Improving the Film Cooling of a Rotor Blade Platform

Transcription

1 Improving the Film Cooling of a Rotor Blade Platform Giovanna Barigozzi, Antonio Perdihizzi, Roberto Abram To ite this version: Giovanna Barigozzi, Antonio Perdihizzi, Roberto Abram. Improving the Film Cooling of a Rotor Blade Platform. 16th International Symposium on Transport Phenomena and Dynamis of Rotating Mahinery (ISROMAC 2016), Apr 2016, Honolulu, United States. <hal > HAL Id: hal Submitted on 3 May 2017 HAL is a multi-disiplinary open aess arhive for the deposit and dissemination of sientifi researh douments, whether they are published or not. The douments may ome from teahing and researh institutions in Frane or abroad, or from publi or private researh enters. L arhive ouverte pluridisiplinaire HAL, est destinée au dépôt et à la diffusion de douments sientifiques de niveau reherhe, publiés ou non, émanant des établissements d enseignement et de reherhe français ou étrangers, des laboratoires publis ou privés. Distributed under a Creative Commons Attribution 4.0 International Liense

2 Improving the Film Cooling of a Rotor Blade Platform Giovanna Barigozzi 1, Antonio Perdihizzi 1*, Roberto Abram 2 Abstrat This paper shows the results of an experimental ativity developed in ooperation between Ansaldo Energia and the Department of Engineering and Applied Siene of Bergamo University with the aim of assessing the impat of newly designed holes on the thermal protetion of a rotor blade platform. The original rotor blade platform featured 10 ylindrial holes loated along the blade pressure side. Moreover, the hannel front side was ooled exploiting the seal purge flow exiting the stator to rotor interfae gap. The front mid hannel, and partiularly the region around the inter-platform gap, remained unooled. To protet this region two sets of ylindrial holes were designed and manufatured on a 7 blade asade model for experimental verifiation. Aerodynami and thermal tests were arried out at low Mah number. To evaluate the interation of injeted flow with seondary flows a 5hole probe was traversed downstream of the trailing edge plane. The thermal behavior was analyzed by using Thermohromi Liquid Crystals tehnique, so to obtain film ooling effetiveness distributions. The 7-hole onfiguration oupled with a low blowing ratio of about 1.0 provided the best thermal protetion without any impat on the aerodynami performane. Keywords gas turbine platform film ooling 1 Department of Engineering and Applied Sienes, University of Bergamo, Italy 2 Ansaldo Energia S.p.A. - Hot Gas Path Engineering, Genova, Italy *Corresponding author: antonio.perdihizzi@unibg.it Nomenlature blade hord D hole diameter DR density ratio H blade height L hole length M 1 pt, p1 inlet loss free blowing ratio p p t,1 1 Ma Mah number Re 2is U 2is isentropi outlet Reynolds number s blade pith S inter-platform line oordinate T temperature Tu turbulene intensity level U u v w loal mean veloity u,v,w streamwise, transverse and spanwise veloity omponents X,Y,Z asade oordinate system flow angle (axial diretion) T T T T film ooling effetiveness aw U 1 s non dimensional vortiity U 2is U 2 U 2is, ms Subsripts 1 inlet 2 exit ax axial aw adiabati wall loal energy loss oeffiient is s INTRODUCTION ooling flow isentropi ondition streamwise free stream In modern gas turbines the ahievement of high performane requires a ontinuous inrease of turbine inlet temperature. Hene an enhanement of thermal protetion of all surfaes diretly exposed to the hot gases, inluding platform regions, is needed. Many researhers have investigated film ooling on vane end walls, both through upstream slots [1-3] and disrete holes within the passage [4-10]. From the above studies, it is found that film ooling is strongly affeted by end wall seondary flows. The ross flow tends to transport the injeted oolant from the pressure side to the sution side of the passage. Only few works in the open literature deal with film ooling of rotor end walls. Blair [11] deeply analysed the effets of flow three dimensionality on the heat transfer in a rotating blade row. The flow struture in the rotor blade passage has been investigated by Wang et al [12]. Goldstein and Spores [13] deteted high heat transfer near the bladeend wall juntions of the passage, due to an intense ativity of orner vorties. Other researh ativities are foused on the analysis of the film ooling effetiveness due to purging of sealing flow through the stator-rotor axial gap. The areas that are typially diffiult to ool,

3 Improving the Film Cooling of a Rotor Blade Platform - 2 i.e. the region near the leading edge, an be ooled effetively with the upstream slot injetion. Gao et al. [14] and Narzary [15] tested a typial labyrinth-like stator-rotor seal under different blowing onditions, looking for the thermal protetion apability of the purge flow. Papa et al. [16] performed heat transfer measurements and flow visualizations for different blowing ratios in a linear rotor with an injetion slot upstream of rotor blades. In all these papers, tests were performed on linear asade wind tunnels with radial purge flow, i.e. without modeling the rotation effets. Barigozzi et al. [17] experimentally investigated the effet of rotation on slot platform ooling. Rotation effet was simulated in a linear asade by installing fins inside of the slot to hange the oolant to mainstream injetion angle in the tangential diretion. They showed that rotation has a signifiant impat on the 3D flow field and film ooling effetiveness, as it reinfores the seondary flows development and migration aross the blade passage. Pau et al. [18] arried out a omplete aero-thermal investigation on the purge and platform film ooling on a transoni turbine stage. Suryanarayanan et al. [19-20] investigated film ooling through the wheel spae avity in a rotating platform. All these studies showed how the interation of the emerging purge flow struture with the upoming main flow is ompliated by the influene of rotation, resulting in a highly skewed flow emerging from the slot. The leading edge region of the platform an be effetively ooled by this purge flow. But suh oolant flow only sueeds in ooling the front end wall part, before being aptured by the passage vortex, leaving most of the hannel pratially unooled. So disrete holes are usually added inside of the passage in some ritial regions. In some reent works, it has been shown that a omplete film ooling protetion on a rotating platform an be provided by ombining upstream stator-rotor gap ejetion and injetion from disrete holes loated in the rear part of the passage [14,21-22]. Disrete holes distributed along the pressure side an properly ool the rear part of the platform [23], but some more holes are still required to protet the mid hannel front region, and partiularly the inter-platform gap region. The present paper deals with the aerodynami and thermal behaviour of a rotor blade asade with a disrete hole ooling sheme on the platform. Starting from the original oolant injetion sheme with holes loated in the rear part of the passage near to the filleted pressure side of the blade (Fig. 1a - [24]) and from the results oming from an upstream slot ooling onfiguration (Fig. 1b - [17]), a new set of holes was oneived to protet the blade to blade inter-platform front region, a region not overed by the aforementioned ooling shemes. The loal effetiveness distribution is strongly dependent on the holes layout: the arrangement of film ooling holes on the platform is a key point. Two ooling shemes were thus designed, with an inreasing number of holes. This paper shows the results of different hole number and the effets of injetion onditions on film ooling effetiveness of a rotor blade platform. M 1 = 2.4 a) b) Figure. 1. Platform Film ooling effetiveness distribution: a) pressure side ooling holes [24] and b) purge gap [17]. 1. EXPERIMENTAL DETAILS Tests have been performed at the Turbomahinery Laboratory of the University of Bergamo in the wind tunnel for linear rotor asades. It is a ontinuous running sution type wind tunnel that assures a omplete optial aess beause entirely made up of Plexiglas (Fig. 2). The asade model onsists of a 7 blade asade whose geometry (Fig. 2) is typial of a first high pressure turbine rotor blade. It has been tested at low Mah number (Ma 2is =0.3) with a low inlet turbulene intensity level (Tu 1 = 0.7 %). Geometrial details of the asade and operating onditions are all summarized in Table 1. On the hub side, the onnetion between the blade and the platform is realized through a 3D fillet, whose trae in the platform plane is depited

4 Improving the Film Cooling of a Rotor Blade Platform - 3 in Fig. 3. Only one hannel is ooled. The original ooling sheme was made of ten 0.7 mm diameter holes loated on the filleted end wall along the blade pressure side (Fig. 3). More details on this geometry are given in Barigozzi et al. [17]. On the same platform five more holes with 1.0 mm diameter have been initially realized. These holes are distributed along the inter-platform gap with an inlination angle of 40 with respet to the end wall surfae. Note that this is a quite high angle ompared to those usually adopted for film ooling, but it was imposed by design onstrains related to the internal gas path of the real blade. Hole length over diameter ratio L/D is Two plenum integrated in the blade platform (Fig. 4) independently feed the two sets of holes (the 10 original holes and the new ones). Starting from the five hole onfiguration, a seond ooling sheme was realized by adding two more holes in line with the first five (Fig. 3). Injetion onditions were ontrolled through a ontinuous monitoring of oolant total pressure in the two feeding hambers. The amount of the injeted mass flow was in fat so small to be pratially notmeasurable with a suffiiently high auray. The inlet loss free blowing ratio M 1 was thus used to identify eah injetion ondition for the two sets of holes. Coolant total pressure and temperature are measured by pressure taps and T-type thermoouples loated on the bak side wall of eah plenum. The unertainty in the M 1 value was omputed on the basis of pressure transduers (± 5.5 Pa) and Pitot probe (± 10 Pa) unertainties. M 1 resulted to be ± 0.04 at a value of M 1 = 1.0 and ± 0.08 at a value of M 1 = 2.0. Aerodynami measurements were performed 8% of the axial hord (X/ ax = 108%) downstream of the trailing edge plane by using a 5-hole miniaturized aerodynami pressure probe. The measurement plane overs two blade passages and extends over half of the blade span. The measurement grid is made of 30 points per pith in tangential diretion times 15 points along the blade height. The grid spaing was redued approahing the end wall surfae. Unertainties in both stati and total pressure have been estimated to be ± 0.15% of dynami pressure. Sprayable wide banded Thermohromi Liquid Crystals (Hallrest BM/R25C10WC17-10) were used to get the film ooling effetiveness distributions. The surfae is illuminated by means of two strips of white light LED while TLC images were aquired by using a Nikon D7100 amera. The TLC alibration was performed in situ, substituting the blade entral passage with a alibration devie onsisting in an instrumented (10 T-type thermoouples, T = ± 0.1 C) flat aluminium plate over whih a temperature gradient an be generated. All alibrations and measurements were performed in the dark, in order to eliminate any influene of bakground illumination. Moreover, an illumination intensity as uniform as possible was provided to the model surfae by properly orienting the lighting system, in the meanwhile avoiding any light refletion onto the amera. Figure 2. The wind tunnel (1: inlet dut; 2: test setion; 3: tailboard; 4: diffuser; 5: fan; 6: AC motor; 7: disharge hannel). Inter-platform gap PS holes New holes Figure 3. Casade and endwall ooling geometry. Table 1. Casade geometry and operating onditions. s/ =0.69 H/ = 1.26 = mm 2 = Tu 1 = 0.7 % Ma 2is = 0.3 Ma 1 = M 1,PS = 1 2 Re 2is = M 1,New = 1 2 During tests the heated seondary flow (DR = 0.95) was injeted into the main flow at ambient temperature. The time history of the TLC image was reorded by the amera, together with the temperature variation inside the feeding hambers T and the main flow temperature T. The RGB to hue onversion [25] was applied to a single image reorded after a time period of about 60 s, i.e. when a stable temperature level inside eah plenum was reahed, as well as on the end wall surfae. Figure 5 is an example of an image seleted. In image seletion partiular attention

5 Improving the Film Cooling of a Rotor Blade Platform - 4 was paid to avoid important ondution phenomena in the most ritial region, i.e. just upstream of PS hole loations. omputation was based on the maximum temperature measured inside of eah plenum, one a stable oolant temperature value was reahed. The relatively large thikness of the end wall (Plexiglas made) assured to omply with wall adiabati ondition over most of the platform. Film ooling effetiveness unertainty depends on TLC (T w = ± 0.3 C) and thermoouple measurements (T = ± 0.1 C and T = ± 0.5 C) and ondution effets. In regions where ondution phenomena do not exist, unertainty will range from ± 5 % with = 0.5, up to about ± 15 % when = 0.1. Larger unertainty will exist if ondution phenomena beome relevant. More details on thermal tests are in Barigozzi et al. [23] vetors are reported in Fig. 6, to provide the seondary flow onfiguration downstream of the asade. Only the plots referred to the seven holes onfiguration and the minimum tested M 1 are shown, as no signifiant variations of seondary flow onfiguration were found for the other sheme or blowing onditions. It is worth noting that very similar results were also obtained testing the unooled asade. This essentially beause the injeted flow rate is always very small and insuffiient to affet the seondary flows development. Moreover, oolant injetion is performed downstream of the passage vortex separation line (Fig. 7); hene most of the injeted flow is not aptured by passage vortex and it is only subjeted to the endwall ross flow from pressure to sution side. Typial and very well defined seondary flows strutures an be observed from Fig. 6. The flow field is dominated by the presene of a well defined passage vortex, orresponding to the positive vortiity region and to the loss ore on the sution side of the blade wake. The high flow turning makes the passage vortex position to be signifiantly shifted towards mid span. A relevant ross flow is present at the end wall, driving low energy boundary layer flow and the injeted oolant from the pressure to the sution side of the vane. No trae of orner vortex loss peak an be observed, probably beause it is onfined in a thin layer outside of the measuring domain. Figure 4. The ooling flow supply lines. Figure 5. Example of seleted image. 2. AERODYNAMIC RESULTS Aerodynami results showing the seondary flow struture have been obtained by traversing 5-hole probe 0.08 ax downstream of the blade trailing edge. The ontour plots of energy loss oeffiient and stream wise vortiity with seondary veloity Figure 6. and distributions at X/ ax =108% for 7 holes (M 1,PS = 1.3 and M 1,7 holes = 1.1).

6 Improving the Film Cooling of a Rotor Blade Platform THERMAL RESULTS Film ooling effetiveness distributions measured for the 5 hole ooling sheme at three different blowing onditions (M 1 about 1.2, 1.6 and 2.0 for both PS holes and 5 new holes) are presented in Fig. 8. Data of eah test have been normalized referring to the maximum measured value η max over all tests. Figure 7. Oil and dye flow visualization (solid end wall). First fousing on the performane of the new holes, an inrease of inlet loss free blowing ratio results in a derease of thermal protetion. In fat, at the lowest blowing ratio (i.e. M 1 = 1.2) oolant is disharged through all the holes, espeially the last one. Coolant remains attahed to the wall, but its trajetory follows the three dimensional separation lines and never rosses it (see traes of these lines S 1 and S 2 on the effetiveness ontour plot of Fig. 8a). Anyways, oolant persisteny is good enough to allow it to almost reah the sution side of the adjaent blade. Unfortunately, the area interested by film ooling is quite limited as oolant quikly mixes with the passage vortex, hene a poor thermal overage is attained. This is surely due to the fat that new holes are loated lose to the separation lines, resulting in a strong interation between jets and passage vortex. Moreover, their injetion angle is quite high (40 ), resulting in a strong mixing. Inreasing the injetion rate up to M 1 = 1.6 jet persisteny downstream of the holes strongly redues. This is due to the relatively high injetion angle of 40, that does not allow the jets to remain attahed to the wall at this injetion ondition. This is even worst at the highest tested M 1 of 2.0 when a omplete jet liftoff takes plae. As a onsequene, pratially no traes of oolant an be deteted at the wall for the largest tested injetion ondition. Fousing on the downstream PS holes, aording to previous tests, a blowing rate inrease always translates into a thermal protetion improvement, that in the limited range of M 1 values tested in this researh is quite small. Previous experiments involved a wider range of injetion onditions, up to M 1 = 4.0. As typial values of blowing ratio for this ooling sheme are lower, injetion onditions were limited here to about 2.0. Trying to improve the ooled area, two more holes were manufatured on the platform and tested under similar injetion onditions (Fig. 9-7 hole ooling sheme at M 1 about 1.1, 1.5 and 2.1 for both PS holes and new holes). The ooling sheme with two more holes show a non negligible inrease of the thermal protetion for the low injetion ondition. More oolant now reahes the sution side of the adjaent blade, even if again it is not allowed to ross the 3D separation lines. Its momentum is not enough to ounterat the passage vortex washing ativity, resulting in a quik deay. But as soon as M 1 is inreased up to 1.5 or even up to 2.1, jet lift off takes plae again, resulting in a lak of thermal protetion. In partiular, the last hole is haraterized by a so high oolant momentum that no trae of this jet on the wall an be seen, both for M 1 = 1.5 and 2.1. This is due again to the high injetion angle and to the hole loation very lose to the passage vortex three dimensional separation lines. The latter does not allow the oolant to travel aross the passage, resulting in an abrupt derease of effetiveness. Finally, it has to be noted that test duration was higher for the 5-hole ooling sheme, and partiularly for the low injetion ondition, resulting in a higher influene of heat ondution lose to hole loation. This explain the differenes between the effetiveness distribution downstream the PS holes. In order to better enlighten oolant jets behavior downstream of the new injetion holes, film ooling effetiveness data have been extrated along the interplatform line (see Fig. 3). Figure 10 ompares suh data belonging to the two tested hole numbers at variable M 1. Data are plotted against the S oordinate, i.e. a oordinate tangent to the inter-platform line, whose origin is at the filleted leading edge plane and that it is normalized using its value at the filleted trailing edge plane. The first hole is loated approximately at S/S TE = 0.14, while the fifth is at about 0.32 and the seventh at First onsidering the 5-hole onfiguration (Fig. 10a), aording with the previous analysis, best performane is attained at the lowest tested injetion rate of about 1.2. max values as large as 0.38 are reahed lose to the 5th hole, also due to a umulative effet. Inreasing the blowing ondition up to 1.6 max dereases down to about 0.2 as a maximum, also limiting its thermal protetion along the inter-platform separation line. Inreasing M 1 up to the maximum tested value of 2.0 results in a further redution of

7 Improving the Film Cooling of a Rotor Blade Platform - 6 a) a) b) b) ) Figure 8. Film ooling effetiveness distributions for the 5 hole sheme and variable injetion onditions. max to about 0.1, but also in a shift of the ooled region. The high momentum oolant seems to be less influened by the end wall ross flow, resulting in a oolant trae more aligned to the platform separation line. Moreover, the jet oming from the last hole is ompletely detahed from the wall, so ooling here is only due to the preeding holes. ) Figure 9. Film ooling effetiveness distributions for the 7 hole sheme and variable injetion onditions. Adding two holes along the same platform separation line slightly inreases the peak effetiveness at the lowest tested ondition (M 1 = 1.1 ) about 0.4. In fat, the same, or about the same M 1 value allows the 7-hole onfiguration to disharge a higher mass flow when ompared to the 5-hole sheme. This higher mass flow is disharged through the last two holes with a high momentum: this is

8 responsible for the high effetiveness levels observed where the last hole is loated (the 7th hole is loated at about S/S TE = 0.41). Inreasing the injetion ondition results in a progressive redution of thermal protetion that, for the largest tested M 1 of 2.1, is even worse than the 5-hole sheme (below 0.1). In fat the 7th hole does not ontribute to ool the platform. Whatever the injetion ondition, the 7- holes onfiguration allows to ool a slightly wider platform region, when ompared with the 5-hole sheme. Improving the Film Cooling of a Rotor Blade Platform - 7 Due to the very small injeted mass flow, no signifiant modifiations of seondary flows onfiguration and aerodynami losses take plae. Whatever the hole number, best thermal protetion is attained at low oolant blowing ratios M 1 of about 1.0. A blowing ratio inrease always result in a worsening of thermal protetion due to jet lift off from the wall aused by the high injetion angle. The 7-hole onfiguration allows the ooling of a larger platform region, extending up to the sution side of the following blade, up to about 0.5 ax. ACKNOWLEDGMENTS The authors wish to thank Ing. P. Epis and L. Padovan for their appreiated support. a) b) Figure 10. distributions along the inter-platform separation line: a) 5 holes and b) 7 holes onfiguration. As a final omment, both the investigated shemes get the best performane at a relatively low injetion rate of about M 1 = 1.0. The ooling sheme with 5 holes shows a lower ooled area, if ompared to the 7 hole sheme. Larger blowing rates always result in a loss of performane: the high injetion angle quikly promotes jet lift off. This is exatly the opposite of what happens rear on the hannel due to the PS ooling holes. Here, the low injetion angle (11.5 ) allows the high momentum fluid to remain attahed to the wall inreasing the blowing rate. CONCLUSIONS An experimental investigation on the thermal effets related to the introdution of new holes inside of the passage to ool the inter-platform gap region has been arried out. From the presented results the following onlusions an be drawn: REFERENCES [1] M.F. Blair. An Experimental Study of Heat Transfer and Film Cooling on Large-Sale Turbine Endwalls. J. Heat Transfer, 96: , [2] R.P. Roy, K.D. Squires, M. Gerendas, S. Song, W.J. Howe and A. Ansari. Flow and Heat Transfer at the Hub Endwall of Inlet Vane Passages Experiments and Simulations. ASME Paper 2000-GT-198, [3] R.A. Oke and T.W. Simon. Film Cooling Experiments with Flow Introdued upstream of a First Stage Nozzle Guide Vane through Slots of Various Geometries. ASME Paper 2002-GT-30169, [4] R.A. Oke, T.W. Simon, S.W. Burd and R. Vahlberg. Measurements in a Turbine Casade over a Contoured Endwall: Disrete Hole Injetion of Bleed Flow. ASME Paper 2000-GT-214, [5] M.J. Jabbari, K.C. Marston, E.R.G. Ekert and R.J. Goldstein. Film Cooling of the Gas Turbine Endwall by Disrete-Hole Injetion. J. of Turbomah, 118: , [6] S. Friedrihs, H.P. Hodson and W.N. Dawes. Distribution of Film-Cooling Effetiveness on a Turbine Endwall Measured with the Ammonia and Diazo Tehnique. J. Turbomah, 118: , [7] F. Kost and M. Niklas. Film-Cooled Turbine Endwall in a Transoni Flow Field: Part I Aero-dynami Measurements. ASME Paper 2001-GT-0145, [8] M. Niklas. Film-Cooled Turbine Endwall in a Transoni Flow Field: Part II Heat Transfer and Film-Cooling Effetiveness. ASME Paper 2001-GT- 0146, [9] D.G. Knost and K.A. Thole. Adiabati Effetiveness Measurements of Endwall Film-Cooling for a First Stage Vane. ASME Paper 2004-GT-53326, [10] G. Barigozzi, G. Benzoni, G. Franhini and A. Perdihizzi. Fan-shaped Hole Effets on the Aero- Thermal Performane of a Film Cooled Endwall. J. Turbomah. 128:43-52, 2006.

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