Calculations were performed to simulate the tip flow and heat transfer on the GE-E3 first-stage turbine, which represents a modern gas turbine blade geometry. Cases considered were a smooth tip, 2 percent recess, and 3 percent recess. In addition, a two-dimensional cavity problem was calculated. Good agreement with experimental results was obtained for the cavity calculations, demonstrating that the k–ω turbulence model used is capable of representing flows of the present type. In the rotor calculations, two dominant flow structures were shown to exist within the recess. Also areas of large heat transfer rate were identified on the blade tip and the mechanisms of heat transfer enhancement were discussed. No significant difference in adiabatic efficiency was observed for the three tip treatments investigated.

A conjugate heat transfer solver has been developed and applied to a realistic film-cooled turbine vane for a variety of blade materials. The solver used for the fluid convection part of the problem is the Glenn-HT general multiblock heat transfer code. The solid conduction module is based on the Boundary Element Method (BEM), and is coupled directly to the flow solver. A chief advantage of the BEM method is that no volumetric grid is required inside the solid — only the surface grid is needed. Since a surface grid is readily available from the fluid side of the problem, no additional gridding is required. This eliminates one of the most time consuming elements of the computation for complex geometries. Two conjugate solution examples are presented — a high thermal conductivity Inconel nickel-based alloy vane case and a low thermal conductivity silicon nitride ceramic vane case. The solutions from the conjugate analyses are compared with an adiabatic wall convection solution. It is found that the conjugate heat transfer cases generally have a lower outer wall temperature due to thermal conduction from the outer wall to the plenum. However, some locations of increased temperature are seen in the higher thermal conductivity Inconel vane case. This is a result of the fact that film cooling is a two-temperature problem, which causes the direction of heat flux at the wall to change over the outer surface. Three-dimensional heat conduction in the solid allows for conduction heat transfer along the vane wall in addition to conduction from outer to inner wall. These effects indicate that the conjugate heat transfer in a complicated geometry such as a film-cooled vane is not governed by simple one-dimensional conduction from the vane surface to the plenum surface, especially when the effects of coolant injection are included.

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A computational analysis of flow in simplex fuel atomizers using the arbitrary-Lagrangian-Eulerian method is presented. It is well established that the geometry of an atomizer plays an important role in governing its performance. We have investigated the effect on atomizer performance of four geometric parameters, namely, inlet slot angle, spin chamber convergence angle, trumpet angle, and trumpet length. For a constant mass flow rate through the atomizer, the atomizer performance is monitored in terms of dimensionless film thickness, spray cone half-angle, and discharge coefficient. Results indicate that increase in inlet slot angle results in lower film thickness and discharge coefficient and higher spray cone angle. The spin chamber converge angle has an opposite effect on performance parameters, with film thickness and discharge coefficient increasing and the spray cone angle decreasing with increasing convergence angle. For a fixed trumpet length, the trumpet angle has very little influence on discharge coefficient. However, the film thickness decreases, and spray cone angle increases with increasing trumpet angle. For a fixed trumpet angle, the discharge coefficient is insensitive to trumpet length. Both the spray cone angle and the film thickness are found to decrease with trumpet length. Analytical solutions are developed to determine the atomizer performance considering inviscid flow through the atomizer. The qualitative trends in the variation of film thickness at the atomizer exit, spray cone angle, and discharge coefficient predicted by inviscid flow analysis are seen to agree well with computational results. NomenclatureA a = air core area at orifice exit A o = orifice area A p = total swirl slot area A t = the trumpet end area A ta = air core area at the trumpet end C d = discharge coefficient,ṁ/A o (2 p/ρ) 0.5 D s = spin chamber diameter d o = orifice diameter d t = trumpet diameter at atomizer exit f = body force K = atomizer constant, A p /(D s d o ) K t = A p /(πr t r s ) K 1 = A p /(πr o r s ) L s = spin chamber length l o = orifice length l t = trumpet length p = static pressure Q = volume flow rate = radial distance from axis to inlet slot r o = orifice radius, d o /2 r s = spin chamber radius, D s /2 r t = d t /2 S(t) = surface enclosing control volume V (t) t = film thickness at exit t * = dimensionless film thickness, t/(d o /2) U = average total velocity at the end of orifice U = arbitrary velocity vector for the control volume V (t) u = axial-velocity component u = velocity vector u o= average axial velocity at the end of orifice u oa = axial velocity at the liquid-air interface at the end of orificē u t = average axial velocity at the end of trumpet u ta = axial velocity at the liquid-air interface at the end of trumpet V (t) = control volume w = tangential velocity component w i = average tangential velocity at the inlet w o = average tangential velocity at the end of orifice w oa = tangential velocity at the liquid-air interface at the end of orificē w t = average tangential velocity ...

Calculations were performed to simulate the tip flow and heat transfer on the GE-E3 first stage turbine, which represents a modern gas turbine blade geometry. Cases considered were a smooth tip, 2% recess, and 3% recess. In addition a two-dimensional cavity problem was calculated. Good agreement with experimental results was obtained for the cavity calculations, demonstrating that the k-ω turbulence model used is capable of representing flows of the present type. In the rotor calculations, two dominant flow structures were shown to exist within the recess. Also areas of large heat transfer rate were identified on the blade tip and the mechanisms of heat transfer enhancement were discussed. No significant difference in adiabatic efficiency was observed for the three tip treatments investigated.