M. A. MUJEEBU, M.K. ABDULLAH, M.Z. ABDULLAH, K.A. AHMAD, S. F. SHAKER

Study on the effect of number of film cooling rows on the thermal performance of gas turbine blade

Bu çalışmada, ticari bir hesaplamalı Akışkanlar Dinamiği (HAD) kodu kullanılarak, gaz türbini kanatçıklarının termal davranışına film soğutmanın etkisi üç boyutlu (3D) sayısal olarak incelenmiştir. Gerçek bir A-4 Skyhawk kanatçığın üç boyutlu geometrisi GAMBIT ön-prosesörü ile oluşturulmuştur. İki farklı soğutma konfigürasyonu: 1) U- dirsekli iç kanallı dört sıra film soğutma; 2) U- dirsekli iç kanallı sekiz sıra film soğutma için türbo-spesifik yansıtmasız sınır şartları altında bir türbin kanatçığı üzerindeki transonik akış simule edilmiştir. Türbülans modeli olarak kayma gerilmesi aktarımı (SST) model kullanılmıştır. Akışın serbest-akış türbülans şiddeti 9% kabul edilmiştir. Isı transfer katsayısı, toplam sıcaklık dağılımı, statik basınç ve hız vektörlerinin değişimi incelenmiştir. Sıcaklık dağılımına, soğutma püskürme basınç oranın $(P_{R,ci})$etkisi de incelenmiştir. Film soğutmalı ısı transfer katsayısının film soğutmasız durumdan daha yüksek olduğu görülmüştür. Sıcaklı profillerinden, U- dirsekli iç kanallı sekiz sıra film soğutmalı kanatçığın U- dirsekli iç kanallı dört sıra film soğutmalı kanatçığa göre daha iyi soğutma performansı sergilediği görülmüştür. Püskürme basınç oranının $(P_{R,ci})$daha fazla artırılması sıcaklıkta daha fazla düşüşe sebep olmuştur ve ayrıca yanal püskürtmenin en iyi soğutma tabakası oluşturduğu görülmüştür.

Gaz türbini kanatçığının termal performansına film soğutma kademe sayısının etkisinin incelenmesi

This paper presents three dimensional (3D) numerical investigations on the effect of film cooling on the thermal behavior of gas turbine blades, using a commercial computational fluid dynamics (CFD) code. The 3D airfoil geometry of the blade which emulates the actual (A-4 Skyhawk) blade is generated in the pre-processor (GAMBIT). Two cooling configurations namely 1) four rows film cooling with U-bend internal channel and 2) eight rows film cooling with U-bend internal channel, have been simulated to be transonic flow over a turbine blade with turbo-specific non-reflecting boundary conditions (NRBCs). Turbulence is represented using the shear-stress transport (SST) model, and the flow is assumed to have a free-stream turbulence intensity of 9%. The heat transfer coefficient, total temperature distribution, static pressure and velocity vector are investigated. The effect of coolant injection pressure ratio on$(P_{R,ci})$ temperature distribution is also investigated. The results show that heat transfer coefficient with film cooling is higher than that without film cooling. From the predicted temperature profile, it is observed that the blade with eight rows film cooling with U-bend internal channel shows better cooling performance than that with four rows. Further, increase in$(P_{R,ci})$ leads to reduction in temperature and moreover the lateral spreading facilitated the best coolant layer.

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