Theodore H. Okiishi

Boundary Layer Development in Axial Compressors and Turbines: Part 3 of 4— LP Turbines

This is Part Three of a four-part paper. It begins with Section 11.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this part, we present the experimental evidence that we used to construct the composite picture for LP turbines that was given in the discussion in Section 5.0 of Part 1. We present and interpret the data from the surface hot-film gages and the boundary layer surveys for the baseline operating condition. We then show how this picture changes with variations in Reynolds number, airfoil loading, and nozzle-nozzle clocking.

Boundary Layer Development in Axial Compressors and Turbines: Part 1 of 4—Composite Picture

Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading. Part 1 of this paper summarizes all of our experimental findings by using sketches to show how boundary layers develop on compressor and turbine blading. Parts 2 and 3 present the detailed experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data. Readers not interested in experimental detail can go directly from Part 1 to Part 4. For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region, which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level, and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien-Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment. Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.

Recommendations for Achieving Accurate Numerical Simulation of Tip Clearance Flows in Transonic Compressor Rotors

The tip clearance flows of transonic compressor rotors are important because they have a significant impact on rotor and stage performance. A wall-bounded shear layer formed by the relative motion between the overtip leakage flow and the shroud wall is found to have a major influence on the development of the tip clearance flow field. This shear layer which has not been recognized by earlier investigators, impacts the stable operating range of the rotor. Simulation accuracy is dependent on the ability of the numerical code to resolve this layer. While numerical simulations of these flows are quite sophisticated, they are seldom verified through rigorous comparisons of numerical and measured data because these kinds of measurements are rare in the detail necessary to be useful in high-speed machines. In this paper we compare measured tip-clearance flow details (e.g., trajectory and radial extent) with corresponding data obtained from a numerical simulation. Laser-Doppler Velocimeter (LDV) measurements acquired in a transonic compressor rotor, NASA Rotor 35, are used. The tip clearance flow field of this transonic rotor is simulated using a Navier-Stokes turbomachinery solver that incorporates an advanced k-e turbulence model derived for flows that are not in local equilibrium. A simple method is presented for determining when the wall-bounded shear layer is an important component of the tip clearance flow field.

An experimental investigation of the flow in a diffusing S-duct

Compressible, subsonic flow through a diffusing S-duct has been experimentally investigated. Benchmark aerodynamic data are presented for flow through a representative S-duct configuration. The collected data would be beneficial to aircraft inlet designers and is suitable for the validation of computational codes. Measurements of the 3D velocity field and total and static pressures were obtained at five cross-sectional planes. Surface static pressures and flow visualization also helped to reveal flow field characteristics. All reported tests were conducted with an inlet centerline Mach number of 0.6 and a Reynolds number, based on the inlet centerline velocity and duct inlet diameter, of 2.6 x 10(exp 6). The results show that a larger region of streamwise flow separation occurred within the duct. Details about the separated flow region, including mechanisms which drive this complicated flow phenomenon, are discussed. Transverse velocity components indicate that the duct curvature induces strong pressure driven secondary flows, which evolve into a large pair of counter-rotating vortices. These vortices convect the low momentum fluid of the boundary layer towards the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

Boundary Layer Development in Axial Compressors and Turbines: Part 2 of 4—Compressors

This is Part Two of a four-part paper. It begins with Section 6.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this part, we present the experimental evidence used to construct the composite picture for compressors given in the discussion in Section 5.0 of Part 1. We show the data from the surface hot-film gages and the boundary layer surveys, give a thorough interpretation for the baseline operating condition, and then show how this picture changes with variations in Reynolds number, airfoil loading, frequency of occurrence of wakes and wake turbulence intensity. Detailed flow features are described using raw time traces. The use of rods to simulate airfoil wakes is also evaluated.

Boundary layer development in axial compressors and turbines: Part 4 of 4 : Computations and analyses

This is Part Four of a four-part paper. It begins with Section 16.0 and concludes the description of the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this paper, the computational predictions made using several modern boundary layer codes are presented. Both steady codes and an unsteady code were evaluated. The results are compared with time-averaged and unsteady integral parameters measured for the boundary layers. Assessments are made to provide guidance in using the predictive codes to locate transition and predict loss. Conclusions from the computational analyses are then presented.

Study of the compressible flow in a diffusing S-duct

Benchmark aerodynamic data are presented for compressible flow through a representative S-duct configuration. Measurements of the three-dimensional velocity field, total pressures, and static pressures were obtained in five cross-sectional planes. Surface static pressure and surface flow visualization data were also acquired. All reported tests were conducted with an inlet centerline Mach number of 0.6. The Reynolds number, based on the inlet centerline velocity and duct inlet diameter, was 2.6 x 10 6. Thin inlet turbulent boundary layers existed. The collected data should be beneficial to aircraft inlet designers and are suitable for the validation of computational codes. The results show that a region of streamwise flow separation occurred within the duct. Measurements indicate that the duct curvature induced strong pressure-driven secondary flows. The crossflows evolved into counter-rotating vortices. These vortices convected low momentum fluid of the boundary layer toward the center of the duct, degrading both the uniformity and magnitude of the total pressure profile.

Stator-Rotor Interactions in a Transonic Compressor— Part 1: Effect of Blade-Row Spacing on Performance

Usually less axial spacing between the blade rows of an axial flow compressor is associated with improved efficiency. However, mass flow rate, pressure ratio, and efficiency all decreased as the axial spacing between the stator and rotor was reduced in a transonic compressor rig. Reductions as great as 3.3% in pressure ratio, and 1.3 points of efficiency were observed as axial spacing between the blade rows was decreased from far apart to close together. The number of blades in the stator blade-row also affected stage performance. Higher stator blade-row solidity led to larger changes in pressure ratio efficiency, and mass flow rate with axial spacing variation. Analysis of the experimental data suggests that the drop in performance is a result of increased loss production due to blade-row interactions. Losses in addition to mixing loss are present when the blade-rows are spaced closer together. The extra losses are associated with the upstream stator wakes and are most significant in the midspan region of the flow.

Stator-Rotor Interactions in a Transonic Compressor: Part 2 — Description of a Loss Producing Mechanism

A previously unidentified loss producing mechanism resulting from the interaction of a transonic rotor blade-row with an upstream stator blade-row is described. This additional loss occurs only when the two blade rows are spaced closer together axially. Time-accurate simulations of the flow and high-response static pressure measurements acquired on the stator blade surface reveal important aspects of the fluid dynamics of the production of this additional loss. At close spacing the rotor bow shock is chopped by the stator trailing edge. The chopped bow shock becomes a pressure wave on the upper surface of the stator that is nearly normal to the flow and that propagates upstream. In the reference frame relative to this pressure wave, the flow is supersonic and thus a moving shock wave that produces an entropy rise and loss is experienced. The effect of this outcome of blade-row interaction is to lower the efficiency, pressure ratio, and mass flow rate observed as blade-row axial spacing is reduced from far to close. The magnitude of loss production is affected by the strength of the bow shock and how much it turns as it interacts with the trailing edge of the stator. At far spacing the rotor bow shock degenerates into a bow wave before it interacts with the stator trailing edge and no significant pressure wave forms on the stator upper surface. For this condition, no additional loss is produced.Copyright

An Investigation of Wake-Shock Interactions in a Transonic Compressor With Digital Particle Image Velocimetry and Time-Accurate Computational Fluid Dynamics

The effects of varying axial gap on the unsteady flow field between the stator and rotor of a transonic compressor stage are important because they can result in significant changes in stage mass flow rate, pressure rise, and efficiency. Some of these effects are analyzed with measurements using digital particle image velocimetry (DPIV) and with time-accurate simulations using the 3D unsteady Navier-Stokes computational fluid dynamics solver TURBO. Generally there is excellent agreement between the measurements and simulations, instilling confidence in both. Strong vortices of the wake can break up the rotor bow shock and contribute to loss. At close spacing vortices are shed from the trailing edge of the upstream stationary blade row in response to the unsteady, discontinuous pressure field generated by the downstream rotor bow shock. Shed vortices increase in size and strength and generate more loss as spacing decreases, a consequence of the effective increase in rotor bow shock strength at the stationary blade row trailing edge. A relationship for the change in shed vorticity as a function of rotor bow shock strength is presented that predicts the difference between close and far spacing TURBO simulations.