Turbine Rotor Inlet Temperature Research Articles

Turbine Inlet Temperature Measurements in an 8200 KW Gas Turbine Using Water Vapor Emission

Abstract The measurement of turbine inlet temperature is challenging because of high temperatures and complicated physical access, but continuous measurement of the turbine inlet temperature is very important for maximizing turbine efficiency and increasing durability. This paper provides in situ turbine rotor inlet temperature measurements in an 8200 kW operating gas turbine engine. The measurements were obtained using integrated spectral infrared emission from the water vapor of the combustion gases entering the turbine rotor. The method utilizes a sapphire optical fiber to convey the signal from the turbine wall to outside the turbine casing. All components are capable of long-term exposure to the turbine operating conditions. The temperature measurements were obtained at 6 operating conditions between 50% and full load. The turbine rotor inlet temperature temperature was also determined using more than 20 test cell inputs and Solar Turbine's commercial test cell engine model. The two temperatures (measured and modeled) were within 11 K (less than 1%) across the load sweep. Uncertainty calculations suggest that the uncertainty of the measurement can be expected to be ±2.9% within a confidence interval of 95%. The method also yields the nozzle guide vane surface temperature, which was found to increase monotonically with increasing load.

Investigation of a Rotor Blade With Tip Cooling Subject to a Nonuniform Temperature Profile

Abstract One of the challenges in the design of a high-pressure turbine blade is that a considerable amount of cooling is required so that the blade can survive high temperature levels during engine operation. Another challenge is that the addition of cooling should not adversely affect blade aerodynamic performance. The typical flat tips used in designs have evolved into squealer form that implements rims on the tip, which has been reported in several studies to achieve better heat transfer characteristics as well as to decrease pressure losses at the tip. This paper demonstrates a numerical study focusing on a squealer turbine blade tip that is operating in a turbine environment matching the typical design ratios of pressure, temperature, and coolant blowing. The blades rotate at a realistic rpm and are subjected to a turbine rotor inlet temperature profile that has a nonuniform shape. For comparison, a uniform profile is also considered as it is typically used in computational studies for simplicity. The effect of tip cooling is investigated by implementing seven holes on the tip near the blade pressure side. Results confirm that the temperature profile nonuniformity and the addition of cooling are the drivers for loss generation, and they further increase losses when combined. Temperature profile migration is not pronounced with a uniform profile but shows distinct features with a nonuniform profile for which hot gas migration toward the blade pressure side is observed. The blade tip also receives higher coolant coverage when subject to the nonuniform profile.

Advanced Cooling in Gas Turbines 2016 Max Jakob Memorial Award Paper

Gas turbines have been extensively used for aircraft engine propulsion, land-based power generation, and industrial applications. Power output and thermal efficiency of gas turbines increase with increasing turbine rotor inlet temperatures (RIT). Currently, advanced gas turbines operate at turbine RIT around 1700 °C far higher than the yielding point of the blade material temperature about 1200 °C. Therefore, turbine rotor blades need to be cooled by 3–5% of high-pressure compressor air around 700 °C. To design an efficient turbine blade cooling system, it is critical to have a thorough understanding of gas turbine heat transfer characteristics within complex three-dimensional (3D) unsteady high-turbulence flow conditions. Moreover, recent research trend focuses on aircraft gas turbines that operate at even higher RIT up to 2000 °C with a limited amount of cooling air, and land-based power generation gas turbines (including 300–400 MW combined cycles with 60% efficiency) burn alternative syngas fuels with higher heat load to turbine components. It is important to understand gas turbine heat transfer problems with efficient cooling strategies under new harsh working environments. Advanced cooling technology and durable thermal barrier coatings (TBCs) play most critical roles for development of new-generation high-efficiency gas turbines with near-zero emissions for safe and long-life operation. This paper reviews basic gas turbine heat transfer issues with advanced cooling technologies and documents important relevant papers for future research references.

Effect of V-shaped Ribs on Internal Cooling of Gas Turbine Blades

Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperature. The rotor inlet temperatures in most gas turbines are far higher than the melting point of the blade material. Hence the turbine blades need to be cooled. In this work, simulations were carried out with the leading edge of gas turbine blade being internally cooled by coolant passages with V-shaped ribs at angles of 30°, 45° or 60° and at three aspect ratios (1:1, 1:2 and 2:3). The trailing edge of the blade was cooled by cylindrical and triangular pin-fin perforations in staggered and inline arrangements. Numerical analyses were carried out for each configuration of the cooling passages. The best cooling passages for leading edge and trailing edge were deduced by comparing the results of these analyses. It was found that using V-shaped ribs and fins induces a swirling flow, which in turn increases the velocity gradient and hence produces an improvement in heat transfer. The results show that under real time flow conditions, the application of V-shaped ribs and pin-fin perforations is a very promising technique for improving blade life.

CFD ANALYSIS ON RADIALCOOLING OF GAS TURBINE BLADE

Gas turbines are extensively used for air craft propulsion, land based power generation and industrial applications. Thermal efficiency of gas turbine improved by increasing turbine rotor inlet temperature. The current rotor inlet temperature in advanced gas turbine is for above the melting point of blade material. A sophisticated cooling scheme must be developed for continuous safe operation of gas turbines with high performance. Gas turbines are cooled externally and internally. Several methods have been suggested for the cooling of blades and vanes. The techniques that involve to cool the blades and vanes by using cooling methods is to have radial holes to pass high velocity cooling air along the blade span. In this thesis, a turbine blade is designed and modelled in CATIA v5and ICEM CFDsoftware. The turbine blades are designed using cooling holes. The turbine blade is designed with 12 holes. CFD analysis is done to determine the pressure distribution, velocity, temperature distribution and heat transfer rate by applying the inlet velocities. Thermaland Structural analysis is done to determine the heat transfer rates and strength of the blade. The present used material for blade is chromium steel. In this thesis, itis replaced with Titanium aluminium alloy. The better material for blade is analyzed

Investigation of parameters affecting exergy and emission performance of basic and intercooled gas turbine cycles

In this article an attempt has been made to analyze the affect of various cycle operating parameters, compressor-pressure-ratio, TRIT (turbine-rotor-inlet-temperature) combustor-primary-zone-temperature, equivalence-ratio, and residence-time on thermodynamic as well as emission performance of the BGT (basic-gas-turbine) and IcGT (intercooled-gas-turbine) cycles on a comparative basis. Thermodynamic assessment of the proposed cycles, shows that rational efficiency of the IcGT (intercooled-gas-turbine cycle) to be 8.39% higher than the BGT (basic-gas-turbine cycle). Overall exergy destruction within the cycles has been found to the 4.42% lower for IcGT cycle as compared to BGT cycle. It has also been observed that the IcGT cycle delivers higher gas turbine specific work and thermal efficiency in comparison to the BGT cycle for the same compressor-pressure-ratio and TRIT. Emission assessment shows that at fixed value of equivalence-ratio and residence-time, NOx emission is higher at higher values of compressor-pressure-ratio (rp,c) for both cycles. The mass of NOx and UHC (unburnt-hydrocarbon) emission increases with increase in equivalence-ratio, whereas CO (carbon-monoxide) emission decreases with increase in equivalence-ratio. Emission performance maps show lower quantum of NOx and CO emission for the IcGT cycle. UHC emission is higher in case of IcGT cycle due to lower combustor inlet air temperature. Overall, both thermodynamic and emission performance of IcGT cycle is superior.

Experimental investigation of heat transfer and flow using V and broken V ribs within gas turbine blade cooling passage

Gas turbines are extensively used for aircraft propulsion, land-based power generation, and various industrial applications. With an increase in turbine rotor inlet temperatures, developments in innovative gas turbine cooling technology enhance the efficiency and power output; these advancements of turbine cooling have allowed engine designs to exceed normal material temperature limits. For internal cooling design, techniques for heat extraction from the surfaces exposed to hot stream of gas are based on an increase in the heat transfer areas and on the promotion of turbulence of the cooling flow. In this study, an improvement in performance is obtained by casting repeated continuous V- and broken V-shaped ribs on one side of the two pass square channels into the core of the blade. A detailed experimental investigation is done for two pass square channels with a 180° turn. Detailed heat transfer distribution occurring in the ribbed passage is reported for a steady state experiment. Four different combinations of 60° V- and broken 60° V-ribs in a channel are considered. A series of thermocouples are used to obtain the temperature on the channel surface and local heat transfer coefficients are obtained for Reynolds numbers 16,000, 56,000 and 85,000 within the turbulent flow regime. Area averaged data are calculated in order to compare the overall performance of the tested ribbed surface and to evaluate the degree of heat transfer enhancement induced by the rib. Flow within the channels is characterized by heat transfer enhancing ribs, bends, rotation and buoyancy effects. A series of experimental measurements is performed to predict the overall performance of the channel. This paper presents an attempt to collect information about the Nusselt number, the pressure drop and the overall performance of the eight different ribbed ducts at the specified Reynolds number. The main contribution of this study is to evaluate the best combination of rib arrangements throughout the two pass cooling channels. After a series of experiments, it can be concluded that the distribution of peaks in heat transfer in the case of inlet V and outlet inverted V is high. The overall performances for broken ribs are higher compared with the continuous ribs in two-pass cooling channels.

Fundamental Gas Turbine Heat Transfer

Gas turbines are used for aircraft propulsion and land-based power generation or industrial applications. Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperatures (RIT). Current advanced gas turbine engines operate at turbine RIT (1700 °C) far higher than the melting point of the blade material (1000 °C); therefore, turbine blades are cooled by compressor discharge air (700 °C). To design an efficient cooling system, it is a great need to increase the understanding of gas turbine heat transfer behaviors within complex 3D high-turbulence unsteady engine-flow environments. Moreover, recent research focuses on aircraft gas turbines operating at even higher RIT with limited cooling air and land-based gas turbines burn coal-gasified fuels with a higher heat load. It is important to understand and solve gas turbine heat transfer problems under new harsh working environments. The advanced cooling technology and durable thermal barrier coatings play critical roles for the development of advanced gas turbines with near zero emissions for safe and long-life operation. This paper reviews fundamental gas turbine heat transfer research topics and documents important relevant papers for future research.

Durability of Oxide/Oxide Ceramic Matrix Composites in Gas Turbine Combustors

An integrated creep rupture strength degradation and water vapor degradation model for gas turbine oxide-based ceramic matrix composite (CMC) combustor liners was expanded with heat transfer computations to establish the maximum turbine rotor inlet temperature (TRIT) for gas turbines with 10:1 pressure ratio. Recession rates and average CMC operating temperatures were calculated for an existing baseline N720/A (N720/Al2O3) CMC combustor liner system with and without protective Al2O3 friable graded insulation (FGI) for 30,000-h liner service life. The potential for increasing TRIT by Y3Al5O12 (YAG) substitution for the fiber, matrix, and FGI constituents of the CMC system was explored, because of the known superior creep and water vapor degradation resistance of YAG compared to Al2O3. It was predicted that uncoated N720/A can be used as a combustor liner material up to a TRIT of ∼1200 °C, offering no TRIT advantage over a conventional metal + thermal barrier coating (TBC) combustor liner. A similar conclusion was previously reached for a SiC/SiC CMC liner with barium strontium aluminum silicate (BSAS)-type environmental barrier coating (EBC). The existing N720/A + Al2O3 FGI combustor liner system can be used at a maximum TRIT of ∼1350 °C, a TRIT increase over metal + TBC, and uncoated N720/A of ∼150 °C. Replacing the Al2O3 with YAG is predicted to increase the maximum allowable TRIT. Substitution of the fiber or matrix in N720/A increases TRIT by ∼100 °C. A YAG FGI improves the TRIT of the N720/A + Al2O3 FGI by ∼50 °C, enabling a TRIT of ∼1400 °C, similar to that predicted for SiC/SiC CMCs with protective rare earth monosilicate EBCs.

Tip Clearance Effects on Inlet Hot Streak Migration Characteristics in High Pressure Stage of a Vaneless Counter-Rotating Turbine

In this paper, three-dimensional multiblade row unsteady Navier–Stokes simulations at a hot streak temperature ratio of 2.0 have been performed to reveal the effects of rotor tip clearance on the inlet hot streak migration characteristics in high pressure stage of a vaneless counter-rotating turbine. The numerical results indicate that the migration characteristics of the hot streak in the high pressure turbine rotor are dominated by the combined effects of secondary flow, buoyancy, and leakage flow in the rotor tip clearance. The leakage flow trends to drive the hotter fluid toward the blade tip on the pressure surface and to the hub on the suction surface. Under the effect of the leakage flow, even partial hotter fluid near the pressure surface is also driven to the rotor suction surface through the tip clearance. Compared with the case without rotor tip clearance, the heat load of the high pressure turbine rotor is intensified due to the effects of the leakage flow. And the results indicate that the leakage flow effects trend to increase the low pressure turbine rotor inlet temperature at the tip region. The air flow with higher temperature at the tip region of the low pressure turbine rotor inlet will affect the flow and heat transfer characteristics in the downstream low pressure turbine.

Trailing-Edge Cooling for Gas Turbines

The trailing-edge section of modern high-pressure turbine airfoils is an area that requires a high degree of attention from turbine performance and durability standpoints. Aerodynamic loss near the trailing edge includes expansion waves, normal shocks, and wake shedding. Thermal issues associated with trailing edge are also very complex and challenging. To maintain effective cooling ensuring metal temperature below design limit is particularly difficult, as it needs to be implemented in a relatively small area of the airfoil. To date, little effort has been devoted to advancing the fundamental understanding of the thermal characteristics in airfoil trailing-edge regions. Described in this paper are the procedures leading to closed-form, analytical solutions for temperature profile for four most representative trailing-edge configurations. The configurations studied are 1) solid wedge shape without discharge, 2) wedge with slot discharge, 3) wedge with discrete-hole discharge, and 4) wedge with pressure-side cutback slot discharge. Comparison among these four cases is made primarily in the context of airfoil metal temperature and resulting cooling effectiveness. Further discussed in the paper are the overall and detail design parameters for preferred trailing-edge cooling configurations as they affect turbine airfoil performance and durability. Also described in this treatment is a current experimental investigation of heat transfer over a trailing-edge configuration. The trailing edge is preceded with an internal cooling channel of pedestal array. The pedestal array consists of both circular pin fins and oblong shaped features or exit teardrops. Downstream to the pedestal array, the trailing edge exits in a pressure side cutback partitioned by the oblong-shaped teardrops. The local heat-transfer coefficient over the entire wetted surface in the internal cooling chamber has been determined using a “hybrid” measurement technique based on transient liquid crystal imaging. The hybrid technique employs the transient conduction model in a semi-infinite solid for resolving the heat-transfer coefficient on the end-wall surface uncovered by the pedestals. The heat-transfer coefficient over a pedestal can be resolved by the lumped capacitance method with an assumption of low Biot number. The overall heat transfer for both the pedestals and end-walls combined shows a significant enhancement compared to the case with thermally developed smooth channel. Near the most downstream section of the suction side, the land, caused by pressure side cutback, is exposed to the stream mixed with hot gas and discharged coolant. Both the adiabatic effectiveness and heat-transfer coefficient on the land section are characterized by using this hybrid liquid-crystal technique. Dr. Cunha is a Principal Engineer at Pratt & Whitney responsible for advanced turbine designs of commercial and military engine programs. Dr. Cunha is also responsible for establishing technical direction for a consortium of Universities as well as in-house research programs in the area of state-of-the-art cooling technologies suitable for use with extremely aggressive gas turbine rotor inlet temperatures. Dr. Cunha’s professional career spans a period of 27 years with turbine original equipment manufacturers including General Electric and Siemens. Dr. Cunha has authored numerous technical papers and gas turbine design standards. Dr. Cunha holds 21 U.S. patents in turbomachinery and is presently a member of the IGTI-ASME heat transfer committee.

Characterization of First‐Stage Silicon Nitride Components After Exposure to an Industrial Gas Turbine

This paper provides a summary of recent efforts undertaken to examine the mechanical properties and microstructural stability of first‐stage silicon nitride blades and nozzles after field testing in an industrial gas turbine. Two commercially available silicon nitrides, i.e., Kyocera SN282 vanes and SN281 blades, were successfully evaluated in the 100 h final phase engine test at Solar Turbines Incorporated. The turbine rotor inlet temperature was increased from 1010° to 1121°C at 100% speed during the engine test with efficiency increasing from 28.8% to 30.1%. Results of scanning electron microscopy showed that apparent materials recession still occurred during the 100 h engine test, especially in the leading and trailing edge regions where the gas pressure or velocity was the highest. The apparent material recession of the airfoils resulted from the volatilization of the normally protective silica layer, evidenced by the increased surface roughness and porous Lu2Si2O7 surface layer features. On the other hand, mechanical results generated using a ball‐on‐ring test technique showed that little strength degradation was measured after the 100 h engine test.

Ceramic Stationary Gas Turbine Development Program—Fifth Annual Summary

A program is being performed under the sponsorship of the United States Department of Energy, Office of Industrial Technologies, to improve the performance of stationary gas turbines in cogeneration through the selective replacement of metallic hot section components with ceramic parts. The program focuses on design, fabrication, and testing of ceramic components, generating a materials properties data base, and applying life prediction and nondestructive evaluation (NDE). The development program is being performed by a team led by Solar Turbines Incorporated, and which includes suppliers of ceramic components, U.S. research laboratories, and an industrial cogeneration end user. The Solar Centaur 50S engine was selected for the development program. The program goals included an increase in the turbine rotor inlet temperature (TRIT) from 1010°C (1850°F) to 1121°C (2050°F), accompanied by increases in thermal efficiency and output power. The performance improvements are attributable to the increase in TRIT and the reduction in cooling air requirements for the ceramic parts. The ceramic liners are also expected to lower the emissions of NOx and CO. Under the program uncooled ceramic blades and nozzles have been inserted for currently cooled metal components in the first stage of the gas producer turbine. The louvre-cooled metal combustor liners have been replaced with uncooled continuous-fiber reinforced ceramic composite (CFCC) liners. Modifications have been made to the engine hot section to accommodate the ceramic parts. To date, all first generation designs have been completed. Ceramic components have been fabricated, and are being tested in rigs and in the Centaur 50S engine. Field testing at an industrial co-generation site was started in May, 1997. This paper will provide an update of the development work and details of engine testing of ceramic components under the program.

Recent Development in Turbine Blade Film Cooling

Gas turbines are extensively used for aircraft propulsion, land‐based power generation, and industrial applications. Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperature (RIT). The current RIT level in advanced gas turbines is far above the .melting point of the blade material. Therefore, along with high temperature material development, a sophisticated cooling scheme must be developed for continuous safe operation of gas turbines with high performance. Gas turbine blades are cooled internally and externally. This paper focuses on external blade cooling or so‐called film cooling. In film cooling, relatively cool air is injected from the inside of the blade to the outside surface which forms a protective layer between the blade surface and hot gas streams. Performance of film cooling primarily depends on the coolant to mainstream pressure ratio, temperature ratio, and film hole location and geometry under representative engine flow conditions. In the past number of years there has been considerable progress in turbine film cooling research and this paper is limited to review a few selected publications to reflect recent development in turbine blade film cooling.

Thermodynamic Evaluation of Gas Turbine Cogeneration Cycles: Part I—Heat Balance Method Analysis

This paper presents a heat balance method of evaluating various open-cycle gas turbines and heat recovery systems based on the first law of thermodynamics. A useful graphic solution is presented that can be readily applied to various gas turbine cogeneration configurations. An analysis of seven commercially available gas turbines is made showing the effect of pressure ratio, exhaust temperature, intercooling, regeneration, and turbine rotor inlet temperature in regard to power output, heat recovery, and overall cycle efficiency. The method presented can be readily programmed in a computer, for any given gaseous or liquid fuel, to yield accurate evaluations. An X–Y plotter can be utilized to present the results.