Walter F. O’Brien

An Improved Streamline Curvature Approach for Off-Design Analysis of Transonic Axial Compression Systems

Abstract : A streamline curvature (SLC) throughflow numerical model was assessed and modified to better approximate the flow fields of highly transonic fans typical of military fighter applications. Specifically, improvements in total pressure loss modeling were implemented to ensure accurate and reliable off-design performance prediction. The assessment was made relative to the modeling of key transonic flow field phenomena, and provided the basis for improvements, central to which was the incorporation of a physics-based shock loss model. The new model accounts for shock geometry changes, with shock loss estimated as a function of inlet relative Mach number, blade section loading (flow turning), solidity, leading edge radius, and suction surface profile. Other improvements included incorporation of loading effects on the tip secondary loss model, use of radial blockage factors to model tip leakage effects, and an improved estimate of the blade section incidence at which minimum loss occurs. Data from a single-stage, isolated rotor and a two-stage, advanced-design (low aspect ratio, high solidity) fan provided the basis for experimental comparisons. The two-stage fan was the primary vehicle used to verify the present work. Results from a three-dimensional, steady, Reynolds-averaged Navier-Stokes model of the first rotor of the two-stage fan were also used to compare with predicted performance from the improved SLC representation. In general, the effects of important flow phenomena relative to off-design performance of the fan were adequately captured. These effects included shock loss, secondary flow, and spanwise mixing. Most notably, the importance of properly accounting for shock geometry and loss changes with operating conditions was clearly demonstrated. The majority of the increased total pressure loss with loading across the important first-stage tip region was shown to be the result of increased shock loss, even at part-speed.

Separated Flow in a Low Speed Two-Dimensional Cascade: Part I Flow Visualization and Time-Mean Velocity Measurements

This study was conducted for the purpose of providing a more fundamental understanding of separated flow in cascades and to provide performance data for fully stalled blade rows. Cascades of a single blade geometry and a solidity of unity were studied for three stagger angles and the full range of angle of attack, extending well into the stalled flow regime. Results are presented from flow visualization and time-mean velocity measurements of stalled flow in the cascade. Surface and smoke flow visualization revealed that the blade stagger angle is a key parameter in determining the location of the separation line and the occurrence of propagating stall. Time-mean velocity measurements obtained with a dual hot split-film probe also showed that the separated velocity profiles within the blade passages and the profiles in the wake have distinctly different characteristics depending on the stagger angle.

A Shock Loss Model for Supersonic Compressor Cascades

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier-Stokes analyses of the L030-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

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 Three-Dimensional Turbine Engine Analysis Compressor Code (TEACC) for Steady-State Inlet Distortion

The direct approach of modeling the flow between all blade passages for each blade row in the compressor is too computationally intensive for practical design and analysis investigations with inlet distortion. Therefore a new simulation tool called the Turbine Engine Analysis Compressor Code (TEACC) has been developed. TEACC solves the compressible, time-dependent, 3D Euler equations modified to include turbomachinery source terms which represent the effect of the blades. The source terms are calculated for each blade row by the application of a streamline curvature code. TEACC was validated against experimental data from the transonic NASA rotor, Rotor 1B, for a clean inlet and for an inlet distortion produced by a 90-deg, one-per-revolution distortion screen. TEACC revealed that strong swirl produced by the rotor caused the compressor to increase in loading in the direction of rotor rotation through the distorted region and decrease in loading circumferentially away from the distorted region.Copyright

A CFD Analysis of the Effects of Two-Phase Flow in a Two-Stage Centrifugal Compressor

The performance of a compressor is known to be affected by the ingestion of liquid droplets. Heat, mass and momentum transfer as well as the droplet dynamics are some of the important mechanisms that govern the two-phase flow. This paper presents numerical investigations of three-dimensional two-phase flow in a two-stage centrifugal compressor, incorporating the effects of the above mentioned mechanisms. The results of the two-phase flow simulations are compared with the simulation involving only the gaseous phase. The implications for the compressor performance, viz. the pressure ratio, the power input and the efficiency are discussed. The role played by the droplet-wall interactions on the rate of vaporization, and on the compressor performance is also highlighted.Copyright

Performance Modeling of a Power Generation Gas Turbine With Wet Compression

In the last 15 years more than 1000 power generation gas turbines have been modified with an OEM or aftermarket module to generate the wet compression phenomenon where “Hot Day” conditions are present on the site. This modification to the gas turbine increases power, but can produce performance problems including reduced compressor surge margin and possibly a shorter maintenance cycle because of resulting problems present in the compressor such as blade vibration and erosion with impingement of water droplets on the surface of the compressor blades[1]. In the last few years researchers in academia and the private sector have worked to understand the principles behind the wet compression process in order to know in depth how to use the application to best advantage with gas turbines. The main areas of the research on wet compression are thermodynamic analyses, computer fluid dynamic analysis, and the use of operational data. Because present technology is unable to obtain detailed operational data on the evaporation process within the compressor, researchers rely on computer simulations based upon aerothermodynamics and physical measurements of the gas turbines, and assumptions based upon available information. These computer simulations are typically aimed toward explaining the performance data from a specific gas turbine model. Most of these computer simulations are cycle analyses of the gas turbine [2–7], although a few are CFD analyses for a specific compressor using either in-house computer programs or commercial CFD software [8–10]. CFD analysis takes into account the fact that an evaporation model should be used in order to predict how the evaporation of the water droplets occurs through the stages of the compressor. Many of the CFD simulations that have been performed for wet compression assume that the mixture of air, liquid water, and water vapor is at equilibrium throughout the compressor. Also, a single water droplet size is sometimes used for the simulation instead of a size distribution for the droplets. These assumptions simplify the calculations for the software. The results of these simulations may over-forecast the effect of the wet compression and the power output of the gas turbine because of incorrect predictions of evaporation models, or because of the lack of a proper droplet size distribution affecting the calculation. An analysis that properly forecasts the power output of a gas turbine with wet compression is important for design, performance prediction, and operation. The intention of this paper is to show how performance predictions for a power generation gas turbine is affected by applying several evaporation models [2, 4, 5, 7] in a gas turbine model with a detailed, stage-by-stage compressor model. Model predictions are compared with available operational performance data. Conclusions are provided regarding the best evaporation model assumptions for accurate predictions of gas turbine performance with wet compression.Copyright

Assessing Fan and Compressor Blade HCF Sensitivity to Non-Uniform Inlet Flows

In recent years, the subject of high cycle forced response excitation of aircraft turbine engine components and resulting fatigue failures (HCF) has been of increasing concern. Prediction and test methods are needed that can identify inlet distortion-related HCF drivers and response, and support assessment and correction of potential problems before failure occurs. The present paper describes a method by which inlet distortions may be examined for flow-distortion-driven HCF excitation potential in a test situation. The method was derived from results of a distortion screen test that produced periodic excitation in a two-stage transonic fan. A CFD method is used to predict the unsteady pressure field on the first stage blades due to a harmonic component of the measured total pressure distortion. A finite element (FE) model of the blade is then used to identify modal frequencies and predict vibratory stress and strain, providing information on the HCF potential and sensitivity to the imposed distortion. The method avoids the need for a full unsteady, coupled CFD-FE model of the entire compressor flowpath, and the calculations can be conducted rapidly. For the case evaluated, it is shown to produce stress predictions that compare to within 5% of the experimental values.© 2004 ASME

Design and Commissioning of an Engine-Like Test Environment

Test and evaluation are critical to any product development program. The validation of engine sensor products is particularly challenging since the test engine required for validation can range in value from thousands to millions of dollars, costing much more than the sensor product itself. As a result, significant sensor testing and validation is required by an engine owner prior to on-engine testing. To support our development activities and to facilitate test validation acceptance, we have created a test and evaluation platform for gas turbine sensors that will allow us to test developmental sensors in an engine-like environment without risking the possibility of engine damage. Driven by the core exhaust of a JT15-D engine, the Dynamic Rotor Research Rig (DR3) test and evaluation platform provides a test capability that is highly representative of the high temperatures, vibrations, gasses, fluids and overall gas turbine engine environment, while providing the means to easily add and replace sensors, add and test custom rotors, control temperature and rotor speeds, and to not risk engine health during test activities.Here we will discuss our sensor testing goals and how they fed into the operational goals and design considerations for the DR3. Early design concepts and the ultimate approach we took with the DR3 design will be explored, along with the candidate test rig component and subassembly fabrication processes that we evaluated and ultimately selected for use. We will review the manufacturing issues that we encountered during the construction phase of the DR3 and overview the commissioning of the DR3, problems that we discovered during start up and how we solved them.Included will be the results of initial turbine blade clearance and blade tip timing sensor testing performed on the DR3 and an evaluation of the DR3 performance, including temperature and speed control of the test rig and other characterization of the operating regime of the rig. Finally, we will present future plans to upgrade the DR3 rig to support future high temperature sensor and blade health monitoring development activities.Copyright

Application of an Improved Streamline Curvature Approach to a Modern, Two-Stage Transonic Fan: Comparison With Data and CFD

A streamline curvature method with improvements to key loss models is applied to a two-stage, low aspect ratio, transonic fan with design tip relative Mach number of approximately 1.65. Central to the improvements is the incorporation of a physics-based shock model. The attempt here is to capture the effects of key flow phenomena relative to the off-design performance of the fan. A quantitative analysis regarding solution sensitivities to model parameters that influence the key phenomena over a wide range of operating conditions is presented. Predictions are compared to performance determined from overall and interstage measurements, as well as from a three-dimensional, steady, Reynolds-averaged Navier-Stokes method applied across the first rotor. Overall and spanwise comparisons demonstrate that the improved model gives reasonable performance trending and generally accurate results. The method can be used to provide boundary conditions to higher-order solvers, or implemented within novel approaches using the streamline curvature method to explore complex engine-inlet integration issues, such as time-variant distortion.Copyright