Mark C. Morris

3D CFD Ingestion Evaluation of a High Pressure Turbine Rim Seal Disk Cavity

The amount of cooling air assigned to seal high pressure turbine rim cavities is critical for performance as well as component life. Less air leads to excessive hot annulus gas ingestion and its penetration deep into the cavity compromising disk life. Excessive purge air adversely affects performance. The minimum purge (i.e. sealing) air requirement to control ingestion is also influenced by annulus circumferential pressure fluctuation present over the rim seal cavity. Its interaction with the platform gap resistance and the amount of purge air needs to be understood in order to reliably predict performance and component life. Work has commenced to investigate opportunities in reducing disk cavity purge flow requirements by studying ways to control ingestion. The study has been initiated with 3D CFD model setup/run mode options to benchmark main/cavity flow field interactions. The selection of the appropriate CFD model fidelity, however, is one of the main goals of this work. The CFD model phase has 3 options to be evaluated; 1) steady solution with mixing plane aft of the cavity, 2) steady solution with mixing plane forward of the cavity, 3) unsteady solution. Option 1 has been completed and is the subject of this paper. A reference HP turbine stage and disk cavity from an engine design was selected for the CFD study. The steady flow solution model captured the oscillatory movement and penetration depth of ingestion by varying purge flow rate and observing the impact on the mixing plane forward and aft of the disk cavity. Moreover, the influence of upstream stator vane airfoil fillet shape was also investigated. The entrained flow was established by starving the cavity and integrating the outflow along the disk. This value along with the nominal and intermediate cavity purge flows were validated against relevant sealing flow design correlations. At a radial location near the rim, an ingestion mixing efficiency value versus purge flow rate was obtained which correlates well with recent unsteady flow results from the literature.Copyright

Review of Platform Cooling Technology for High Pressure Turbine Blades

With the relatively large surface area of the platform of the gas turbine blades being exposed directly to the hot, mainstream gas, it is vital to efficiently cool this region of the blades. This region is particularly difficult to protect due to the strong secondary flows developed at the airfoil junction (formation of the leading edge horseshoe vortex) and circumferentially across the blade passage (strengthening passage vortex moving from the pressure side to the suction side of the passage). Over the past decade, researchers and engine designers have attempted to combat the enhanced heat transfer to the blade platform by implementing both frontside and backside novel cooling techniques. This paper presents a review of platform cooling technology ranging from frontside film cooling via stator-rotor purge flow, mid-passage purge flow, and discrete film holes to backside cooling achieved via impinging jet arrays or cooling channels. To gain a full understanding of state-of-the-art cooling technology, recent patents, journal articles, and conference proceedings are included in this review.Copyright

Combined Effects of Jet Plate Thickness and Fillet Radius on Leading Edge Jet Impingement With Round and Racetrack Shaped Jets

The effect of jet plate thickness is considered as regionally averaged Nusselt numbers are measured on a concave surface, which models the leading edge of modern gas turbine blades. The performance of both round and racetrack shaped orifices for leading edge impingement is considered. Regionally averaged heat transfer coefficient distributions are obtained in a steady state experiment using heated aluminum plates. From this traditional heat transfer technique, the heat transfer afforded by jet plates of varying thickness is quantified. The thickness of the jet plate is varied from 1.33 to 4.0 diameters (for both the round and racetrack shaped jets). To model the modern, cast airfoil, the effect of an inlet and outlet radius on the jet orifice is also investigated. For all cases, the jet – to – target surface spacing (z/djet) is 4, the jet – to – jet spacing (s/djet) is 8, and the target surface diameter – to jet diameter (D/djet) is 5.33. Target surface Nusselt numbers are obtained for three separate Reynolds numbers. For the round orifices, jet Reynolds numbers of 14,000, 28,100, and 42,100 are used while the corresponding Reynolds numbers for the racetrack shaped jets are 11,800, 23,600, and 35,400. Although the Reynolds number is reduced for the racetrack shaped jets, the mass flow through each jet remains constant (from the round to the racetrack jets). The Nusselt numbers measured in the stagnation region of the target surface are relatively insensitive to the jet plate thickness. For all cases considered, the flow is not developed as it exits the orifice, so the flow structures of the jets ensuing from each of the plates are similar. When the inlet of the jet is rounded, the vena contracta effect within the orifice is minimized, and a more symmetrical jet develops within the orifice. For a fixed flow rate, the racetrack shaped jets provide enhanced heat transfer compared to the round jets for all geometries considered.Copyright

High Pressure Turbine Low Radius Radial TOBI Discharge Coefficient Validation Process

A TOBI (tangential on board injection) or preswirl system is a critical component of a high pressure turbine cooling delivery system. Its efficient performance and characterization are critical because the blade and disk life depend on the accuracy of delivering the required flow at the correct temperature and pressure. This paper presents a TOBI flow discharge coefficient validation process applied to a low radius radial configuration starting from a 1D network flow analysis to a 3D frozen rotor CFD analysis of the rotor cooling air delivery system. The analysis domain commences in the combustor plenum stationary reference frame, includes the TOBI, transitions to the rotating reference frame as the flow travels through the rotating cover plate orifice, continues up the turbine disk into the slot bottom blade feed cavity, and terminates in the turbine blade. The present effort includes matching a 1D network model with 3D CFD results using simultaneous goal-matching of pressure predictions throughout the circuit, defining test rig pressure measurements at critical “non-disturbing” locations for quanification of pressure ratio across the TOBI, and finally comparing the TOBI flow coefficient resulting from stationary cold flow tests with what was obtained from the 3D CFD results. Analysis of the results indicates that the discharge coefficient varies with pressure ratio and that the traditional method of using a constant discharge coefficient extracted from a cold flow test run under choked conditions leads to over-predicting turbine cooling flows. TOBI flow coefficient prediction for the present study compares well with the stationary data published by author researchers for the configuration under investigation, and the process described in this paper is general for any TOBI configuration.Copyright