Stage Turbine Vane Research Articles

Application of deep-learning method in the conjugate heat transfer optimization of full-coverage film cooling on turbine vanes

• An optimization method using the conjugate heat transfer CFD and the cGAN model is developed for the arrangement of film holes on turbine blades. • The optimized scattered arrangements of film holes can improve the uniformity of wall temperature under heterogeneous thermal loads. • Sampling principles of cGAN model are suggested for the prediction of complex phenomena including backflow. Cooling design optimization with complex nonuniform heat loads is a typical challenge in the development of new generation gas turbines. Combined inlet hot streaks and swirls caused by lean-burn combustor force film cooling to attenuate uneven heat loads on the blade surface, especially in the spanwise direction. Besides, the complex coupling relationship among hundreds of design variables of film cooling arrangement hinders the development of the optimal design. The deep learning model shows a strong fitting ability when dealing with high-dimensional nonlinear problems, which could fit the mapping relationship between design variables and the temperature field. In this paper, a turbine cooling design optimization methodology based on conjugate heat transfer (CHT) simulation and conditional generative adversarial network (cGAN) is developed, and the film cooling design of the 1st stage turbine vanes is optimized through the multi-objective genetic algorithm (MOGA). An initial sample containing 96 cases is constructed by CHT computational fluid dynamics (CFD) simulation with inlet hot streaks and swirls. Based on the initial sample, a cGAN model that predicts the temperature distribution of the vane surface is trained and tested. The film hole arrangement of the vane surface is described by a 276-bit binary optimization variable. The 5% maximum temperature predicted by cGAN and the regressed coolant mass flow is used as the dual optimization objectives. The optimal film hole arrangements in rows and the scattered film hole arrangements found by MOGA are compared, which shows that the optimal scattered arrangements perform better due to well adaptability to the nonuniform thermal load in the spanwise direction. The impact of sample size and sample selection on the performance of the cGAN model is discussed. The retrained cGAN model indicates that a proper abundance of specific samples can improve the prediction of complex coupling phenomena such as backflow.

Metallurgical Failure Analysis of Cracked First Stage Vane of Heavy-Duty Industrial Gas Turbine Engine for Power Generation

Abstract A first stage turbine vane of a heavy-duty gas turbine engine, exhibiting an outward bulge of approximately 0.6 mm and a visible crack in the top coating, was delivered by the client for a detailed destructive failure investigation. The subject component was an experimental part used in a test field engine. The suction side wall of the casting was cut off to reveal a gaping crack on the inside of the aft cavity that was also visible by borescopic inspection. The cooling insert of the aft cavity was clean and did not have any blocked cooling holes. The cooling pattern on the inside wall of the aft cavity was unremarkable. No evidence whatever of impact damage was found on the suction side surface of the subject vane in the cracked region. The hypothesis as to the most probable root cause of the cracking that is still standing is a combination of low wall thickness, non-optimal process parameters for bond coat pre-heat, and potentially high thermal/mechanical loading during engine operation. It appears this is not a wide-spread phenomenon but rather an isolated occurrence.

Dynamics of Coherent Structures and Random Turbulence in Pressure Side Film Cooling on a First Stage Turbine Vane

This paper collects the final results of a combined experimental and numerical investigation on pressure side (PS) film cooling in a high-pressure turbine vane, including two staggered rows of cylindrical holes and a trailing edge cutback, fed by one plenum. Having learned that the scale resolving simulation technique is essential to get reasonable predictions of adiabatic film cooling effectiveness, the stress-blended eddy simulation (SBES) model has been selected as the best among the available hybrid RANS–LES options. Mainstream conditions were limited to low speed and low turbulence intensity due to the need of high temporal and spatial resolution. The choice of one only coolant-to-mainstream mass flow ratio equal to MFR = 1.5% was dictated by the hole discharge: on the one side, mainstream injection into the cooling holes and, on the other side, jet liftoff were avoided to get an effective thermal coverage downstream of the holes. SBES potential was evaluated on the basis of qualitative and quantitative characteristics of the flow along the interface between coolant and mainstream because of their ultimate effect on vane surface temperature. The focus was set on shape and dynamics of coherent structures: SBES provided evidence of shear layer Kelvin–Helmholtz instability and hairpin vortices, downstream of cooling holes, with a Strouhal number (St) of 1.3 and 0.3–0.4, respectively. Simulated vortex shedding in the cutback region was characterized by St of 0.32 to be compared against the measured St value of 0.40.

Stress-Blended Eddy Simulation of Coherent Unsteadiness in Pressure Side Film Cooling Applied to a First Stage Turbine Vane

Within the framework of scale resolving simulation techniques, this paper considers the application of the stress-blended eddy simulation (SBES) model to pressure side (PS) film cooling in a high-pressure turbine nozzle guide vane. The cooling geometry exhibits two rows of film cooling holes and a trailing edge cutback, fed by the same plenum chamber. The blowing conditions investigated were in the range of coolant-to-mainstream mass flow ratio (MFR) from 1% to 2%. The flow regime resembles that in a real engine (exit isentropic Mach number of Ma2is = 0.6), but also low speed conditions (Ma2is = 0.2) were considered for comparison purposes. The predicted results were validated with measurements of surface adiabatic effectiveness and instantaneous off-wall visualizations of the flow field downstream of cooling holes and cutback slot. The focus is on SBES ability of developing shear layer structures, because of their strong influence on velocity field, entrainment mechanisms and, thus, vane surface temperature. Special attention has been paid to the development and dynamics of coherent unsteadiness, since measured values of shedding frequency were also available for validation. SBES provided significant improvement in capturing the unsteady physics of cooling jet-mainstream interaction. The effects of changes in flow regime and blowing conditions on vortex structures were well predicted along the cutback surface. As regards the cooling holes, the high speed condition made it difficult to match the experimental Kelvin–Helmholtz breakdown in the shear layer, in the case of high velocity jets.

Research on cooling effectiveness in stepped slot film cooling vane

As one of the most important developments in air cooling technology for hot parts of the aero-engine, film cooling technology has been widely used. Film cooling hole structure exists mainly in areas that have high temperature, uneven cooling effectiveness issues when in actual use. The first stage turbine vanes of the aero-engine consume the largest portion of cooling air, thereby the research on reducing the amount of cooling air has the greatest potential. A new stepped slot film cooling vane with a high cooling effectiveness and a high cooling uniformity was researched initially. Through numerical methods, the affecting factors of the cooling effectiveness of a vane with the stepped slot film cooling structure were researched. This paper focuses on the cooling effectiveness and the pressure loss in different blowing ratio conditions, then the most reasonable and scientific structure parameter can be obtained by analyzing the results. The results show that 1.0 mm is the optimum slot width and 10.0 is the most reasonable blowing ratio. Under this condition, the vane achieved the best cooling result and the highest cooling effectiveness, and also retained a low pressure loss.

Impact of the Combustor-Turbine Interface Slot Orientation on the Durability of a Nozzle Guide Vane Endwall

The combustor-turbine interface is an essential component in a gas turbine engine as it allows for thermal expansion between the first stage turbine vanes and combustor section. Although not considered as part of the external cooling scheme, leakage flow from the combustor-turbine interface can be utilized as coolant. This paper reports on the effects of orientation of a two-dimensional leakage slot, simulating the combustor-turbine interface, on the net heat flux reduction to a nozzle guide vane endwall. In addition to adiabatic effectiveness and heat transfer measurements, time-resolved, digital particle image velocimetry (TRDPIV) measurements were performed in the vane stagnation plane. Four interface slot orientations of 90 deg, 65 deg, 45 deg, and 30 deg located at 17% axial chord upstream of a first vane in a linear cascade were studied. Results indicate that reducing the slot angle to 45 deg can provide as much as a 137% reduction to the average heat load experienced by the endwall. Velocity measurements indicate the formation of a large leading edge vortex for coolant injected at 90 deg and 65 deg while coolant injected at 45 deg and 30 deg flows along the endwall and washes up the vane surface at the endwall junction.

Showerhead Film Cooling Performance of a Turbine Vane at High Freestream Turbulence in a Transonic Cascade

This paper experimentally investigates the effect of blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane. The vane midspan was instrumented with single-sided platinum thin film gauges to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. Three different exit Mach numbers of Mex = 0.57, 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 9.7 × 105, 1.1 × 106 and 1.5 × 106, respectively—were tested with an inlet free stream turbulence intensity (Tu) of 16% and an integral length scale normalized by vane pitch (Λx/P) of 0.23. A showerhead cooling scheme with five rows of cooling holes was tested at blowing ratios of BR = 0, 1.5, 2.0, and 2.5 and a density ratio of DR = 1.3. Nusselt number and adiabatic film cooling effectiveness distributions were presented on the vane surface over a range of s/C = −0.58 on the pressure side to s/C = 0.72 on the suction side of the vane. The primary effects of coolant injection were to augment the Nusselt number and reduce the adiabatic wall temperature downstream of the injection on the vane surface as compared to no film injection case (BR = 0) at all exit Mach number conditions. In general, an increase in blowing ratio (BR = 1.5 to 2.5) showed noticeable Nusselt number augmentation on pressure surface as compared to suction surface at exit Mach 0.57 and 0.75; however, the Nusselt number augmentation for these blowing ratios was found to be negligible on the vane surface for exit Mach 1.0 case. At exit Mach 1.0, an increase in blowing ratio (BR = 1.5 to 2.5) was observed to have an adverse effect on the adiabatic effectiveness on the pressure surface but had negligible effect on suction surface. The effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition in the region s/C = 0.28 to s/C = 0.45 at all blowing ratio and exit Mach number conditions. An increase in Reynolds number from exit Mach 0.76 to 1.0 increased heat transfer levels on the vane surface at all blowing ratio conditions. A large increase in Reynolds number adversely affected adiabatic effectiveness on the pressure surface at all blowing ratio conditions. On the suction surface, a large increase in Reynolds number also affected adiabatic effectiveness in the favorable pressure gradient and boundary layer transition region.

Heat transfer and film-cooling for the endwall of a first stage turbine vane

Secondary flows that result in turbomachines from inherent pressure gradients in airfoil passages, are the main contributors to aerodynamic losses and high heat transfer to the airfoil endwalls. The endwalls present a challenge to durability engineers in maintaining the integrity of the airfoils. One means of preventing degradation in the turbine is to film-cool components whereby coolant is extracted from the compressor and injected through small cooling holes in the airfoil surfaces. In addition to film-cooling, leakage flows from component interfaces, such as the combustor and turbine, can provide cooling in localized areas but also provide a change to the inlet boundary condition to the passage. This paper presents measurements relevant to the endwall region of a vane, which indicate the importance of considering the inlet flow condition.

Elevated freestream turbulence effects on heat transfer for a gas turbine vane

High freestream turbulence levels have been shown to greatly augment the heat transfer along a gas turbine airfoil, particularly for the first stage nozzle guide vane. For this study, augmentations in convective heat transfer have been measured for a first stage turbine vane in the stagnation region, along the mid-span, and along the platform resulting from an approach freestream turbulence level of 19.5%. In addition to quantifying surface heat transfer, boundary layer measurements have been made to better understand high freestream turbulence effects. Although there are a number of correlations that have been developed for scaling freestream turbulence augmentations to heat transfer, the results of this study indicate that these correlations are not successful in predicting heat transfer for various regions along a turbine vane.

Evaluation of NiCrAlYSi overlay coating on Ni3Al based alloy IC-6 after an engine test

Abstract The second stage turbine vanes of a recently developed high performance jet aero-engine were cast using a Ni 3 Al based alloy IC-6, and then coated with NiCrAlYSi by vacuum arc deposition. After a 379 h engine test, the morphology, microstructure, phase structure and chemical composition of the coating were examined and analyzed by scanning electron microscopy with X-ray energy dispersion spectrometer, X-ray diffraction and electron probe microanalysis techniques. The analytical emphasis was placed on the effect of black spots which appeared on the coating surface after the test on the coating and base alloy IC-6. In addition, the origin of black spots was intensely investigated.The results indicated that the black spots were a kind of high-temperature hot corrosion caused by Na 2 SO 4 melt salt from the combusted gases of the engine, the existence of black spots accelerated the internal oxidation of the coating, but had not impaired the base alloy after 379 h engine operation. It may be summarized that NiCrAlYSi coating can provide an effective protection against oxidation and corrosion of base alloy IC-6 at elevated temperature of the present test.

A study of thermodynamic cycle and system configurations of hydrogen combustion turbines

A hydrogen combustion turbine is powered by steam generated from the internal combustion of hydrogen as a fuel mixed with stoichiometric oxygen. As it is possible to use a closed cycle system, benefits in cycle efficiency and a reduction of environmental pollution effects. Three different closed hydrogen combustion turbine cycles are evaluated. These are the Bottoming reheat cycle (A), the Topping extraction cycle (B) designed by Jericha and Ratzesberger, and the Rankine cycle (C). Calculations have been carried out to investigate the best cycle. This investigation consists of the comparison of thermodynamic efficiency, first stage turbine vane height of the high temperature, high pressure turbine, and maximum operating temperature of the heat exchangers. In these investigations, the component efficiencies are assumed to be the values which are expected to be achieved in the near future. As a result, the thermal efficiency of cycles (A) and (B) is the same value of 61.5%. That of cycle (C), which has the feed water heating with optimized pressure ratio of the intermediate turbine, is 58.8%. Cycle (B) has the largest first stage turbine vane height of the high temperature/high pressure turbine. The larger vane height has an advantage from the point of view of both the manufacturing of the complex cooling passage inside the vane and the turbine aerodynamic efficiency. The maximum operating temperature of the heat exchanger of cycles (A) and (B) is 870 °C, while that of cycle (C) is more than 1000 °C where some problems are anticipated in the feasibility of this heat exchanger. This investigation shows that the Topping extraction cycle (B) is considered to be the best cycle from the point of view of both the thermal efficiency and the feasibility of manufacturing.

Advanced Cooled Turbine Airfoil Aerodynamic Investigation

Abstract : Several connectively cooled trailing edge designs were investigated with the objective of eliminating the need for film cooling on the airfoil suction side. A 43.4% reduced solidity first stage turbine vane having potential application for an advanced Air Force fighter engine was selected for the evaluation. The final design eliminates film cooling on the suction side and uses the wavy criss-cross slot as the cooling scheme for the trailing edge section. The cooling design was incorporated into a cascade test airfoil using the radial wafer fabrication technique. The airfoil was constructed by photoetching the cooling design into the individual wafers, bonding the wafers together and machining the bonded block into the airfoil shape. The airfoil was subsequently evaluated in an airfoil cascade test to determine the aerodynamic and cooling performance. The aerodynamic profile loss of the reduced solidity radial wafer airfoil was reduced 56% relative to a baseline 43.4% reduced solidity configuration with film cooled suction surface and was 30% under the program goal. The wavy criss-cross slot design used in the trailing edge section proved to be an efficient cooling technique, and eliminated the need for suction side film cooling.

External Heat-Transfer Distribution on Film Cooled Turbine Vanes

Experiments were performed to determine the heat transfer to the suction surface of two film cooled first stage turbine vanes using two-row discrete hole cooling. Evaluation of the heat transfer to the film cooled vanes was made in a turbulent environment downstream of an aircraft turbojet engine combustor at moderately high primary gas stream temperatures and pressures. Results of these experiments provided quantitative data on the adiabatic film effectiveness for a number of blowing rates, and show decreased effectiveness levels when compared to measured values for a continuous slot in a wind tunnel environment. In addition the external heat transfer coefficients were obtained based on the difference between the film cooled adiabatic wall temperature and the wall temperature. The presence of film cooling holes increased the heat-transfer coefficients even without blowing. Further increases in the heat-transfer coefficients were measured in the presence of blowing with the largest increases in the region nearest the ejection.

MAR-M 302

Abstract MAR-M 302 is a cast cobalt-base alloy having superior oxidation and thermal shock resistance. It was designed primarily for use as a first stage turbine vane to operate in the temperature range of 1500-2000 F. This datasheet provides information on composition, physical properties, hardness, and tensile properties as well as creep. It also includes information on high temperature performance and corrosion resistance as well as casting, heat treating, machining, and joining. Filing Code: Co-43. Producer or source: Martin Metals Division, Martin Marietta Corporation.

SM-302

Abstract SM 302 is a cast cobalt-base alloy having superior oxidation and thermal shock resistance. It was designed primarily for use as a first stage turbine vane to operate in the temperature range of 1500-2000 F. This datasheet provides information on composition, physical properties, hardness, and tensile properties as well as creep. It also includes information on high temperature performance and corrosion resistance as well as casting, heat treating, machining, and joining. Filing Code: Co-32. Producer or source: Sierra Metals Corporation.