5 up to 1.5% of the blade span. Prior to manufacturing, the blade geometry was modified to compensate for the radial deformation of the rotating assembly under centrifugal loads. A numerical procedure was implemented to minimize the residual unbalance of the rotor in rainbow configuration and to optimize the placement of every single airfoil within each sector. Subsequently, the rotor was balanced in situ to reduce the vibrations and satisfy the international standards for high balance quality. Three fast-response capacitive probes located at distinct circumferential locations around the rotor annulus measured the single-blade tip clearance in rotation. Additionally, the minimum running blade clearance is captured with wear gauges located at five axial positions along the blades chord. The capacitance probes are self-calibrated using a multitest strategy at several rotational speeds. The in situ calibration methodology and dedicated data reduction techniques allow the accurate measurement of the distance between the turbine casing and the local blade tip features (rims and cavities) for each rotating airfoil separately. General guidelines are given for the design and calibration of a tip clearance measurement system that meets the required measurement accuracy and resolution in function of the sensor uncertainty, nominal tip clearance levels, and tip seal geometry.

Design Considerations for Tip Clearance Control and Measurement on a Turbine Rainbow Rotor With Multiple Blade Tip Geometries

2016 ◽  
Author(s):  
S. Lavagnoli ◽  
C. De Maesschalck ◽  
V. Andreoli
Keyword(s):  
High Performance ◽  
Large Scale ◽  
Numerical Procedure ◽  
International Standards ◽  
Tip Clearance ◽  
Radial Deformation ◽  
Numerical Predictions ◽  
Blade Tip ◽  
Clearance Control

The accurate design, control and monitoring of the running gaps between static and moving components is vital to preserve the mechanical integrity and ensure the correct functioning of any compact rotating machinery. Throughout engine service, the rotor tip clearance undergoes large variations due to installation tolerances or as the result of different thermal expansion rates of the blades, rotor disk and casing during speed transients. Hence, active tip clearance control concepts and engine health monitoring systems rely on precise real-time gap measurements. Moreover, this tip gap information is crucial for engine development programs to verify the mechanical and aerothermal design, and validate numerical predictions. This paper presents an overview of the critical design requirements for testing engine-representative blade tip flows in a rotating turbine facility. The manuscript specifically focuses on the challenges related with the design, verification and monitoring of the running tip clearance during a turbine experiment. In the large-scale turbine facility of the von Karman Institute, a rainbow rotor was mounted for simultaneous aerothermal testing of multiple blade tip geometries. The tip shapes are a selection of high-performance squealer-like and contoured blade tip designs. On the rotor disc, the blades are arranged in seven sectors operating at different clearance levels from 0.5 up to 1.5% of the blade span. Prior to manufacturing, the blade geometry was modified to compensate for the radial deformation of the rotating assembly under centrifugal loads. A numerical procedure was implemented to minimize the residual unbalance of the rotor in rainbow configuration, and to optimize the placement of every single airfoil within each sector. Subsequently, the rotor was balanced in-situ to reduce the vibrations and satisfy the international standards for high balance quality. The single-blade tip clearance in rotation was measured by three fast-response capacitive probes located at three distinct circumferential locations around the rotor annulus. Additionally, the minimum running blade clearance is captured with wear gauges located at five axial positions along the blades chord. The capacitance probes are self-calibrated using a multi-test strategy at several rotational speeds. The in-situ calibration methodology and dedicated data reduction techniques allow the accurate measurement of the distance between the turbine casing and the local blade tip features (rims and cavities) for each rotating airfoil separately. General guidelines are given for the design and calibration of a tip clearance measurement system that meets the required measurement accuracy and resolution in function of the sensor uncertainty, nominal tip clearance levels and tip seal geometry.

scholarly journals Influences of Tip Cooling Injection on Tip Clearance Control at Design and Off-Design Incidences

2009 ◽  
Vol 2009 ◽  
pp. 1-12 ◽  
Author(s):  
Maosheng Niu ◽  
Shusheng Zang
Keyword(s):  
Heat Transfer ◽  
Tip Clearance ◽  
Air Injection ◽  
Turbine Cascade ◽  
Tip Clearance Flow ◽  
Clearance Flow ◽  
Blade Tip ◽  
Design Value ◽  
Tip Injection ◽  
Clearance Control

A numerical investigation has been performed to study the influences of cooling injection from the blade tip surface on controlling tip clearance flow in an unshrouded, high-turning axial turbine cascade. Emphasis is put on the analysis of the effectiveness of tip injection when the approaching flow is at design and off-design incidences. A total of three incidence angles are investigated, 7.4°, 0°, 0°, 0°, and 7.6°, 0° relative to the design value. The results indicate that even at the off-design incidences, tip injection can also act as an obstruction to the tip clearance flow and weaken the interaction between the passage flow and the tip clearance flow. It is also found that tip injection causes the tip clearance loss to be less sensitive to the incidences. Moreover, with injection, at all these incidences the heat transfer conditions are improved significantly on the blade tip surface in the middle and aft parts of blade. Thus, tip injection is proved to be an effective method of controlling tip clearance flow, even at off-design conditions. Beside that, an indirect empirical correlation is observed to be able to perform well in predicting the losses induced by tip clearance flow at design and off-design conditions, no matter whether air injection is active or not.

Blade Tip Heat Transfer and Aerodynamics in a Large Scale Turbine Cascade With Moving Endwall

2006 ◽  
Author(s):  
P. Palafox ◽  
M. L. G. Oldfield ◽  
P. T. Ireland ◽  
T. V. Jones ◽  
J. E. LaGraff
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Large Scale ◽  
Infrared Camera ◽  
Suction Side ◽  
Constant Heat Flux ◽  
Tip Clearance ◽  
Airfoil Surface ◽  
Nusselt Number Distribution ◽  
Blade Tip

High resolution Nusselt number (Nu) distributions were measured on the blade tip surface of a large, 1.0 meter-chord, low-speed cascade representative of a high-pressure turbine. Data was obtained at a Reynolds number of 4.0 × 105 based on exit velocity and blade axial chord. Tip clearance levels ranged from 0.56% to 1.68% design span or equally from 1% to 3% of blade chord. An infrared camera, looking through the hollow blade, made detailed temperature measurements on a constant heat flux tip surface. The relative motion between the endwall and the blade tip was simulated by a moving belt. The moving belt endwall significantly to shifts the region of high Nusselt number distribution and reduces the overall averaged Nusselt number on the tip surface by up to 13.3%. The addition of a suction side squealer tip significantly reduced local tip heat transfer and resulted in a 32% reduction in averaged Nusselt number. Analysis of pressure measurements on the blade airfoil surface and tip surface along with PIV velocity flow fields in the gap give an understanding of the heat transfer mechanism.

Experimental Study of Effects of Grooved Tip Clearances on the Flow Field in a Compressor Cascade Passage

2010 ◽  
Author(s):  
Hongwei Ma ◽  
Jun Zhang ◽  
Jinghui Zhang ◽  
Zhou Yuan
Keyword(s):  
Flow Field ◽  
Large Scale ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Probe ◽  
Leakage Flow ◽  
Compressor Cascade ◽  
High Loss ◽  
Static Pressure Distribution ◽  
Blade Tip

This paper presents an experimental investigation of effects of grooved tip clearances on the flow field of a compressor cascade. The tests were performed in a low-speed large-scale cascade respectively with two tip clearance configurations, including flat tip and grooved tip with a chordwise channel on the blade top. The flow field at 10% chord downstream from the cascade trailing edge was measured at four incidence angles using a mini five-hole pressure probe. The static pressure distribution was measured on the tip endwall. The results show that the pressure gradient from the pressure side to the suction side on the blade tip is reduced due to the existence of the channel. As a result, the leakage flow is weakened. The high-blockage and high-loss region caused by the leakage flow is narrower with the grooved tip. In the meantime, the leakage flow migrates to lower spanwise position. The combined result is that the flow capacity in the tip region is improved at the incidence angles of 0° and 5° with the grooved tip. However, the loss is slightly greater than that with the flat tip at all the incidence angles.

A Novel Generalized Predictive Control and its Application on Active Clearance Control of Aero-Engine

2012 ◽  
Vol 616-618 ◽  
pp. 1922-1925
Author(s):  
Kai Peng ◽  
Ding Fan ◽  
Lei Zhang ◽  
Qiu Xia Wang
Keyword(s):  
Predictive Control ◽  
Gas Turbines ◽  
Dynamic Performance ◽  
Tip Clearance ◽  
Generalized Predictive Control ◽  
Aero Engine ◽  
Blade Tip ◽  
Additional Service ◽  
And Control ◽  
Clearance Control

Turbine blade tip clearance continues to be a concern in the design and control of gas turbines. Ever increasing demands for improved efficiency and higher operating temperatures require more stringent tolerances on turbine tip clearance. An implicit active generalized predictive control with AR error modification and fuzzy adjustment on control horizon of aero-engine turbine tip clearance is presented and evaluated. The results show the resultant active tip clearance control system has good steady and dynamic performance and benefits of increased efficiency, reduced specific fuel consumption, and additional service life.

Transonic Passage Turbine Blade Tip Clearance With Scalloped Shroud: Part III — Heat Transfer in Engine Configuration

2004 ◽  
Author(s):  
F. Casey Wilkins ◽  
Gregory M. Feldman ◽  
Wayne S. Strasser ◽  
James H. Leylek
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Large Scale ◽  
Numerical Study ◽  
Viscous Sublayer ◽  
Tip Clearance ◽  
Transfer Coefficients ◽  
Constant Wall Heat Flux ◽  
Blade Tip ◽  
Large Scale Land

This work presents a numerical study that was done to investigate the heat transfer characteristics of a transonic turbine blade with a scalloped shroud operating at realistic engine conditions typical of those found in a large scale, land-based gas turbine. The geometry under investigation was an infinite, linear cascade composed of the same blade and shroud design used in an experimental test rig by the research sponsor. This simulation was run for varying nominal tip clearances of 20, 80, and 5.08 mm. For each of these clearances, the simulation was run with and without the scrubbing effects of the outer casing, resulting in a total of six cases that could be used to determine the influence of tip clearance and relative casing motion on heat transfer. A high quality grid (ranging from approximately 10–12 million finite volumes depending on tip clearance) with y+ for first layer cells at or below 1.0 everywhere was used to resolve the flow down to the viscous sublayer. The “realizable” k-ε turbulence model was used for all cases. A constant wall heat flux was imposed on all the surrounding surfaces to obtain heat transfer data. Results produced include a full map of heat transfer coefficients for the suction and pressure surfaces of the blade as well as the tip, shroud, and outer casing for every case. Physical mechanisms responsible for the final heat transfer outcome for all six cases are documented.

Detailed Heat Transfer Coefficient Distributions on a Large-Scale Gas Turbine Blade Tip

2000 ◽  
Vol 123 (4) ◽  
pp. 803-809 ◽  
Author(s):  
Shuye Teng ◽  
Je-Chin Han ◽  
G. M. S. Azad
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Wind Tunnel ◽  
Transfer Coefficient ◽  
Turbine Blade ◽  
Large Scale ◽  
Tip Clearance ◽  
Low Speed ◽  
Blade Tip ◽  
Low Speed Wind Tunnel

Measurements of detailed heat transfer coefficient distributions on a turbine blade tip were performed in a large-scale, low-speed wind tunnel facility. Tests were made on a five-blade linear cascade. The low-speed wind tunnel is designed to accommodate the 107.49 deg turn of the blade cascade. The mainstream Reynolds number based on cascade exit velocity was 5.3×105. Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. The central blade had a variable tip gap clearance. Measurements were made at three different tip gap clearances of about 1.1 percent, 2.1 percent, and 3 percent of the blade span. Static pressure distributions were measured in the blade mid-span and on the shroud surface. Detailed heat transfer coefficient distributions were measured on the blade tip surface using a transient liquid crystal technique. Results show that reduced tip clearance leads to reduced heat transfer coefficient over the blade tip surface. Results also show that reduced tip clearance tends to weaken the unsteady wake effect on blade tip heat transfer.

Blade Tip Heat Transfer and Aerodynamics in a Large Scale Turbine Cascade With Moving Endwall

2011 ◽  
Vol 134 (2) ◽  
Author(s):  
P. Palafox ◽  
M. L. G. Oldfield ◽  
P. T. Ireland ◽  
T. V. Jones ◽  
J. E. LaGraff
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Large Scale ◽  
Infrared Camera ◽  
Suction Side ◽  
Constant Heat Flux ◽  
Tip Clearance ◽  
Airfoil Surface ◽  
Nusselt Number Distribution ◽  
Blade Tip

High resolution Nusselt number distributions were measured on the blade tip surface of a large, 1.0 m chord, low-speed cascade representative of a high-pressure turbine. Data were obtained at a Reynolds number of 4.0×105 based on exit velocity and blade axial chord. Tip clearance levels ranged from 0.56% to 1.68% design span or equally from 1% to 3% of the blade chord. An infrared camera, looking through the hollow blade, made detailed temperature measurements on a constant heat flux tip surface. The relative motion between the endwall and the blade tip was simulated by a moving belt. The moving belt endwall significantly shifts the region of high Nusselt number distribution and reduces the overall averaged Nusselt number on the tip surface by up to 13.3%. The addition of a suction side squealer tip significantly reduced local tip heat transfer and resulted in a 32% reduction in averaged Nusselt number. Analysis of pressure measurements on the blade airfoil surface and tip surface along with particle image velocimetry velocity flow fields in the gap gives an understanding of the heat transfer mechanism.

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