Effects of Particle Size, Gas Temperature and Metal Temperature on High Pressure Turbine Deposition in Land Based Gas Turbines From Various Synfuels

2007 ◽  
Author(s):  
Jared M. Crosby ◽  
Scott Lewis ◽  
Jeffrey P. Bons ◽  
Weiguo Ai ◽  
Thomas H. Fletcher
Keyword(s):  
Particle Size ◽  
High Pressure ◽  
Coal Ash ◽  
Capture Efficiency ◽  
Metal Temperature ◽  
Test Series ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
The Third ◽  
Power Turbine

Four series of tests were performed in an accelerated deposition test facility to study the independent effects of particle size, gas temperature, and metal temperature on ash deposits from two candidate power turbine synfuels. The facility matches the gas temperature and velocity of modern first stage high pressure turbine vanes while accelerating the deposition process. This is done by matching the net throughput of particulate out of the combustor with that experienced by a modern power turbine. In the first series of tests, four different size particles were studied by seeding a natural-gas combustor with finely-ground coal ash particulate. The entrained ash particles were accelerated to a combustor exit flow Mach number of 0.25 before impinging on a thermal barrier coated (TBC) target coupon at 1183°C. Particle size was found to have a significant effect on capture efficiency with larger particles causing significant TBC spallation during a 4-hour accelerated test. In the second series of tests, different gas temperatures were studied while the facility maintained a constant exit velocity of 170m/s (Mach = 0.23–0.26). Coal ash with a mass mean diameter of 3 μm was used. Particle deposition rate was found to decrease with decreasing gas temperature. The threshold gas temperature for deposition was approximately 960°C. In the third and fourth test series impingement cooling was applied to the backside of the target coupon to simulate internal vane cooling. Ground coal and petcoke ash particulates were used for the two tests respectively. Capture efficiency was reduced with increasing massflow of coolant air, however at low levels of cooling the deposits attached more tenaciously to the TBC layer. Post exposure analyses of the third test series (scanning electron microscopy and x-ray spectroscopy) show decreasing TBC damage with increased cooling levels. Implications for the power generation goal of fuel flexibility are discussed.

Effects of Temperature and Particle Size on Deposition in Land Based Turbines

2008 ◽  
Vol 130 (5) ◽  
Author(s):  
Jared M. Crosby ◽  
Scott Lewis ◽  
Jeffrey P. Bons ◽  
Weiguo Ai ◽  
Thomas H. Fletcher
Keyword(s):  
Particle Size ◽  
Particle Deposition ◽  
Deposition Process ◽  
Capture Efficiency ◽  
Test Series ◽  
Test Facility ◽  
Back Side ◽  
Gas Temperature ◽  
The Third ◽  
Barrier Coating

Four series of tests were performed in an accelerated deposition test facility to study the independent effects of particle size, gas temperature, and metal temperature on ash deposits from two candidate power turbine synfuels (coal and petcoke). The facility matches the gas temperature and velocity of modern first stage high pressure turbine vanes while accelerating the deposition process. Particle size was found to have a significant effect on capture efficiency with larger particles causing significant thermal barrier coating (TBC) spallation during a 4 h accelerated test. In the second series of tests, particle deposition rate was found to decrease with decreasing gas temperature. The threshold gas temperature for deposition was approximately 960°C. In the third and fourth test series, impingement cooling was applied to the back side of the target coupon to simulate internal vane cooling. Capture efficiency was reduced with increasing mass flow of coolant air; however, at low levels of cooling, the deposits attached more tenaciously to the TBC layer. Postexposure analyses of the third test series (scanning electron microscopy and X-ray spectroscopy) show decreasing TBC damage with increased cooling levels.

High Pressure Turbine Deposition in Land Based Gas Turbines From Various Synfuels

2005 ◽  
Author(s):  
Jeffrey P. Bons ◽  
Jared Crosby ◽  
James E. Wammack ◽  
Brook I. Bentley ◽  
Thomas H. Fletcher
Keyword(s):  
High Pressure ◽  
Gas Turbines ◽  
Test Facility ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
Deposit Thickness ◽  
Barrier Coating ◽  
Power Turbine ◽  
Fuel Ash ◽  
Deposition Test

Ash deposits from four candidate power turbine synfuels were studied in an accelerated deposition test facility. The facility matches the gas temperature and velocity of modern first stage high pressure turbine vanes. A natural-gas combustor was seeded with finely-ground fuel ash particulate from four different fuels: straw, sawdust, coal, and petroleum coke. The entrained ash particles were accelerated to a combustor exit flow Mach number of 0.31 before impinging on a thermal barrier coating (TBC) target coupon at 1150°C. Post exposure analyses included surface topography, scanning electron microscopy, and x-ray spectroscopy. Due to significant differences in the chemical composition of the various fuel ash samples, deposit thickness and structure vary considerably for each fuel. Biomass products (e.g. sawdust and straw) are significantly less prone to deposition than coal and petcoke for the same particle loading conditions. In a test simulating one turbine operating year at a moderate particulate loading of 0.02 parts per million by weight, deposit thickness from coal and petcoke ash exceeded 1 mm and 2 mm respectively. These large deposits from coal and petcoke were found to detach readily from the turbine material with thermal cycling and handling. The smaller biomass deposit samples showed greater tenacity in adhering to the TBC surface. In all cases, corrosive elements (e.g. Na, K, V, Cl, S) were found to penetrate the TBC layer during the accelerated deposition test. Implications for the power generation goal of fuel flexibility are discussed.

High-Pressure Turbine Deposition in Land-Based Gas Turbines From Various Synfuels

2005 ◽  
Vol 129 (1) ◽  
pp. 135-143 ◽  
Author(s):  
Jeffrey P. Bons ◽  
Jared Crosby ◽  
James E. Wammack ◽  
Brook I. Bentley ◽  
Thomas H. Fletcher
Keyword(s):  
High Pressure ◽  
Gas Turbines ◽  
Test Facility ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
Deposit Thickness ◽  
Barrier Coating ◽  
Power Turbine ◽  
Fuel Ash ◽  
Deposition Test

Ash deposits from four candidate power turbine synfuels were studied in an accelerated deposition test facility. The facility matches the gas temperature and velocity of modern first-stage high-pressure turbine vanes. A natural gas combustor was seeded with finely ground fuel ash particulate from four different fuels: straw, sawdust, coal, and petroleum coke. The entrained ash particles were accelerated to a combustor exit flow Mach number of 0.31 before impinging on a thermal barrier coating (TBC) target coupon at 1150°C. Postexposure analyses included surface topography, scanning electron microscopy, and x-ray spectroscopy. Due to significant differences in the chemical composition of the various fuel ash samples, deposit thickness and structure vary considerably for each fuel. Biomass products (e.g., sawdust and straw) are significantly less prone to deposition than coal and petcoke for the same particle loading conditions. In a test simulating one turbine operating year at a moderate particulate loading of 0.02 parts per million by weight, deposit thickness from coal and petcoke ash exceeded 1 and 2mm, respectively. These large deposits from coal and petcoke were found to detach readily from the turbine material with thermal cycling and handling. The smaller biomass deposit samples showed greater tenacity in adhering to the TBC surface. In all cases, corrosive elements (e.g., Na, K, V, Cl, S) were found to penetrate the TBC layer during the accelerated deposition test. Implications for the power generation goal of fuel flexibility are discussed.

Effect of Particle Size and Trench Configuration on Deposition From Fine Coal Flyash Near Film Cooling Holes

2009 ◽  
Author(s):  
Weiguo Ai ◽  
Robert G. Laycock ◽  
Devin S. Rappleye ◽  
Thomas H. Fletcher ◽  
Jeffrey P. Bons
Keyword(s):  
Particle Size ◽  
Film Cooling ◽  
Collection Efficiency ◽  
Particle Surface ◽  
Coal Ash ◽  
Capture Efficiency ◽  
Deposition Pattern ◽  
Fine Coal ◽  
Temperature Maps ◽  
Film Cooling Holes

Particulate deposition experiments were performed in a turbine accelerated deposition facility to examine the effects of flyash particle size and trench configuration on deposits near film cooling holes. Deposition on two bare metal Inconel coupons was studied, with hole spacings (s/d) of 3.375 and 4.5. Two sizes of sub-bituminous coal ash particles were used, with mass mean diameter of 4 and 13 microns, respectively. The effect of a cooling trench at the exit of the cooling holes was also examined in this deposition facility. Experiments were performed at different angles of impaction. Particles were accelerated to a combustor exit flow Mach number of 0.25 and heated to 1183°C before impinging on a target coupon. The particle loading in the 1-hr tests was 160 ppmw. Blowing ratios were varied in these experiments from 0 to 4.0. Particle surface temperature maps were measured using two-color pyrometry based on RGB signals from a camera. Deposits generated from finer particles were observed to stick to the surface more tenaciously than larger particles. The capture efficiency measured for the small particles was lower than for the larger particles, especially at low blowing ratios. However, the finer particles exhibited a greater variation in deposition pattern as a function of hole spacing than seen with larger particles. The effect of trench configuration on deposition was examined by performing deposition tests with and without the trench for the same hole spacing and blowing ratio. The effects of trench configuration on capture efficiency, deposition pattern, and surface topography are reported. Deposition experiments at impingement angles from 45° to 15° showed changes in both deposit thickness and temperature. The trench increased cooling effectiveness, but did not change the particulate collection efficiency because the trench acted as a particulate collector.

Unsteady Half-Annulus Computational Fluid Dynamics Calculations of Thermal Migration Through a Cooled 2.5 Stage High-Pressure Turbine

2014 ◽  
Vol 136 (8) ◽  
Author(s):  
James A. Tallman
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
High Pressure ◽  
Cfd Simulation ◽  
Domain Size ◽  
Steady State Solution ◽  
Rotor Blades ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
Model Domain

This paper presents an industrial perspective on the potential use of multiple-airfoil row unsteady computational fluid dynamics (CFD) calculations in high-pressure turbine design cycles. A sliding-mesh unsteady CFD simulation is performed for a high-pressure turbine section of a modern aviation engine at conditions representative of engine take-off. The turbine consists of two stages plus a center-frame strut upstream of the low-pressure turbine. The airfoil counts per row are such that a half-annulus model domain must be simulated for periodicity. The total model domain size is 170 MM computational grid points and the solution requires approximately nine days of clock time on 6288 processing cores of a Cray XE6 supercomputer. Airfoil and endwall cooling flows are modeled via source term additions to the flow. The endwall flowpath cavities and their purge/leakage flows are resolved in the computational meshes to an extent. The time-averaged temperature profile solution is compared with static rake data taken in engine tests. The unsteady solution shows a considerable improvement in agreement with the rake data, compared with a steady-state solution using circumferential mixing planes. Passage-to-passage variations in the gas temperature prediction are present in the 2nd stage, due to nonperiodic alignment between the nozzle vanes and rotor blades. These passage-to-passage differences are quantified and contrasted.

A Generalized Procedure to Estimate Matching of Power Turbine With Aeroderivative Gas Generator at High Speed Settings

2005 ◽  
Author(s):  
Inam U. Haq
Keyword(s):  
High Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Gas Generator ◽  
High Pressure Turbine ◽  
Data Manipulation ◽  
Turbine Inlet ◽  
Engine Design ◽  
Power Turbine ◽  
Set Up

This paper encapsulates generalized considerations of power turbine matching with aeroderivative gas generator at high power settings. A computation route is set up to estimate the magnitude of the desired parameters from design point knowledge of a gas generator. Then, a method is delineated to verify matching of power turbine inlet nozzle area with exhaust of gas generator by measuring tangible tested parameters. Data manipulation revealed that there exists a favorable correlation between pressure ratio of high pressure turbine and gas generator speed that may directly reflect the influence of physical area change of power turbine inlet nozzle area. A practical example is presented to demonstrate the procedure. From engine design to retirement, the generalized considerations may be applied on several occasions where question of matching may become important and require explanation for performance and financial justifications. Some generalized rules of matching are condensed and their applications are suggested.

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.

Ceramic Matrix Composites for High Pressure Turbine Vanes

2014 ◽  
Author(s):  
Robert J. Boyle ◽  
Ankur H. Parikh ◽  
Vinod K. Nagpal ◽  
Michael C. Halbig ◽  
James A. DiCarlo
Keyword(s):  
High Pressure ◽  
Ceramic Matrix Composite ◽  
Ceramic Matrix Composites ◽  
Ceramic Matrix ◽  
Thermal Loads ◽  
Matrix Composites ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Barrier Coating

Through thickness, hoop, and spanwise component stresses were calculated for two Ceramic Matrix Composite (CMC) vane configurations. The analyses are for the first stage vane of a High Pressure Turbine. One configuration is for a vane with trailing edge ejection, and the other has no trailing edge ejection. The effects of analyzing separate pressure and thermal loads, as well as combining these loads, are examined. For the case without trailing edge ejection the effects of variations in the stiffness modulus are given. Results are discussed for the midspan region as well as for the entire span. Pressure loads were determined assuming a mainstream gas and coolant pressure of 50 atm. Thermal loads were determined assuming a gas temperature of 2141°K(3394°F), and a maximum Environmental Barrier Coating temperature of 1756°K(2700°F). The desired maximum CMC temperature was 1589°C(2400°F).

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.

Unsteady, Half-Annulus CFD Calculations of Thermal Migration Through a Cooled, 2.5 Stage High-Pressure Turbine

2013 ◽  
Author(s):  
James A. Tallman
Keyword(s):  
High Pressure ◽  
Cfd Simulation ◽  
Domain Size ◽  
Steady State Solution ◽  
Rotor Blades ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
Model Domain ◽  
Grid Points ◽  
Two Stages

This paper presents an industrial perspective on the potential use of multiple-airfoil row, unsteady CFD calculations in high-pressure turbine design cycles. A sliding-mesh unsteady CFD simulation is performed for a high-pressure turbine section of a modern aviation engine at conditions representative of engine take-off. The turbine consists of two stages plus a center-frame strut upstream of the low-pressure turbine. The airfoil counts per row are such that a half-annulus model domain must be simulated for periodicity. The total model domain size is 170MM computational grid points, and the solution requires approximately 9 days of clock time on 6,288 processing cores of a Cray XE6 supercomputer. Airfoil and endwall cooling flows are modeled via source term additions to the flow. The endwall flowpath cavities and their purge/leakage flows are resolved in the computational meshes to an extent. The time-averaged temperature profile solution is compared with static rake data taken in engine tests. The unsteady solution shows a considerable improvement in agreement with the rake data, compared with a steady-state solution using circumferential mixing planes. Passage-to-passage variations in gas temperature prediction are present in the 2nd stage, due to non-periodic alignment between the nozzle vanes and rotor blades. These passage-to-passage differences are quantified and contrasted.

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