Douglas C. Rabe

Role of Tip-Leakage Vortices and Passage Shock in Stall Inception in a Swept Transonic Compressor Rotor

The current paper reports on investigations aimed at advancing the understanding of the flow field near the casing of a forward-swept transonic compressor rotor. The role of tip clearance flow and its interaction with the passage shock on stall inception are analyzed in detail. Steady and unsteady three-dimensional viscous flow calculations are applied to obtain flow fields at various operating conditions. The numerical results are first compared with available measured data. Then, the numerically obtained flow fields are interrogated to identify the roles of flow interactions between the tip clearance flow, the passage shock, and the blade/endwall boundary layers. In addition to the flow field with nominal tip clearance, two more flow fields are analyzed in order to identify the mechanisms of blockage generation: one with zero tip clearance, and one with nominal tip clearance on the forward portion of the blade and zero clearance on the aft portion. The current study shows that the tip clearance vortex does not break down, even when the rotor operates in a stalled condition. Interaction between the shock and the suction surface boundary layer causes the shock, and therefore the tip clearance vortex, to oscillate. However, for the currently investigated transonic compressor rotor, so-called breakdown of the tip clearance vortex does not occur during stall inception. The tip clearance vortex originates near the leading edge tip, but moves downward in the spanwise direction inside the blade passage. A low momentum region develops above the tip clearance vortex from flow originating from the casing boundary layer. The low momentum area builds up immediately downstream of the passage shock and above the core vortex. This area migrates toward the pressure side of the blade passage as the flow rate is decreased. The low momentum area prevents incoming flow from passing through the pressure side of the passage and initiates stall inception. It is well known that inviscid effects dominate tip clearance flow. However, complex viscous flow structures develop inside the casing boundary layer at operating conditions near stall.Copyright

Experimental Investigation of Stepped Tip Gap Effects on the Performance of a Transonic Axial-Flow Compressor Rotor

The effects of stepped-tip gaps and clearance levels on the performance of transonic axial-flow compressor rotor were experimentally determined. A two-stage compressor with no inlet guide vanes was tested in a modern transonic compressor research facility. The first-stage rotor was unswept and was tested for an optimum tip clearance with variations in stepped gaps machined into the casing near the aft tip region of the rotor. Nine casing geometries were investigated consisting of three step profiles at each of three clearance levels. For small and intermediate clearances, stepped tip gaps were found to improve pressure ratio, efficiency, and flow range for most operating conditions. At 100 percent design rotor speed, stepped tip gaps produced a doubling of mass flow range with as much as a 2.0 percent increase in mass flow and a 1.5 percent improvement in efficiency. This study provides guidelines for engineers to improve compressor performance for an existing design by applying an optimum casing profile.

The Application of Pressure- and Temperature-Sensitive Paints to an Advanced Compressor

An innovative pressure-measurement technique that employs the tools of molecular spectroscopy has been widely investigated by the aerospace community. Measurements are made via oxygen-sensitive molecules attached to the surface of interest as a coating, or paint. The pressure-sensitive-paint (PSP) technique is now commonly used in stationary wind-tunnel tests; this paper presents the use of this technique in advanced turbomachinery applications. New pressure- and temperature-sensitive paints (P/TSPs) have been developed for application to a state-of-the-art transonic compressor where pressures up to 1.4 atm and surface temperatures to 90°C are expected for the suction surface of the first-stage rotor. PSP and TSP data images have been acquired from the suction surface of the first-stage rotor at 85 percent of the corrected design speed for the compressor near-stall condition. A comparison of experimental results with CFD calculations is discussed.

Experimental and Computational Investigation of Stepped Tip Gap Effects on the Flowfield of a Transonic Axial-Flow Compressor Rotor

The effects of stepped tip gaps and clearance levels on the flowfield of a transonic axial-flow compressor rotor were experimentally and computationally determined. This paper complements a previous experimental study by the authors regarding the effects of stepped tip gaps and clearance levels on the performance of an axial-flow compressor rotor. In the current study, the generation of blockage associated with the variation of geometry in the rotor tip region was examined. The shock-vortex interaction generating the blockage was characterized, and a theory and mechanism for relocation of blockage in the rotor tip region was developed. A two-stage compressor with no inlet guide vanes was tested in a modern transonic compressor research facility. The first-stage rotor was unswept and was tested with stepped gaps machined into the casing near the aft tip region of the rotor. Nine casing geometries were investigated consisting of three step profiles at each of three clearance levels. Computational Fluid Dynamic modeling of tip geometry effects also was performed. Increased tip clearance was found to increase the amount of flow blockage near the rotor tip. Stepped tip gaps were found to be an effective means of reducing the effects of tip region blockage, resulting in improved pressure ratio, efficiency, and mass flow. This study provides guidelines for engineers to improve compressor performance for an existing design by applying an improved casing profile.Copyright

Predicting Separation and Transitional Flow in Turbine Blades at Low Reynolds Numbers

There is increasing interest in design methods and performance prediction for aircraft engine turbines operating at low Reynolds numbers. In this regime, boundary layer separation may be more likely to occur in the turbine flow passages. For accurate CFD predictions of the flow, correct modeling of laminar-turbulent boundary layer transition is essential to capture the details of the flow. To investigate possible improvements in model fidelity, CFD models were created for the flow over two low pressure turbine blade designs. A new three-equation eddy-viscosity type turbulent transitional flow model originally developed by Walters and Leylek was employed for the current RANS CFD calculations. Previous studies demonstrated the ability of this model to accurately predict separation and boundary layer transition characteristics of low Reynolds number flows. The present research tested the capability of CFD with the Walters and Leylek turbulent transitional flow model to predict the boundary layer behavior and performance of two different turbine cascade configurations. Flows over the Pack-B turbine blade airfoil and the midspan section of a typical low pressure turbine (TLPT) blade were simulated over a Reynolds number range of 15,000–100,000, and predictions were compared to experimental cascade results. The turbulent transitional flow model sensitivity to turbulent flow parameters was investigated and showed a strong dependence on free-stream turbulence intensity with a second order effect of turbulent length scale. Focusing on the calculation of the total pressure loss coefficients to judge performance, the CFD simulation incorporating Walters and Leylek’s turbulent transitional flow model produced adequate prediction of the Reynolds number performance for the TLPT blade cascade geometry. Furthermore, the correct qualitative flow response to separated shear was observed for the Pack-B blade airfoil. Significant improvements in performance predictions were shown over predictions of conventional RANS turbulence models that cannot adequately model boundary layer transition.

A Mixing Plane Model Investigation of Separation and Transitional Flow at Low Reynolds Numbers in a Multistage Low Pressure Turbine

Flow separation with increased losses is known to occur when low pressure turbine (LPT) blades are operated at high altitudes with a reduced inlet Reynolds number. Under these conditions, boundary layer separation is more likely to be present within the flowfield of the LPT stages due to thickening of the boundary layers and an increase in the portion of the airfoil experiencing laminar flow. More accurate CFD predictions are needed in order to improve design methods and performance prediction for LPT stages operating at low Reynolds numbers. Steady flow CFD simulations of multistage LPT flow were completed at nominal and high altitude conditions with the conventional Spalart-Allmaras turbulence model. This model was used in combination with a mixing plane model for the simulation of flow through domains with one or more regions in relative rotational motion. Flow visualizations were completed using surface flow and streamline calculations to help identify vortical structures present within the flowfield. Also, the total pressure loss coefficient was calculated for each blade row. Qualitative comparisons indicate that the simulated high altitude condition had an increase in the amount of separated flow present within the flowfield compared to the nominal altitude condition. This can be attributed to the reduction in the inlet Reynolds number. Initial investigations with a recently-developed three-equation eddy-viscosity type turbulent transitional flow model are also reported. Comparisons of flow predictions for the 1st turbine stage with the two models revealed that large vortices predicted with the Spalart-Allmaras model were not present, and the wake loss coefficient was significantly lower with the three-equation turbulence model. Based on these and previous results, the CFD with the three-equation model is considered to have potential to provide improved prediction of separation and transitional flow in low Reynolds number turbine applications.

Turbulence Model Comparisons for Mixing Plane Simulations of a Multistage Low Pressure Turbine Operating at Low Reynolds Numbers

There has been a need for improved flow prediction methods for low pressure turbine (LPT) blades operating at high altitudes with a reduced inlet Reynolds number. These conditions present an increased amount of laminar-to-turbulent transitional flow within the boundary layers on the LPT blade surfaces. Also, boundary layer separation is more likely to occur within the flowfield of the LPT stages due to the lower freestream velocities in the regions of adverse pressure gradients on the suction surfaces. More accurate predictions of aerodynamic losses due to low Reynolds effects are needed for CFD to provide more accurate input to the design process for LPT stages operating at high altitudes. Steady flow CFD simulations of flow in a multistage LPT geometry were completed at nominal and high altitude conditions with the conventional Spalart-Allmaras turbulence model and a recentlydeveloped three-equation eddy-viscosity type transitional flow model. These models were used in combination with a mixing plane model for the simulation of flow through a three stage low pressure turbine. Flow visualizations were completed using surface flow and streamline calculations to help identify vortical structures present within the flowfield. Also, the total pressure loss coefficient was calculated for each blade row. Qualitative comparisons indicate the amount and the location of the flow separation differed significantly depending on the chosen turbulence model. Overall, the high altitude condition had an increased amount of separated flow compared to the nominal altitude condition resulting in an increase in the loss coefficient. The altitude effect on the laminar-to-turbulent transition location was studied using the three-equation model. The model provided a more detailed understanding of the aerodynamic loss mechanisms present in low Reynolds number flows, since it accounted for transitional boundary layer flow effects. Based on the these results, the CFD using the three-equation model has the potential to be a more effective method for turbine flow prediction at low Reynolds numbers compared to conventional RANS turbulence models.