This paper presents time-averaged and phase-resolved measurements of the surface pressure data for the vane and blade of a transonic single-stage research turbine. The data are compared and contrasted with predictions from an unsteady Euler/Navier–Stokes code. The data were taken in a shock-tunnel facility in which the flow was generated with a short-duration source of heated and pressurized air. Surf ace-mounted high-response pressure transducers were used to obtain the pressure measurements. The turbine was operating at the design flow function, the design stage pressure ratio, and 100 percent corrected speed. A matrix of data was obtained at two vane exit conditions and two vane/rotor axial spacings.
A comprehensive study has been performed to determine the influence of vane-blade spacing on transonic turbine stage aerodynamics. In Part I of this paper, an investigation of the effect of turbine vane–blade interaction on the time-mean airfoil surface pressures and overall stage performance parameters is presented. Experimental data for an instrumented turbine stage are compared to two- and three-dimensional results from four different time-accurate Navier–Stokes solvers. Unsteady pressure data were taken for three vane-blade row spacings in a short-duration shock tunnel using surface-mounted, high-response pressure sensors located along the midspan of the airfoils. Results indicate that while the magnitude of the surface pressure unsteadiness on the vane and blade changes significantly with spacing, the time-mean pressures and performance numbers are not greatly affected.
A multisweep space-marching solver based on a modified version of the SIMPLE algorithm was employed to study the three-dimensional flow field through a linear cascade. Three cases were tested: one with moderate loading, one with high loading, and one with high loading and tip clearance. The results of the numerical simulation were compared with available experimental measurements, and the agreement between the two was found satisfactory. The numerical simulation provided insight into several important endwall flow phenomena such as the interaction between the leakage and passage vortices, the interaction between the leakage vortex and the wake, the effect of leakage flow and loading on losses and secondary kinetic energy, the suction side corner separation, and the blowing of this separation by the leakage flow.
goal is the use of film-cooling, wherein a layer of relatively The overall objective of this study was to examine cool gas is iqjected near the metal surfaces to provide a the aerodynamics and heat transfer characteristics of dis-buffer layer between the hot gas and the protected sursite film-cooled turbine &rf0&. The ob-faces. As operational temperature lev& rise in modern jective was to attempt to predict the three-dimensional gas turbine engines, accurately controlling turbine coolflow about the C3X turbine -e m a d e with a lading ing flows presents one of the more difficult engineering edge showerhead film cooling arrangement. The moti-challenges in the overall turbomaehinery d e s b process. vation behind this work was to validate and assess the Turbine airfoil blade row flows are characterized by accuracy of present 3-D Navier-Stokes predictions for re-large temperature gradients, Ma& numbers ranging from ticf film-cooled airfoil heat transfer predictions through low subsonic (< 0.15) to the transonic range (> LO), and comparisons with experimental data. high levels of freestream turbulence with strong, small Calculations were performed using a Cartesian COOP scale interactions between surface boundary layer convecdinate system and taking advantage of the spanwise p e tive flows, diffusive transport, and turbulent shear transriodicity of the C3X geometry. The film cooling flow was port. Aerodynamic and thermal design techniques curmodeled as injection at discrete holes and was introduced rently available to turbine airfoil designers have deficieninto the blade passage through the use of separate mesh cies which do not permit a priori designs which meet systems and transpiration/iection boundary conditions. desired design goals without expensive experimental deThe grid generation for the C3X m a d e involved mod-velopment iterations. As such, the airfoil/eoolant flow eling the film cooling holes as discrete objects in a 3-D injection design (hole size, placement, shaping, etc) is ofmesh. Calculations were performed for the C3X at multi-ten based on experience and/or empirical databases. Inple operating points, both with and without coolhg holes c r d utilieationof computational fluid dynamics (CFD) activated. Predicted results were compared with experi-techniques in the design process for turbmomachinery airmental data.foils and flowpaths has naturally led to the use of these tools for predicting airfoil surface heat transfer and film + . , cooling effectiveness. Unfortunately, our lack of comprchensive turbulence models capable of accurately predict ing heat transfer in high Reynolds number turbulent flows The trend in thermodynamic design Of gas turbine en-has prevented widespread acceptance of CFD techdques in the heat design arena. The of CFD for details ofthe gas path/coolant flow interaction can be very useful for determining trends a better of
A 3-D Navier-Stokes analysis for turbomachinery flows on C- or O-type grids is presented. The analysis is based on the Beam and Warming implicit algorithm for solution of the unsteady Navier-Stokes equations and is derived from an early version of the ARC3D flow code developed at NASA Ames Research Center. The Navier-Stokes equations are written in a Cartesian coordinate system rotating about the z-axis, and then mapped to a general body-fitted coordinate system. All viscous terms are calculated and the turbulence effects are modelled using the Baldwin-Lomax turbulence model. The equations are discretized using finite differences on stacked body-conforming grids. Modifications made to convert ARC3D from external flow to internal turbomachinery flows and to improve solution accuracy are given in detail. The body-conforming grid construction procedure is also presented. Calculations for several rotor flows have been made, and results of code experimental verification studies are presented. Comparisons of the solutions obtained on the C- and O-type grids are also presented, with particular attention to shock resolution.
A rapid, time-marching, numerical scheme based on the hopscotch method is presented for solution of steady, two-dimensional, transonic flow in turbomachinery cascades. The scheme is applied to the strong-conservation form of the unsteady Euler equations written in arbitrary curvilinear coordinates. Cascade solutions are obtained on an orthogonal, body-centered coordinate system. Numerical solution results for two turbine cascades are presented and compared with experimental data to demonstrate the accuracy and computational efficiency of the analysis method.
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