Volume 6A: Heat Transfer — Combustors; Film Cooling最新文献

{"title":"Turbine Vane Passage Experiments Documenting Evolution of Secondary Flows With Changes in Combustor Coolant Injection Flowrates","authors":"Kedar P. Nawathe, Y. Kim, T. Simon","doi":"10.1115/gt2022-82697","DOIUrl":"https://doi.org/10.1115/gt2022-82697","url":null,"abstract":"\u0000 A secondary flow system with a dominant passage vortex pattern has been observed in many gas turbine vane passage studies in which there is no upstream coolant injection or only near-passage endwall coolant injection (no combustor cooling). However, it was shown in recent studies that combustor coolant introduced upstream of the vane passage changes secondary flow patterns in the passage. This results in a different secondary flow vortex system, called the ‘impingement vortex’ system. It was discussed in recent literature having combustor coolant injection. Until now, there has been no study on how increases in combustor coolant momentum effect transition from the passage vortex system to the impingement vortex system. Such a study is presented in the present paper. Velocity component measurements are taken using a five-hole probe at three axial locations in the vane passage to document secondary flow development throughout the passage. Four combustor coolant flowrate cases are considered along with a comparison case having no coolant injection. It is shown that as the combustor coolant flowrate increases, the passage vortex system weakens and, at a sufficiently high combustor coolant flowrate, the impingement vortex system appears. Knowing the detailed flow physics of this transition between the two secondary flow systems is helpful for turbine thermal designers who wish to understand how secondary flows transport coolant within the turbine vane passage.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"54 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121564593","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Assessment of a Conjugate Heat Transfer Method on an Effusion Cooled Combustor Operated With a Swirl Stabilized Partially Premixed Flame","authors":"A. Amerini, S. Paccati, L. Mazzei, A. Andreini","doi":"10.1115/gt2022-81345","DOIUrl":"https://doi.org/10.1115/gt2022-81345","url":null,"abstract":"\u0000 Computational fluid dynamics play a crucial role in the design of cooling systems in gas turbine combustors due to the difficulties and costs related to experimental measurements performed in pressurized reactive environments. Despite the massive advances in computational resources in the last years, reactive unsteady and multi-scale simulations of combustor real operating conditions are still computationally expensive.\u0000 Modern combustors often employ cooling schemes based on effusion technique, which provides uniform protection of the liner from hot gases, combining the heat removal by means of heat sink effect with liner coverage and protection by film cooling. However, a large number of effusion holes results in a relevant increase of computational resources required to perform a CFD simulation capable of correctly predicting the thermal load on the metal walls within the combustor. Moreover, a multi-physics and multi-scale approach is mandatory to properly consider the different characteristic scales of the several heat transfer modes within combustion chambers to achieve a reliable prediction of aero-thermal fields within the combustor and wall heat fluxes and temperatures. From this point of view, loosely-coupled approaches permit a strong reduction of the calculation time, since each physics is solved through a dedicated solver optimized according to the considered heat transfer mechanism. The object of this work is to highlight the capabilities of a loosely-coupled unsteady multi-physics tool (U-THERM3D) developed at the University of Florence within ANSYS Fluent.\u0000 The coupling strategy will be employed for the numerical analysis of the TECFLAM effusion cooled swirl burner, an academic test rig well representative of the working conditions of a partially premixed combustion chamber equipped with an effusion cooling system, developed by the collaboration of the Universities of Darmstadt, Heidelberg, Karlsruhe, and the DLR. The highly detailed numerical results obtained from the unsteady multi-physics and multi-scale simulation will be compared with experimental data to validate the numerical procedure.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"115 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124005741","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Numerical Study of Combustor-Turbine Interaction by Using Hybrid RANS-LES Approach","authors":"S. G. Tomasello, A. Andreini, R. Meloni, S. Cubeda, L. Andrei, V. Michelassi","doi":"10.1115/gt2022-82139","DOIUrl":"https://doi.org/10.1115/gt2022-82139","url":null,"abstract":"\u0000 The complex flow field of gas turbine lean combustors is meant to reduce NOx emissions and maintain a stable flame by controlling the local temperature and promoting high turbulent mixing. Still, this may produce large flow and temperature unsteady distortions capable of disrupting the aerodynamics and heat transfer of the first high-pressure-turbine cooled nozzle. Therefore, the interaction between the combustion chamber and the turbine nozzle is analyzed first with the help of scale-resolving simulations that notably also include a realistic turbine nozzle cooling system. To determine the nature and severity of the interaction, and the risks associated to performing decoupled simulation, the results of the coupled computer simulation are analyzed and compared with those of decoupled simulations. In this case, the combustor is computed by replacing the turbine nozzle with a discharge convergent with the same throat area, and the conditions at the interface plane are used as inlet boundary conditions for a conventional RANS of the nozzle. The analyses of the coupled and decoupled simulation reveal that the combustion chamber is weakly affected by the presence of the nozzle, whereas the two thermal fields of the nozzle surface differ considerably, as well as the disruption of the film cooling by the incoming flow distortions.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"599 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123331528","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Experimental and Numerical Study on Effusion Cooling Configuration for the Swirl CMC Combustor Platform","authors":"Kun Du, Qihao Chen, Tingrui Liang, YiHao Jia, Cun-liang Liu","doi":"10.1115/gt2022-81567","DOIUrl":"https://doi.org/10.1115/gt2022-81567","url":null,"abstract":"\u0000 With advantageous thermal, strength, and weight properties at high temperatures, Ceramic Matrix Composite (CMC) is regarded as the potential structural material for modern aeroengines with the extreme increase of operating temperature. However, the anisotropic thermal conductivities caused by the weaving type call for a novel cooling structure design since it brings a completely different thermal conduction performance inside the hot components compared to the superalloy. The infrared thermographic experiment was carried out based on the SiC/SiC composite platform prepared by a 2-D plain weave braid structure with the CVI process with three staggered effusion hole configurations in this paper to explore the cooling characteristic of the CMC. The surface overall cooling effectiveness (ϕ) of the platform was measured under a temperature ratio of 1.5 (Tg/Tc = 1.5) and seven mass flow ratios. And a separate numerical simulation with different swirl effects was conducted as a supplement for the combustor operation condition. The results indicate that the thermal conductivity along the thickness direction is of great importance for the platform cooling. The overall cooling effectiveness of the CMC platform was smaller in comparison with the superalloy platform because of its smaller through-thickness thermal conductivity. Moreover, the concentrated coolant outflow is beneficial for the CMC cooling performance given that the area-averaged overall cooling effectiveness disparity of the two materials was narrowed with the smaller hole spacing. The swirling flow strongly inhibits the outflow of coolant near the swirling core but expands the lateral coverage of the film simultaneously. Such behavior alters the film cooling potential as well as the cooling pattern. The difference between CMC and superalloy was increased with the low swirl situation and decreased with the high swirl situation.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"20 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126936572","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Film Cooling Hole Shape Effects on Turbine Blade Heat Transfer – Part I: Computational Comparison to Experiment","authors":"Spencer J. Sperling, Louis E. Christensen, Randall M. Mathison, H. Aksoy, Jong-Shang Liu, Jeremy B. Nickol","doi":"10.1115/gt2022-78244","DOIUrl":"https://doi.org/10.1115/gt2022-78244","url":null,"abstract":"\u0000 Gas turbines generate highly unsteady flow fields, which are further complicated by cooling. Film cooling technology is dependent on experimental and computational research performed for simplified geometries, which can be difficult to translate to turbine domains. To help explain disconnects between flat plate experiments and turbine operation, this study performs computational research, grounded with experimental measurements, examining the film cooling performance of several different hole shapes on rotating turbine blades.\u0000 Round, fan, and advanced anti-vortex hole geometries are incorporated into experimental hardware and computational models, and pressure and heat flux results form a basis for comparisons between computation and experiment as well as between the cooling geometries. Unsteady and Steady simulations are both evaluated, and results indicate significant model accuracy improvement from the inclusion of unsteady consideration.\u0000 Shaped film cooling holes are observed to provide stronger film effectiveness traces on the blade. Both the fan and advanced shaped film cooling holes generate stronger cooling jet cores that remain close to the blade wall on the pressure surface, suction surface, and near the leading edge. Advanced shaped holes provide increased lateral spread and increased resistance to radial migration.\u0000 The results of this study help identify benefits of using shaped film cooling on the turbine blade as well as the mechanisms generating the cooling benefits. This will help designers weigh the manufacturing costs of shaped film cooling holes as well as identify areas of the blade where shaped film cooling is needed and where it is not. Additionally, this study observes significant improvements in blade heat transfer predictions through unsteady treatment alone, indicating better computational agreement can be achieved by leveraging lower-cost RANS simulation tools.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"26 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114522309","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"The Cooling Effect of Combustor Exit Louver Scheme on a Transonic Nozzle Guide Vane Endwall","authors":"Shuo Mao, Daniel Van Hout, Kai Zhang, W. Ng, Hongzhou Xu, M. Fox","doi":"10.1115/gt2022-82800","DOIUrl":"https://doi.org/10.1115/gt2022-82800","url":null,"abstract":"\u0000 The ever-increasing combustor exit temperature in modern turbine engine designs raises challenges for the nozzle guide vane cooling. Due to the challenges of NGV cooling design, the cooling effect from the combustor cooling features can prove valuable. This study investigates, experimentally and numerically, the cooling effect of a louver cooling scheme near the combustor exit on the NGV endwall. The wind tunnel testing and CFD simulation are carried out with engine-representative conditions of Maexit = 0.85, Reexit = 1.5 × 106, Tu = 16%, and DR = 2.1. Various coolant mass flow rates from 1% to 4% are tested to demonstrate the effect of the coolant rate.\u0000 For the geometry studied, the results found a critical MFR between 1%∼2%. By exceeding this value, the coolant forms a uniform film which provides good coverage upstream of the NGV passage inlet. As for the cooling of the NGV passage, the MFR of the range investigated is not sufficient for desirable cooling performance. The pressure side endwall proves most difficult for the coolant to reach. In addition, the fishmouth cavity at the combustor-NGV passage causes a three-dimensional cavity vortex that transports the coolant in the pitch-wise direction. The coolant transport pattern is dependent on the coolant MFR. Based on the results, it is proposed to combine this louver scheme with the upstream jump cooling scheme for a desirable NGV cooling system.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"17 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114787208","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Effect of Inlet Preswirl on the Film Cooling Characteristics of Turbine Vane Surface","authors":"Xinnan Chen, Zhigang Li, Jun Li, Qingzong Xu, Qiang Du","doi":"10.1115/gt2022-82755","DOIUrl":"https://doi.org/10.1115/gt2022-82755","url":null,"abstract":"\u0000 Conjugate heat transfer (CHT) analyses were conducted to validate the accuracy of numerical method in the prediction of aerodynamic and thermal performance of the film cooled C3X vane. The effect of inlet preswirl on the film cooling performance of the vane surface was investigated, while the Nu and the net heat flux reduction (NHFR) were also calculated to analyze the heat transfer characteristics. In particular, the effects of clocking positions and swirl orientations of the inlet preswirl with varied mass flow ratios of film coolant were studied individually. The results indicate that the inlet preswirl aligned to the vane leading edge (LE) would significantly enhance the heat transfer on the vane surface while the inlet preswirl aligned to the passage would weaken the heat transfer but lead to a nonuniform distribution of Nu. The inlet preswirl aligned to the LE would also seriously affect the flow structure of the showerhead film coolant and do deteriorate to the film cooling performance on the vane surface especially for the suction surface, in contrast, the inlet preswirl aligned to the passage would even improve the film cooling performance under some mass flow rates of film coolant. Moreover, the disadvantage of inlet preswirl to the downstream film cooling performance is almost negligible, while the inlet preswirl is even beneficial to the film coolant coverage under some mass flow rates of coolant in this paper, especially for the suction surface (SS) with a large mass flow rate and the pressure surface (PS) with a small mass flow rate.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"68 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"117318844","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Effects of Part-to-Part Flow Variations on Overall Effectiveness and Life of Rotating Turbine Blades","authors":"Brian F. Knisely, Reid A. Berdanier, J. Wagner, K. Thole, A. Arisi, C. Haldeman","doi":"10.1115/gt2022-83216","DOIUrl":"https://doi.org/10.1115/gt2022-83216","url":null,"abstract":"\u0000 As firing temperatures in gas turbine engines continue to increase to achieve high efficiencies, components in the main gas path must be protected with cooling flows to ensure lifing targets are met. Manufacturing variations, however, influence the performance and life characteristics of components with the same nominal design. This study presents blade flow and overall cooling effectiveness measurements for nine true-scale, aero engine turbine blades with realistic manufacturing variations. Flow measurements were made through each blade at a fixed pressure ratio to determine flow variability between holes and between blades. Infrared thermography was used to capture spatially-resolved temperature measurements reported as overall effectiveness on the same nine blades under high-speed rotating conditions at the Steady Thermal Aero Research Turbine Laboratory. Thermal performance was correlated with blade flow performance indicating substantial blade-to-blade variations resulting from manufacturing differences. Measurements also indicated wide variations in cooling jet trajectories as well as overall cooling effectiveness. Finally, the observed blade-to-blade variations in effectiveness were scaled to engine conditions with lifing estimates showing some blades would be expected to last only half as long as others due to manufacturing variability.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"18 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"115454552","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Influence of Porosity on Double-Walled Effusion-Cooled Systems for Gas Turbine Blades","authors":"M. Courtis, P. Ireland","doi":"10.1115/gt2022-80377","DOIUrl":"https://doi.org/10.1115/gt2022-80377","url":null,"abstract":"\u0000 Double wall effusion cooling (DWEC) systems for gas turbine blades utilise two skins connected by pedestals and take advantage of cooling benefits provided by impingement jets and film holes. The latter exhausts coolant externally onto the blade surface forming a protective layer against the high external heat loads, which can be enhanced via the beneficial influence of adjacent films. Consequently, increasingly porous outerskins are being considered in order to provide greater thermal protection and/or reduce the required coolant mass consumption. To realise such systems, further research must understand how the internal aerothermal field is affected by high porosity.\u0000 A semi-decoupled unit-cell computational fluid dynamics (CFD) method is applied to a range of DWEC systems to understand overall cooling effectiveness as well as internal characteristics. A comparison of internal convection highlights a shift in the breakdown of cooling performance, due to the large changes in wetted surface area of the outerskin. For low porosity, most of the internal cooling occurs through the jet impingement on the internal outerskin wall, while the addition of more film holes provides an increasingly greater proportion of convective heat transfer. On the external surface, porosity increased film effectiveness due to film superposition, provided a more uniform film coverage, and reduced the likelihood of jet-lift-off.\u0000 Coupling the benefits of internal cooling and film effectiveness resulted in a reduction of mean metal temperature, peak temperature and temperature gradient between the outer and inner walls. Criteria reflecting the main drivers for thermal fatigue. Despite these benefits, for the most porous DWEC configuration a variation in mass flow between film holes was observed, and in some cases the risk of hot gas ingestion was evident. DWEC components would benefit from further understanding of the susceptibility to blockage, the pressure margin limits and the extent of flow migration.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"5 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129057789","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

{"title":"Efficient Modelling of Blade Film Cooling in Gas Turbines","authors":"J. Penrose, L. Zori, J. Morales, S. Patil, D. Pons, S. Rida","doi":"10.1115/gt2022-80700","DOIUrl":"https://doi.org/10.1115/gt2022-80700","url":null,"abstract":"\u0000 One of the most effective ways to mitigate thermal fatigue in a high-pressure turbine’s blades is by cooling the blade from inside and outside via an intricate cooling system. The cooling flow passes from the blade interior through many small holes to form a cooling film on the blade surface. In the early blade design phases, the designers must accurately determine the location, the pattern, the distribution density, the shape, and the size of cooling holes to maximize the blade film cooling whilst maintaining the external aerodynamics. Hundreds of simulation cycles may be needed to reach an optimal design.\u0000 In this work, a film-cooling model and associated workflow are proposed. The model implementation within Ansys-Fluent uses a virtual-boundary concept which does not require the explicit resolution of the holes. The benefit of this is its compatibility with existing turbomachinery solution methods, and the consistency with subsequent mesh refinement toward resolved hole geometry. The workflow will allow designers to dissociate the uncooled aerodynamic geometry and mesh from the hole/film-cooling design, during the early design iterations.\u0000 For verification of this approach, results for a simplified single hole-setup on a flat plate are first presented. A typical gas turbine vane configuration is then used to demonstrate industrial application. The results from an aerodynamic mesh refinement study show good agreement with the resolved model. Flexibility within the workflow for changing the location, shape and properties of the holes is also demonstrated. This approach, therefore, represents a useful design tool where multiple hole pattern configurations could be quickly assessed.","PeriodicalId":267158,"journal":{"name":"Volume 6A: Heat Transfer — Combustors; Film Cooling","volume":"32 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2022-06-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"117010365","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}