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Optimization of Turbine Blade Cooling Using Combined Cooling Techniques a

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Mohammad Mehdi Zolfagharian , Mehran Rajabi-Zargarabadi , A S Mujumdar , M S a

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Valipour & Mojtaba asadollahi a

Department of Mechanical Engineering, Semnan University, P.O. Box 35131-191, Semnan, Iran b

Department of Mechanical Engineering, National University of Singapore, 9 Engineering Drive 1, 119260, Singapore c

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Department of Chemical and Bimolecular Engineering, Hong Kong University of Science and Technology, Clear Water Bay, Kowloon, Hong Kong Published online: 19 Nov 2014.

To cite this article: Mohammad Mehdi Zolfagharian, Mehran Rajabi-Zargarabadi, A S Mujumdar, M S Valipour & Mojtaba asadollahi (2014) Optimization of Turbine Blade Cooling Using Combined Cooling Techniques, Engineering Applications of Computational Fluid Mechanics, 8:3, 462-475, DOI: 10.1080/19942060.2014.11015529 To link to this article: http://dx.doi.org/10.1080/19942060.2014.11015529

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Engineering Applications of Computational Fluid Mechanics Vol. 8, No. 3, pp. 462–475 (2014)

OPTIMIZATION OF TURBINE BLADE COOLING USING COMBINED COOLING TECHNIQUES Mohammad Mehdi Zolfagharian^*, Mehran Rajabi-Zargarabadi^, AS Mujumdar #+ MS Valipour^ and Mojtaba asadollahi ^ ^

Downloaded by [108.61.197.7] at 22:49 26 August 2015

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Department of Mechanical Engineering, Semnan University, P.O. Box 35131-191, Semnan, Iran Department of Mechanical Engineering, National University of Singapore, 9 Engineering Drive 1, 119260, Singapore + Department of Chemical and Bimolecular Engineering, Hong Kong University of Science and Technology, Clear Water Bay, Kowloon, Hong Kong *E-Mail: [email protected] (Corresponding Author)

ABSTRACT: This paper presents analysis and optimization of turbine bade cooling systems. Since the temperature of combustion gases is very high sometimes reaching 2400 K, the turbine blade cannot sustain the resulting thermal stress. Moreover, for higher efficiency for advanced gas turbines, increase of inlet temperature is needed. Common blade cooling methods are film cooling, convection cooling, impingement cooling and combined cooling. In this paper, a numerical solution of the thermal and flow fields in film cooling technique on the AGTB expand symmetrical turbine blade was obtained and the results were validated with experimental data. Then the turbine blade geometry was changed and two combined cooling (impingement/convection cooing and impingement/film cooling) techniques were evaluated. The low Reynolds number k–ε turbulence model (AKN) was used for the turbulent flow simulations at various blowing ratios for two blade thicknesses. Comparisons of the results between the available experimental and numerical data showed that the AKN model is capable of predicting the turbulent flow and heat transfer in turbine blade cooling. Combined techniques (impingement/convection

cooling and impingement/film cooling) were also carried out and more cooling effectiveness and uniform temperature distribution were found than film cooling method only. Keywords: turbine blade, film cooling, impingement cooling, convection cooling, combined cooling, cooling effectiveness, turbulence modeling

their results with the available experimental data (Nemdili et al., 2008) and examined the performance of the SSG (Speziale, Sarkar, and Gatski) Reynolds Stress Model for the prediction of film cooling of a symmetrical blade. They found that the RSM model yields reasonably good agreement with measured data at low blowing ratios. Baheri et al. (2008) investigated the effect of mainstream turbulence intensity on the film cooling effectiveness. They reported that sensitivity of film cooling effectiveness to turbulence intensity decreases for trench shaped holes. Florschuetz et al. (1981) and Höglund (1999) evaluated experimentally and numerically the influence of the distance between the jet orifice and target plates (H/D) and jet-to-set spacing (L/D) on the averaged heat transfer. Their results indicated that for larger H/D, the heat transfer is decreased; with a further increase of H/D, the heat transfer is also decreased for an increased L/D. Wang and Mujumdar (2005) carried out a comparative study of the heat transfer under a

1. INTRODUCTION Increase of gas turbine thermal efficiency can be achieved by increasing the turbine inlet temperature. Despite considerable progress in blade metallurgy, such increase of gas temperature can only be afforded when the blade can be cooled effectively. Film cooling, convection cooling, impingement cooling and combined cooling are some of the most commonly used cooling techniques. Major effort has been devoted in recent years to developing these techniques. Haslinger and Hennecke (1997) investigated experimentally leading-edge film cooling of a symmetrical model of AGTB turbine blade. They showed that by increasing the blowing rate, the jets penetrate deeper into the main stream resulting in a decrease of the effectiveness. Film cooling of the same blade model, was investigated numerically by Lakehal et al. (2001). They predicted the effect of blowing rates on the flow and temperature fields and successfully compared Received: 3 Sep. 2013; Revised: 9 May 2014; Accepted: 23 May 2014 462


Engineering Applications of Computational Fluid Mechanics Vol. 8, No. 3 (2014)

curved effusion wall of combustion chamber. Their results showed that with decrease of the effusion hole angle and effusion hole-hole spacing, cooling effectiveness is enhanced. Sensitivity of the overall effectiveness of film cooling and internal cooling of a turbine vane suction side was investigated experimentally by Williams et al. (2013). Overall cooling effectiveness and adiabatic film effectiveness were measured downstream of a single row of round holes positioned on the suction side of the vane. According to their results, while the adiabatic film effectiveness decreased when using high momentum flux ratios for the film cooling, due to coolant jet separation, the overall cooling effectiveness increased at higher momentum flux ratios. Overall effectiveness of a blade endwall with jet impingement and film cooling was also investigated experimentally by Mensch and Thole (2013). In their work, a conjugate heat transfer model for the endwall of a seven-blade cascade was developed to examine the impact of both convective cooling and solid conduction through the endwall. Experiments with only film cooling showed high effectiveness around film-cooling holes due to convective cooling within the holes. Internal impingement cooling provided more uniform effectiveness than film cooling. A few experimental and numerical studies have been performed for combined cooling techniques, but almost all of these numerical investigations have been carried out for flat plate. The present work investigates numerical simulation of two combined impingement/ convection cooling (CICC) and combined impingement/film cooling (CIFC) techniques over the AGTB symmetrical turbine blade geometry. Film cooling for this geometry studied experimentally by Haslinger and Hennecke, (1997) and numerically by Haslinger and Hennecke (1997), Lakehal et al. (2001), Nemdili et al. (2008) and Baheri et al. (2008). In the present study firstly, numerical solution of heat transfer and flow field for film cooling over the AGTB symmetrical turbine blade was obtained. The results were validated with relevant experimental data. By changing the geometry, two CICC and CIFC techniques were evaluated numerically. Flow field, temperature field and cooling effectiveness were calculated for different blowing ratios ranging from 0.3 to 0.7. Finally, by changing the blade thickness, the effect of conduction heat transfer (blade thickness) on the cooling effectiveness was also investigated.

turbulent slot jet using five low Reynolds number k–ε models. They compared numerical results with the available experimental data (Van Heiningen, 1982). Their results showed that the Yap correction, which decreases the turbulence length scale in the near wall region, improves the predicted local Nusselt number in good agreement with the experimental data in both stagnation and further downstream regions. Liu et al. (2013) reported detailed impingement heat transfer measurements on grooved surfaces using the well-known transient liquid crystal method. Their results showed that heat transfer is enhanced near grooves. However, the grooves directly underneath the jet holes hinder the impingement ﬂow and the impingement heat transfer is reduced locally. The effect of impingement cooling on the cooling of the leading edge region in a commercial turbine high pressure first stage rotor blade was investigated using computational fluid dynamic simulations for various cases which are different from those of Rajamani et al. (2013). In one case, passage was smoothened at the apex to reduce the dead zone and to enhance spreading of the jet. This results in 3% increase in Nuavg. Also in another case the coolant flow passage was further modified to improve the spreading of flow and a 5% increase in the Nuavg was observed as compared to base case. Many researchers have studied impingement cooling and film cooling flows. However, investigations on the flow and heat transfer characteristics of combined cooling techniques have not been widely considered. This work aims to make a contribution to this area. Halila et al. (1982) investigated an experimental model of combined impingement/film cooling technique. They reported qualitatively that the overall cooling effectiveness is enhanced in this combined model. Immarigeon and Hassan (2006) evaluated numerically an advanced impingement/film cooling scheme. Their results indicated that greater airfoil protection is provided by combined impingement/film cooling than traditional film cooling. Zhang et al. (2009), investigated numerically the flow and heat transfer characteristics of combined impingement/effusion cooling. They calculated the film cooling effectiveness for different blowing ratios and found that wall cooling effectiveness increases as the center- to-center spacing of adjacent holes decreases. Yang et al. (2011) investigated experimentally the impingement - effusion cooling behavior on a 463


blade stagnation line. The air is ejected out through the film cooling hole at S/D = 3.1 to the blade surface to create an air film to protect the surface from the hot combustion gases as shown in Fig. 2. Boundary conditions were taken according to the formulation of Nemdili et al. (2008). At the inflow boundary, a uniform stream wise velocity profile was assumed with U∞ = 30 m/s with T∞=293. The Reynolds number based on D and U∞ was Re = 7950. Uniform distributions were also specified for k and ε, corresponding to a free-stream turbulence intensity of 0.5% and dimensionless eddy viscosity of μt /μ = 30. Similarly, a uniform velocity profile was set at the inlet of the discharge pipe. Here as well uniform distributions of k and ε were specified, based on 3% and a length scale of k3/2/ε = 0.3D. The injection flow velocity was computed according to the blowing ratio values. The outflow boundary condition was set at constant zero relative pressure, while on the blade surface and the pipe internal walls, adiabatic wall, k = 0 and no-slip conditions were employed. The density ratio ρJet/ρ∞ and temperature ratio TJet/T∞ were approximately equal to one.

Fig. 1 Film cooling model blade geometry (Haslinger and Hennecke, 1997).


2. PROBLEM STATEMENT 2.1 Film cooling geometry and boundary condition The physical model for film cooling of the turbine blade used in the present study, has been studied experimentally by Haslinger and Hennecke (1997). The geometry of this blade is shown in Fig. 1.This blade is symmetrical with a length of 515 mm and a maximum width of 72 mm. Fig. 2 displays the computational domain and related boundary condition as well as a schematic view of the present film cooling model. Due to geometric and physical symmetry, simulations were carried out only for the flow field within the half domain. The film hole length-to-diameter ratio, L/D, was held fixed at 4. In the streamwise direction, the holes are inclined at angle of 110 deg. to the surface and located so that the trailing edge of the hole is at S/D = 3.1, where S is the length along the blade from the stagnation point, and S/D = 0 is the

2.2 Combined cooling geometry and boundary condition As use of impingement and convection cooling instead of film cooling leads to avoiding the mixing of the coolant and main stream flow (see Fig. 3a), and also due to the high heat transfer rate for impinging jet, the geometry was altered; a new cooling model (CICC) which includes both impingement and convection cooling, instead of simple film cooling was

Fig. 2 Boundary condition (Left) and schematic view (Right) of film cooling. 464



Fig. 3 Boundary conditions (Left) and schematic view (Right) of a) Impingement/Convection cooling b) Impingement/Film cooling system.

evaluated in present work. Fig. 3a displays a schematic view of the computational domain and the boundary condition employed in the present CICC model. It comprises a jet impingement hole, an impingement/convection cooling cavity and mainstream flow. As shown in Fig. 3a, the cooling stream is injected through the impinging hole into the impingement/convection cooling cavity when it impinges on the leading edge of blade. It then flows along the impingement/convection cooling cavity. In this cavity heat is transferred by conduction through the blade, and then by convection from the air flowing inside of the blade. Eventually hot gases enter into the main stream through the end blade holes (See Fig. 3a). The circular jet impingement hole diameter, D, is 4 mm that is identical to the film hole diameter in film cooling technique. The distance between the impingement hole and the blade is H~3D, when blade thickness is considered t=0 and H~2.5D when t= D/2. All boundary conditions

in this geometry are identical to those for the film cooling geometry (Fig. 3a). The side planes of the mainstream channel, hole and cooling cavity were considered as symmetry boundary condition. Boundary condition at the inlet of the discharge pipe for the CICC technique was the same as that for film cooling. Fig. 3b displays the computational domain and boundary condition of the present CIFC model. The computational domain and all boundary conditions for CIFC are identical to those used for simulating CICC. In the CIFC technique, impingement cooling was used at the leading edge of the AGTB turbine blade and combined with film cooling in the rest of the blade to compensate for the reduction of cooling efficiency downstream of the impingement zone. The region downstream of the impingement zone was divided into approximately 3 equal zones, 2.5

Engineering Applications of Computational Fluid ...

Engineering Applications of Computational Fluid ...

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