Development of an Experimental Capability to Produce Controlled Blade Tip/Shroud Rubs at Engine Speed

2004 ◽  
Author(s):  
Corso Padova ◽  
Jeffery Barton ◽  
Michael G. Dunn ◽  
Steve Manwaring ◽  
Gamaliel Young ◽  
...  
Keyword(s):  
Engine Speed ◽  
Facility Design ◽  
Bladed Disks ◽  
Current Configuration ◽  
Blade Tip ◽  
Design Speed ◽  
Tip Shroud ◽  
Compressor Disk

Development of an in-ground spin-pit facility specifically designed to investigate aeromechanic phenomena for engine hardware rotating at design speed is reported in this paper. The purpose of this paper is to describe the facility design and operation and to demonstrate utility by providing typical results from a recently completed measurement program. The facility is designed to allow insertion of a segment of engine casing into the path of single-bladed or multiple-bladed disks. In the current configuration, a 90-degree sector of a representative engine casing is forced to rub the tip of a single-bladed compressor disk with predetermined blade incursion into the casing for rotational speeds in the vicinity of 20,000 rpm.

Development of an Experimental Capability to Produce Controlled Blade Tip∕Shroud Rubs at Engine Speed

2004 ◽  
Vol 127 (4) ◽  
pp. 726-735 ◽  
Author(s):  
Corso Padova ◽  
Jeffrey Barton ◽  
Michael G. Dunn ◽  
Steve Manwaring ◽  
Gamaliel Young ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Engine Speed ◽  
Gas Turbine Engine ◽  
Specific Information ◽  
Bladed Disks ◽  
Turbine Engine ◽  
Blade Tip ◽  
Tip Shroud ◽  
Compressor Disk

An experimental capability using an in-ground spin-pit facility specifically designed to investigate aeromechanic phenomena for gas turbine engine hardware rotating at engine speed is demonstrated herein to obtain specific information related to prediction and modeling of blade-casing interactions. Experiments are designed to allow insertion of a segment of engine casing into the path of single-bladed or multiple-bladed disks. In the current facility configuration, a 90deg sector of a representative engine casing is forced to rub the tip of a single-bladed compressor disk for a selected number of rubs with predetermined blade incursion into the casing at rotational speeds in the vicinity of 20,000rpm.

Experimental Results From Controlled Blade Tip/Shroud Rubs at Engine Speed

2006 ◽  
Author(s):  
Corso Padova ◽  
Jeffery Barton ◽  
Michael G. Dunn ◽  
Steve Manwaring
Keyword(s):  
Computational Models ◽  
Engine Speed ◽  
Leading Edge ◽  
Experimental Results ◽  
Trailing Edge ◽  
Metal On Metal ◽  
Non Linear ◽  
Blade Tip ◽  
Force Components ◽  
Tip Shroud

Experimental results obtained for an Inconel compressor blade rubbing a steel casing at engine speed are described. Load cell, strain gauge and accelerometer measurements are discussed and then applied to analyze the metal-on-metal interaction resulting from sudden incursions of varying severity, defined by incursion depths ranging from 13 μm to 762 μm (0.0005-in to 0.030-in). The results presented describe the transient dynamics of rotor and casing vibro-impact response at engine operational speed similar to those experienced in flight. Force components at the blade tip in axial and circumferential directions for a rub of moderate incursion depth (140 μm) are compared to those for a severe rub (406 μm). Similar general trends of variation during the metal-to-metal contact are observed. However, in the nearly three-fold higher incursion the maximum incurred circumferential load increases significantly, while the maximum incurred axial load increases much less, demonstrating the non-linear nature of the rub phenomena. Concurrently, the stress magnification on the rubbing blade at root mid-chord, at tip leading edge, and at tip trailing edge is discussed. The results point to the possibility of failure occurring first at the airfoil trailing edge. Such a failure was in fact observed in the most severe rub obtained to date in the laboratory, consistent with field observations. Computational models to analyze the non-linear dynamic response of a rotating beam with periodic pulse loading at the free-end are currently under development and are noted.

Experimental Results From Controlled Blade Tip/Shroud Rubs at Engine Speed

2006 ◽  
Vol 129 (4) ◽  
pp. 713-723 ◽  
Author(s):  
Corso Padova ◽  
Jeffery Barton ◽  
Michael G. Dunn ◽  
Steve Manwaring
Keyword(s):  
Computational Models ◽  
Engine Speed ◽  
Leading Edge ◽  
Experimental Results ◽  
Trailing Edge ◽  
Metal On Metal ◽  
Non Linear ◽  
Blade Tip ◽  
Force Components ◽  
Tip Shroud

Experimental results obtained for an Inconel® compressor blade rubbing a steel casing at engine speed are described. Load cell, strain gauge, and accelerometer measurements are discussed and then applied to analyze the metal-on-metal interaction resulting from sudden incursions of varying severity, defined by incursion depths ranging from 13μm to 762μm (0.0005in. to 0.030in.). The results presented describe the transient dynamics of rotor and casing vibro-impact response at engine operational speed similar to those experienced in flight. Force components at the blade tip in axial and circumferential directions for a rub of moderate incursion depth (140μm) are compared to those for a severe rub (406μm). Similar general trends of variation during the metal-to-metal contact are observed. However, in the nearly threefold higher incursion the maximum incurred circumferential load increases significantly, while the maximum incurred axial load increases much less, demonstrating the non-linear nature of the rub phenomena. Concurrently, the stress magnification on the rubbing blade at root mid-chord, at tip leading edge, and at tip trailing edge is discussed. The results point to the possibility of failure occurring first at the airfoil trailing edge. Such a failure was in fact observed in the most severe rub obtained to date in the laboratory, consistent with field observations. Computational models to analyze the non-linear dynamic response of a rotating beam with periodic pulse loading at the free-end are currently under development and are noted.

Controlled Fan Blade Tip/Shroud Rubs at Engine Conditions

2011 ◽  
Author(s):  
Corso Padova ◽  
Michael Dunn ◽  
Jeffrey Barton ◽  
Kevin Turner ◽  
Tod Steen
Keyword(s):  
Titanium Alloy ◽  
State University ◽  
Facility Design ◽  
University Development ◽  
Blade Tip ◽  
Experimental Facilities ◽  
The Ohio State University ◽  
Fan Blade ◽  
Engine Components ◽  
Tip Shroud

The purpose of this paper is to describe the new facility design and operation improvements, and to demonstrate utility by providing typical results obtained as part of a typical measurement program. Since 2002 a number of experiments have been conducted to generate a broad database for tip rubs using two unique experimental facilities at the Gas Turbine Laboratory of The Ohio State University. Development of an in-ground spin-pit facility specifically designed to investigate rub-in-systems for jet engine components using real hardware rotating at representative engine speeds was reported several years ago. While the original smaller facility is still in use, more recently a much larger in-ground spin-pit facility for which the basic design and operation of the blade tip/shroud incursion technique is very different from the original facility design has been commissioned, and the results of a measurement program completed using a full-scale titanium-alloy fan blade rubbing an abradable casing are presented. The Large Spin-Pit Facility [LSPF] is designed to allow rotating engine hardware from low RPM [typically a few thousands] to 18,000 rpm, using two interchangeable spindle arrangements mounted above ground onto an in-ground containment tank. The LSPF is also designed to allow the progressive insertion of a casing segment into the path of a single-bladed or multiple-bladed disk. Segments extending 90 or 120 degrees are in use for different applications. For the configuration discussed in this paper, a 90-degree segment of a representative fan casing is forced to rub the tip of a titanium-alloy fan blade at a rotational speed in the vicinity of 6000 rpm.

scholarly journals Vibration Parameters Estimation by Blade Tip-Timing in Mistuned Bladed Disks in Presence of Close Resonances

2020 ◽  
Vol 10 (17) ◽  
pp. 5930
Author(s):  
Saeed Bornassi ◽  
Christian Maria Firrone ◽  
Teresa Maria Berruti
Keyword(s):  
Frequency Response ◽  
Optimization Technique ◽  
Numerical Test ◽  
Parameters Estimation ◽  
Least Square ◽  
Response Curves ◽  
Bladed Disks ◽  
Bladed Disk ◽  
Resonance Frequencies ◽  
Blade Tip

The present paper is focused on the post processing of the data coming from the Blade Tip-Timing (BTT) sensors in the case where two very close peaks are present in the frequency response of the vibrating system. This type of dynamic response with two very close peaks can occur quite often in bladed disks. It is related to the fact that the bladed disk is not perfectly cyclic symmetric and the so called “mistuning” is present. A method based on the fitting of the BTT sensors data by means of a 2 degrees of freedom (2DOF) dynamic model is proposed. Nonlinear least square optimization technique is employed for identification of the vibration characteristics. A numerical test case based on a lump parameter model of a bladed disk assembly is used to simulate different response curves and the corresponding sensors signals. The Frequency Response Function (FRF) constructed at the resonance region is compared with the traditional Sine fitting results, the resonance frequencies and damping values estimated by the fitting procedure are also reported. Accurate predictions are achieved and the results demonstrate the considerable capacity of the 2DOF method to be used as a standalone or as a complement to the standard Sine fitting method.

Dynamic Response of a Metal and a CMC Turbine Blade During a Controlled Rub Event Using a Segmented Shroud

2015 ◽  
Vol 137 (6) ◽  
Author(s):  
Nisrene Langenbrunner ◽  
Matt Weaver ◽  
Michael G. Dunn ◽  
Corso Padova ◽  
Jeffery Barton
Keyword(s):  
Dynamic Response ◽  
Turbine Blade ◽  
Engine Performance ◽  
Ceramic Matrix Composites ◽  
Matrix Composites ◽  
Cooperative Research ◽  
Temperature Capability ◽  
Blade Tip ◽  
The Ohio State University ◽  
Design Speed

Ceramic matrix composites (CMCs) provide several benefits over metal blades including weight and increased temperature capability, and have the potential for increased engine performance by reduction of the cooling flow bled from the compressor and by allowing engines to run at higher turbine inlet temperatures. These CMC blades must be capable of surviving fatigue (high cycle and low cycle), creep, impact, and any tip rub events due to the engine missions or maneuvers that temporarily close blade tip/shroud clearances. As part of a cooperative research program between GE Aviation and the Ohio State University Gas Turbine Laboratory, OSU GTL, the response of a CMC stage 1 low-pressure turbine blade has been compared with the response of an equivalent metal turbine blade using the OSU GTL large spin-pit facility (LSPF) as the test vehicle. Load cells mounted on the casing wall, strain gages mounted on the airfoils, and other instrumentation are used to assess blade tip rub interactions with a 120-deg sector of a representative turbine stationary casing. The intent of this paper is to present the dynamic response of both the CMC and the metal blades with the turbine disk operating at design speed and with representative incursion rates and depths. Casing wear and blade tip wear are both characterized for several types of rub conditions including a light, medium, and heavy rub at room temperature. For each condition, the rub primary dynamic modes have been evaluated, and the corresponding blade tip loads have been calculated. The preliminary results suggest that a CMC blade has similar abilities to a metal blade during a rub event.

Experimental and Computational Investigation of Heat Transfer Effectiveness and Pressure Distribution of a Shrouded Blade Tip Section

2004 ◽  
Author(s):  
Nirm V. Nirmalan ◽  
Jeremy C. Bailey ◽  
Mark E. Braaten
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Pressure Distribution ◽  
Coolant Flow ◽  
Tip Clearance ◽  
Computational Investigation ◽  
Blade Tip ◽  
Flow Through ◽  
Tip Shroud ◽  
Unique Study

An experimental and computational investigation was conducted to study the detailed distribution of heat transfer effectiveness and pressure on an attached tip-shroud of a turbine blade. Temperatures and pressures were measured on the airfoil-side and gap-side surfaces of the shrouded tip in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at gaps varying of 0.25% to 1.67% of blade span and at an airfoil exit Reynolds number of 1.26×106 and Mach number of 0.95. The effect of coolant flow through the radial-cooled airfoil was also studied. The experimental results are compared with a computational model using the commercially available code, CFX. This unique study presents the influence of gap and coolant flow on the pressure distribution and heat transfer effectiveness of an attached tip-shroud surface.

Aeroperformance Measurements for a Fully Cooled High-Pressure Turbine Stage

2013 ◽  
Vol 136 (3) ◽  
Author(s):  
Charles Haldeman ◽  
Michael Dunn ◽  
Randall Mathison ◽  
William Troha ◽  
Timothy Vander Hoek ◽  
...  
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Free Stream ◽  
Pressure Ratio ◽  
Specific Work ◽  
Gas Temperature ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Tip Shroud

A detailed aero performance measurement program utilizing fully cooled engine hardware (high-pressure turbine stage) supplied by Honeywell Aerospace Advanced Technology Engines is described. The primary focus of this work was obtaining relevant aerodynamic data for a small turbine stage operating at a variety of conditions, including changes in operating conditions, geometry, and cooling parameters. The work extraction and the overall stage performance for each of these conditions can be determined using the measured acceleration rate of the turbine disk, the previously measured moment of inertia of the rotating system, and the mass flow through the turbine stage. Measurements were performed for two different values of tip/shroud clearance and two different blade tip configurations. The vane and blade cooling mass flow could be adjusted independently and set to any desired value, including totally off. A wide range of stage pressure ratios, coolant to free stream temperature ratios, and corrected speeds were used during the course of the investigation. A combustor emulator controlled the free stream inlet gas temperature, enabling variation of the temperature ratios and investigation of their effects on aero performance. The influence of the tip/shroud gap is clearly seen in this experiment. Improvements in specific work and efficiency achieved by reducing the tip/shroud clearance depend upon the specific values of stage pressure ratio and corrected speed. The maximum change of 3%–4% occurs at a stage pressure ratio and corrected speed greater than the initial design point intent. The specific work extraction and efficiency for two different blade tip sets (one damaged from a rub and one original) were compared in detail. In general, the tip damage only had a very small effect on the work extraction for comparable conditions. The specific work extraction and efficiency were influenced by the presence of cooling gas and by the temperature of the cooling gas relative to the free stream gas temperature and the metal temperature. These same parameters were influenced by the magnitude of the vane inlet gas total temperature relative to the vane metal temperature and the coolant gas temperature.

Experimental Investigation of Aerodynamic Losses of Different Shapes of a Shrouded Blade Tip Section

2005 ◽  
Author(s):  
Nirm V. Nirmalan ◽  
Jeremy C. Bailey
Keyword(s):  
Experimental Investigation ◽  
Reynolds Number ◽  
Mach Number ◽  
Tip Clearance ◽  
Exit Plane ◽  
Airfoil Surface ◽  
Blade Tip ◽  
Different Shapes ◽  
Tip Shroud ◽  
Aerodynamic Losses

An experimental investigation was conducted to study the effects on aerodynamic losses of different tip shroud shapes of a shrouded turbine blade. Pressures were measured on the airfoil surface near the tip and a plane downstream of the exit plane in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at tip-clearances varying from 0.25% to 1.67% and at an exit Reynolds number of 1.25 × 106 and Mach Number of 0.95. The paper presents the influence of three tip-shroud shapes and five different tip-clearances on the aerodynamic losses.

Aeroperformance Measurements for a Fully Cooled High-Pressure Turbine Stage

2012 ◽  
Author(s):  
Charles Haldeman ◽  
Michael Dunn ◽  
Randall Mathison ◽  
William Troha ◽  
Timothy Vander Hoek ◽  
...  
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Free Stream ◽  
Pressure Ratio ◽  
Specific Work ◽  
Gas Temperature ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Tip Shroud

A detailed aero performance measurement program utilizing fully cooled engine hardware (high-pressure turbine stage) supplied by Honeywell Aerospace Advanced Technology Engines is described. The primary focus of this work was obtaining relevant aerodynamic data for a small turbine stage operating at a variety of conditions, including changes in operating conditions, geometry, and cooling parameters. The work extraction and the overall stage performance for each of these conditions can be determined using the measured acceleration rate of the turbine disk, the previously measured moment of inertia of the rotating system, and the mass flow through the turbine stage. Measurements were performed for two different values of tip/shroud clearance and two different blade tip configurations. The vane and blade cooling mass flow could be adjusted independently and set to any desired value including totally off. A wide range of stage pressure ratios, coolant to freestream temperature ratios, and corrected speeds were used during the course of the investigation. A combustor emulator controlled the free stream inlet gas temperature, enabling variation of the temperature ratios and investigation of their effects on aero performance. The influence of tip/shroud gap is clearly seen in this experiment. Improvements in specific work and efficiency achieved by reducing the tip/shroud clearance depend upon the specific values of stage pressure ratio and corrected speed. The maximum change of 3% to 4% occurs at a stage pressure ratio and corrected speed greater than the initial design point intent. The specific work extraction and efficiency for two different blade tip sets (one damaged from a rub and one original) were compared in detail. In general, the tip damage only had a very small effect on the work extraction for comparable conditions. The specific work extraction and efficiency were influenced by the presence of cooling gas and by the temperature of the cooling gas relative to the free stream gas temperature and the metal temperature. These same parameters were influenced by the magnitude of the vane inlet gas total temperature relative to the vane metal temperature and the coolant gas temperature.

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