The outlet temperature of combustor is commonly monitored by thermocouples at the turbine exhaust. In order to establish the corresponding relationship between the temperature measured by each thermocouple and the working state of each burner, the azimuthal migration of the combustion hot/cold streaks in the multi-stage turbines needs to be quantified. Experiments to measure this migration have high cost and considerable error. It is also difficult to quantify the migration under multiple working conditions. Three-dimensional full-annulus unsteady simulation can obtain this migration. But the unsteady simulation of a single working condition could take several weeks, which is too expensive for engineering usage. A method named Steady-state Computation of Azimuthal Migration (SCAM) is proposed in this paper. By establishing and solving the transport equation of the migration angle, the azimuthal migration of hot/cold streaks can be predicted by steady-state numerical simulation using the mixing plane at rotor-stator interface. The migration computed by this method is compared with the full-annulus unsteady simulation results in multiple working conditions. The results of SCAM method show good agreement with full-annulus simulations, while costing only 0.01% of the CPU hours. It is also found that the error of SCAM is mainly caused by the fixed boundary value at coolant source terms. The optimal spanwise location of the thermocouples at turbine exhaust is discussed based on the results. The method proposed could be applied to the fault diagnosis and precise repair of the combustors of gas turbines.

Performance degradation of gas turbine is a common phenomenon during operation. Maintenance of the degraded gas turbines and improving their performance at a low cost are important in engineering. In this paper, the maintenance method based on reassembling degraded components of existing gas turbines was studied. This research was based on a type of 2MW gas turbine engine. Blue ray scanning was carried out to rebuild the 3D flow-path geometries of the compressor and turbine of a degraded engine. Then CFD simulations were carried out to compare the characteristic maps of new and degraded components. Secondly, performance tests of six engines were carried out. A correction method was developed to get the specific component characteristics using test data, which can also analyze and quantify the degradations. Also, a gas turbine performance prediction program was used to find the promising component-exchange plan within 5 given gas turbines to improve total thermal efficiency. Finally, additional test was carried out to verify the performance of the reassembled gas turbine. Through the developed method including 3D scanning, CFD simulation, and correction of component characteristics with engine test data, the component performance degradation of a specific gas turbine can be obtained in quantity. The gas turbine performance predictions based on the acquired characteristic maps showed good agreement with test data. With the help of the method developed in this work, a new gas turbine engine was obtained through exchanging the components of degraded engines, which is at a very low cost and in a short time. The improvement in total thermal efficiency was about 0.3 percentage, which was verified by engine tests.

This study explores the suitable criterion for the temperature and pressure modeling to measure the overall cooling effectiveness and the reason leading to modeling deviation at relatively low temperature and pressure test condition (Part II), taking the internally-cooled Mark II vane as an example. The method used in this study includes Artificial Neural Network (ANN), Conjugate Heat Transfer (CHT) Computational Fluid Dynamics (CFD) and experiments (Part II). The average overall cooling effectiveness of the vane was selected as the target modeling parameter. After comparing the modeling error, the results show that matched-Biot number criterion has the best performance at both constant coolant temperature and temperature ratio conditions. The maximum modeling error is 6.4% and 1.2% for those two conditions, respectively. Furthermore, the performance of these similarity criteria based on empirical correlations without the ANN model were also investigated, which is more feasible in engineering use. The accuracy of matched-Biot number criterion has an obvious decrease in this condition, but it is still the best selection at constant coolant temperature conditions. While the mass flow ratio criterion becomes the most accurate one at constant temperature ratio conditions.

In the second part of this two-part paper, several CHT (Conjugate Heat Transfer) CFD (Computational Fluid Dynamics) simulations were conducted at low temperature and pressure conditions which a 100KW or a 400KW blower can provided, in order to investigate the reason caused the modeling deviation when the matched-Biot number criterion is used for modeling. The CFD method used in this paper is the same as the Part I. Afterwards, several experiments were carried out at a mainstream condition which the 100 KW blower can provide to validate whether the deviations predicted by CFD would appear at that target modeling conditions. The mainstream inlet temperature and pressure are 307 K and 113 kPa, respectively. The model used in this paper is the Mark II turbine as well. The CFD and experiment results show that the main source of the modeling deviation comes from the suction side, the minimum deviation for the four modeling tests has over 0.05 (minimum wall temperature deviation has over 0.75 K). Apart from that, the reason leading to the suction side deviation is the variation of temperature boundary layer, which is caused by the unmatched aerodynamic parameters (such as mainstream Mach number and Reynolds number).

The inhomogeneity of temperature in a turbine is related to the nonuniform heat release and air injections in combustors. In addition, it is influenced by the interactions between turbine cascades and coolant injections. Temperature inhomogeneity results in nonuniform flow temperature at turbine outlets, which is commonly measured by multiple thermal couples arranged in the azimuthal direction to monitor the operation of a gas turbine engine. Therefore, the investigation of temperature inhomogeneity transportation in a multistage gas turbine should help in detecting and quantifying the over-temperature or flameout of combustors using turbine exhaust temperature. Here the transportation of temperature inhomogeneity inside the four-stage turbine of a 300-MW gas turbine engine was numerically investigated using 3D CFD. The computational domain included all four stages of the turbine, consisting of more than 500 blades and vanes. Realistic components (N 2 , O 2 , CO 2 , and H 2 O) with variable heat capacities were considered for hot gas and cooling air. Coolants were added to the computational domain through more than 19,000 mass and momentum source terms. his was simple compared to realistic cooling structures. A URANS CFD run with over-temperature/flameout at 6 selected combustors out of 24 was carried out. The temperature distributions at rotor–stator interfaces and the turbine outlet were quantified and characterized by Fourier transformations in the time domain and space domain. It is found that the transport process from the hot-streaks/cold-streaks at the inlet to the outlet is relatively stable. The cold and hot fluid is redistributed in time and space due to the stator and rotor blades, in the region with a large parameter gradient at the inlet, strong unsteady temperature field and composition field appear. The distribution of the exhaust gas composition has a stronger correlation with the inlet temperature distribution and is less susceptible to interference.

The supercritical CO 2 gas turbine is considered to achieve a high cycle efficiency by reducing compressor work near the critical point. But test loops made by international community show that the performance of the compressor is still quite far from the target at the design point. The paper focuses on the influence of relative velocity ratio on the performances of centrifugal impellers operating with real gas CO 2 and ideal gas CO 2 . At the same time, comparisons between characteristic curves of impellers operating at near critical, supercritical, and subcritical inlet condition with real gas CO 2 are also demonstrated. Relative velocity ratio demonstrates the same trend with real and ideal gas CO 2 in the same impeller, but the specific value is different. Impellers with real gas CO 2 could achieve a high isentropic efficiency when relative velocity ratio is in the range of 1.05∼1.5. The results show that relative velocity ratio plays an important role in compressor performance.

Nonuniformity of combustor-exit-flow temperature, also named hot-streak, could strongly increase the complexity in unsteady flow and heat transfer of gas turbine cascades. Detailed investigation into the hot-streak transportation in multi-stage gas turbine should provide a valuable reference to both fault diagnosis of combustors based on the distribution of turbine exhaust temperature. In this study, transportation of hot-streaks inside a four-stage gas turbine was numerically investigated using three-dimensional (3D), Unsteady Reynolds Averaged Navier-Stokes equations (URANS) Computational Fluid Dynamics (CFD) study with simulation domain including eight full-annulus cascades. The two-dimensional (2D) temperature fields at the rotor-stator interfaces and turbine exit was acquired by the full-annulus URANS CFD, and then discussed through comparing the circumferential and radial distributions and also the ranges of time-dependent fluctuations. The time-average temperature distributions were analyzed by Discrete Fourier Transform (DFT) method in circumferential direction to characterize the distribution. The URANS CFD results show that the hot-streaks could travel through the entire turbine and result in nonuniformity of turbine exhaust temperature. The rotor-stator interaction could change the 2D shape of hot-streaks and intensity the time-dependent fluctuation of flow temperature. DFT results show that the typical frequencies of circumferential temperature distribution are coinciding with the number of hot-streaks and the blade number of upstream/ downstream cascades. The DFT results also indicated that the hot-streaks migrate 35°∼45° in the circumferential direction, and the maximal relative amplitude of hot-streaks is reduced from 28% to 5%. The redistribution of hot/cold fluid in stator cascade could affect the mixing of hot-streaks depending on clocking position.

Film cooling of 1 st stage rotor blades strongly relates to the aerodynamic performance and life expectation of gas turbine. In this piece of work, a parametric study was carried out based on 3D RANS CFD methods to systematically investigate both cooling performance and aerodynamic loss of various film holes’ arrangements on the pressure side (PS) of a realistic rotor blade. In each CFD case, cylindrical film holes were arranged in one row with P/D around 4.0. More than one hundred CFD cases were carried out with various dimensionless streamwise locations (0.1, 0.2, ..., and 0.7), compound angles (−60°, −30°, ..., and 90°), and relative coolant massflows (0.2%, 0.3%, 0.45%, 0.7%, and 1.0% of mainstream massflow). Detailed distributions of adiabatic film cooling effectiveness on blade surface and aerodynamic loss with film cooling were computed for each case. It was found that coolant massflow was the decisive factor of aerodynamic efficiency φ. For near-LE film holes, −60° compound angle gave higher and more uniform η distributions at various coolant massflows. And the separation near LE could be suppressed, improving local cooling at near-hub on PS. For downstream film holes, 60° and −60° compound angles provided equivalent cooling performances, which were also better than other cases. The cooling characteristics of film holes with −30° compound angle were more susceptible to both coolant massflow and the location of ejection, which was mainly because of the blow-off and reattachment of coolant jets. For the cases with 90° compound angle, the distributions of η were generally similar, and the values increased monotonically as coolant massflows get larger.

This paper discusses the approach of cooling design optimization of a HPT endwall with 3D Conjugate Heat Transfer (CHT) CFD applied. This study involved the optimization of the spacing of impingement jet array and the exit width of shaped holes, which were different for each cooling cavity. The optimization objectives were to reduce the wall temperature level and also to increase the aerodynamic performance of the gas turbine. The optimization methodology consisted of an in-house parametric design & CFD mesh generation tool, a CHT CFD solver, a database of wall temperature distributions, a metamodel, and a genetic algorithm (GA) for evolutionary multi-objective optimization. The CFD tool was validated against experimental data of an endwall at CHT conditions. The metamodel, which could efficiently predict the aerodynamic loss and the wall temperature distribution of a new individual based on the database, was developed and coupled with Non-dominated Sorting Genetic Algorithm II (NSGA-II) to accelerate the optimization process. Through optimization search, the Pareto front of the problem was found costing only tens of CFD runs. By comparing with additional CFD results, it was demonstrated that the design variables in the Pareto front successfully reached the optimal values. The optimal spacing of each impingement array was decided accommodating the local thermal load while avoiding jet lift-off of film coolant. It was also suggested that using cylindrical film holes near throat could benefit both aerodynamics and cooling.

This paper discusses the methodology of impingement cooling optimization of a gas turbine 2 nd stage vane with 3D conjugate heat transfer (CHT) CFD analysis applied. The vane is installed with a novel impingement cooling structure in the leading cavity and a pin-fin array in the trailing edge. This study involves the optimization of the impingement cooling structure, including the location of the jet holes and the diameter of each hole. The generation of 3D model and CHT mesh was realized using an in-house code developed specifically for turbine cooling optimization. A constant pressure drop was assumed within the cooling system during optimization. To make the optimization computationally faster, a metamodel which can predict the detailed distribution of metal temperature on the vane surface was used in the second-level search together with a genetic algorithm. An optimal nonuniform impingement cooling structure in the leading cavity was automatically designed by the optimization process costing only dozens of CFD runs, which provided a more uniform temperature distribution on the vane surface and required no more coolant amount compared with the initial impingement cooling structure.

The present work experimentally quantifies the effects of thermal boundary conditions, i.e., conjugate and convective boundary conditions, on heat transfer performance for the pin fin channel in trailing edge of gas turbine blade. The geometry of pin fin arrays is typical of x/D=y/D=2.5 and H/D=1. For conjugate case, model is constructed with a relatively high conductivity material so that the Biot number of the model matches engine condition. Uniform heat flux is imposed along the external wall of pin fin arrays and highly resolved temperature distributions of internal wall is obtained with steady liquid crystal, meanwhile external temperature is measured through thermocouples. For convective case, model is constructed with low thermal conductivity material to ensure the usage of transient liquid crystal to obtain heat transfer coefficients of the internal wall on the same configuration. Both the measurements are used as boundary conditions to conduct simulation of solid part of pin fin array. Internal and external wall non-dimensional temperature distributions, as well as isothermal lines distribution, of the two cases are compared, results indicate that it will produce large errors in temperature predictions without considering conjugate effect. Further analysis are made about the mechanism of thermal boundary conditions in determining wall temperature, which demonstrates the necessity of taking conjugate heat transfer effect in turbine cooling design.

Impingement cooling is widely used in turbine vanes and combustors. With the increase of turbine inlet temperature, high heat transfer coefficient and low pressure drop are required for cooling structures. A series of new impingement configurations combined with corrugated orifice and pin-fins were developed in the present work. Both transient liquid crystal (TLC) and pressure measurement were applied on the impingement cooling structures. A 3D numerical method was also used for conjugate heat transfer simulation. Corrugated orifice helps decrease the pressure drop by decreasing the speed of cross flow. Experimental data show that, corrugated orifice is helpful in reducing pin-fin induced pressure drop but contributes little to heat transfer. Pin-fins increases both the heat transfer and pressure drop lightly. Conjugate heat transfer simulation shows that pin-fins significantly reduce the metal temperature by conduction. Structures with pin-fins can make a good use of the large surface area of corrugated orifices.

As the advanced heavy-duty gas turbine develops, the turbine inlet temperature and pressure have increased quite significantly to achieve better performance. The flow and heat transfer conditions of hot components including combustor and turbine become even more extreme than ever which need corresponding aerodynamic and cooling design development. The issue of combustor-turbine interaction has been proposed as a complicated research topic. Currently the hot streak, turbulence intensity, swirling flow, radiation are the four important factors for combustor-turbine interaction research according to the literature. Especially as the turbine inlet temperature increases, the radiative heat transfer plays a more and more important role. In this paper, a first stage vane is selected for the conjugate heat transfer simulation including radiative heat transfer since it is almost impossible to identify the radiative effect in experiment. The goal is to examine the effects of radiative heat flux and temperature increment caused by radiation. Several radiative factors including the inlet radiation, gas composition, vane surface emissivity and outlet reflection are investigated. The temperature distribution and heat flux enhancement under different conditions are compared, which can provide reference to the turbine heat transfer design. The general information of radiative effect can be summarized by quantitative analysis. Results show that the temperature increases obviously when considering the radiation effect as expected. However, these factors show distinct influence on the vane temperature distribution. The inlet radiation has significant impact on the vane leading edge and pressure side. Besides the gas radiation plays quite uniform on the whole vane surface.

A tripod cylindrical film hole with asymmetric side holes is studied numerically and experimentally on a flat plate for higher film cooling effectiveness. Firstly, the influences of geometrical parameters are studied and the optimum configurations of the asymmetric tripod hole are found in a DoE optimization study based on an improved numerical model for film cooling prediction, in which more than one hundred 3D CFD simulations are carried out. Then one optimum configuration of the asymmetric tripod hole is examined experimentally using pressure-sensitive paint (PSP) measurements, and compared against the experimental results of the simple cylindrical film hole and a well-designed shaped film hole. The flow and heat transfer characteristics of the asymmetric tripod holes were explored from the DoE results. The side holes can form a shear vortex system or an anti-kidney vortex system when proper spanwise distances of them are adopted, which laterally transports the coolant and form a favorable coolant coverage. According to the experimental results, the cooling performance of the optimized asymmetric tripod hole is significantly better than that of the simple cylindrical hole, especially at high blowing ratios. And the optimized asymmetric tripod hole can provide almost the same or even higher film cooling effectiveness on the flat plate compared with the shaped hole in the same flow conditions.

Shaped film holes can achieve higher film cooling effectiveness compared with the simple cylindrical film holes. According to former studies, the geometry of the shaped film holes has significant influence on the cooling performance. In order to maximize the film cooling effectiveness of the shaped holes, a two-level design optimization methodology of the hole exit shaping is developed in the present study. The optimization methodology consists of a parametric design and CFD mesh generation tool called Coolmesh, a RANS CFD solver, a database of film cooling effectiveness distributions, a metamodel, and a genetic algorithm (GA) for evolutionary optimization. A binary parametric representation of the 2D hole exit shaping is initiated based on the B-spline methods. The metamodel can efficiently predict the detailed distribution of film cooling effectiveness using the CFD results in the database, which is continuously updated for higher accuracy. In each first-level iteration, a second-level GA optimization search is carried out coupled with the metamodel, and then the optimal geometry is evaluated using CFD methods and added to the database. An anisotropic turbulence model is applied to the CFD solver for higher accuracy according to a detailed experimental validation using PSP measurements. In the present study, three design optimizations of the shaped holes without and with compound angles are carried out on a flat plate. The optimization methodology can efficiently find the optimal geometries of shaped holes using only hundreds of CFD runs. For the shaped holes with compound angle, the optimized geometry can generate a back flow vortex which interacts with the shear vortex and weakens the mixing of coolant and hot gas, resulting in a higher film cooling effectiveness on the plate.

Cooling design for the air-cooled turbine blades is a critical issue in modern gas turbine engineering. Advances in CFD technology is providing new prospects for turbine cooling design, as the optimum cooling structures of the blades could be designed through the optimization search coupled with the Conjugate Heat Transfer (CHT) analysis. In this paper, the optimization study for the rib arrangement of a rib-roughed channel in a rotor blade is discussed. The optimization study introduced is realized utilizing a parametric analysis platform, which consists of the parametric design and mesh generation tool and the commercial CHT solver ANSYS CFX. For the optimization study, firstly a group of Design of Experiments (DoE) analysis of a rib-roughed rectangular channel is performed in order to find the optimum rib arrangement and to explore the objective of the optimization search. Then, the optimization search of the optimum rib arrangement is performed for a rib-roughed channel within a rotor blade based on the multi-island Genetic Algorithms (GA) of iSIGHT. During optimization search, a constant pressure drop is assumed within the cooling system, and the CHT simulations are approached for the interior only in order to make the search computationally faster. According to the DoE analysis, minimizing the averaged wall temperature on blade surface is chosen as the optimization objective for the design of rib arrangement. The results of the GA search shows that the optimal rib arrangement with best cooling performance can be decided, and the optimal mass flow rate for the cooling channel is found simultaneously. The optimum schemes of the rib arrangement found by the DoE analysis and GA search are quite identical, which further validates the feasibility of design optimization for the blade cooling structure with the GA and CHT simulations.