Robert Frederick Bergholz

Experimental Investigation of Vane Clocking in a One and One-Half Stage High Pressure Turbine

Aerodynamic measurements were acquired on a modern single-stage, transonic, high-pressure turbine with the adjacent low-pressure turbine vane row (a typical civilian one and one-half stage turbine rig) to observe the effects of low-pressure turbine vane clocking on overall turbine performance. The turbine rig (loosely referred to in this paper as the stage) was operated at design corrected conditions using the Ohio State University Gas Turbine Laboratory Turbine Test Facility. The research program utilized uncooled hardware in which all three airfoils were heavily instrumented at multiple spans to develop a full clocking dataset. The low-pressure turbine vane row (LPTV) was clocked relative to the high-pressure turbine vane row (HPTV). Various methods were used to evaluate the influence of clocking on the aeroperformance (efficiency) and the aerodynamics (pressure loading) of the LPTV, including time-resolved and time-averaged measurements. A change in overall efficiency of approximately 2-3% due to clocking effects is demonstrated and could be observed using a variety of independent methods. Maximum efficiency is obtained when the time-average surface pressures are highest on the LPTV and the time-resolved surface pressure (both in the time domain and frequency domain) show the least amount of variation. The overall effect is obtained by integrating over the entire airfoil, as the three-dimensional (3D) effects on the LPTV surface are significant. This experimental data set validates several computational research efforts that suggested wake migration is the primary reason for the perceived effectiveness of vane clocking. The suggestion that wake migration is the dominate mechanism in generating the clocking effect is also consistent with anecdotal evidence that fully cooled engine rigs do not see a great deal of clocking effect. This is consistent since the additional disturbances induced by the cooling flows and/or the combustor make it extremely difficult to find an alignment for the LPTV given the strong 3D nature of modern high-pressure turbine flows.

Heat Transfer Measurements and Predictions for a Modern, High-Pressure, Transonic Turbine, Including Endwalls

This paper presents both measurements and predictions of the hot-gas-side heat transfer to a modern, 1 1 / 2 stage high-pressure, transonic turbine. Comparisons of the predicted and measured heat transfer are presented for each airfoil at three locations, as well as on the various endwalls and rotor tip. The measurements were performed using the Ohio State University Gas Turbine Laboratory Test Facility (TTF). The research program utilized an uncooled turbine stage at a range of operating conditions representative of the engine: in terms of corrected speed, flow function, stage pressure ratio, and gas-to-metal temperature ratio. All three airfoils were heavily instrumented for both pressure and heat transfer measurements at multiple locations. A 3D, compressible, Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD) solver with k-ω turbulence modeling was used for the CFD predictions. The entire 11 2 stage turbine was solved using a single computation, at two different Reynolds numbers. The CFD solutions were steady, with tangentially mass-averaged inlet/exit boundary condition profiles exchanged between adjacent airfoil-rows. Overall, the CFD heat transfer predictions compared very favorably with both the global operation of the turbine and with the local measurements of heat transfer. A discussion of the features of the turbine heat transfer distributions, and their association with the corresponding flow-physics, has been included.

Heat-Flux Measurements and Predictions for the Blade Tip Region of a High-Pressure Turbine

High-pressure turbine blade tips operate in a highly complex flow environment that makes designing new blades for increased life difficult. Computational fluid dynamics simulations of the tip flow field may be able to guide new designs to improve the blade life, but the analysis techniques need to be verified against detailed measurements before they can be applied. The current paper presents measurements of heat flux and pressure in the blade tip region of a modern one-and-one-half stage high-pressure turbine operating at design corrected conditions in a rotating rig. Both flat tip and recessed, or squealer, tip blades were used in the experiments. The measurements indicate that the recessed tip, used in the majority of modern turbines to minimize blade damage from rubs, increases the blade heat load overall, and creates several hot spots on the floor of the recess for an uncooled airfoil. The tip data also showed there were significant unsteady variations in the heat load at the vane passing frequency. Steady state CFD calculations were completed for both flat and squealer tip configurations to examine if the analysis could capture the details that were measured. The CFD, while not capable of estimating the unsteady heat load component and generally over predicting the overall heat flux by 10–25%, did capture the measured heat flux trends in the recessed tip. These results show that steady-state CFD analysis can be useful in predicting the complex flow field and heat load distribution in turbine blade tips to help guide future blade designs.Copyright

Aerodynamic and Heat-flux Measurements with Predictions on a Modern One and One-Half State High Pressure transonic Turbine

Aerodynamic and heat-transfer measurements were acquired using a modern stage and 1/2 high-pressure turbine operating at design corrected conditions and pressure ratio. These measurements were performed using the Ohio State University Gas Turbine Laboratory Turbine Test Facility. The research program utilized an uncooled turbine stage for which all three airfoils are heavily instrumented at multiple spans to develop a full database at different Reynolds numbers for code validation and flow-physics modeling. The pressure data, once normalized by the inlet conditions, was insensitive to the Reynolds number. The heat-flux data for the high-pressure stage suggests turbulent flow over most of the operating conditions and is Reynolds number sensitive. However, the heat-flux data do not scale according to flat plat theory for most of the airfoil surfaces. Several different predictions have been done using a variety of design and research codes. In this work, comparisons are made between industrial codes and an older code called UNSFLO-2D initially published in the late 1980s. The comparisons show that the UNSFLO-2D results at midspan are comparable to the modern codes for the time-resolved and time-averaged pressure data, which is remarkable given the vast differences in the processing required. UNSFLO-2D models the entropy generated around the airfoil surfaces using the full Navier-Stokes equations, but propagates the entropy invisicidly downstream to the next blade row, dramatically reducing the computational power required. The accuracy of UNSFLO-2D suggests that this type of approach may be far more useful in creating time-accurate design tools, than trying to utilize full time-accurate Navier-Stokes codes which are often currently used as research codes in the engine community, but have yet to be fully integrated into the design system due to their complexity and significant processor requirements.

Rotor / Stator Heat Transfer Measurements and CFD Predictions for Short-Duration Turbine Rig Tests

This paper describes aerodynamic and heat transfer measurements, improved data analysis techniques, and CFD predictions for an extensive series of high-pressure turbine tests carried out in a large-scale, short-duration test facility. The focus of this work is on the test methods, aerodynamic and heat transfer instrumentation, and the heat flux data obtained on the HPT blade and vane pressure and suction surfaces, and on the blade platform. Heat transfer data are presented along three spanwise stream surfaces, and at five platform locations, for two Reynolds numbers, and for three different initial blade temperature levels. An analytical method for improving the interpretation of transient heat flux gage measurements and estimating the uncertainties in heat transfer coefficients is applied based on test facility characteristics, established 2D boundary layer heat transfer theory, and a detailed calculation of the transient temperature response of the heat flux gages as installed in the airfoils. The result is a systematic procedure for improved data reduction and correction of bias errors resulting from in-situ heat flux gage behavior in complex rotating rig tests. Three-dimensional aerodynamic calculations are compared with time-averaged unsteady pressure measurements on the airfoils. Boundary-layer heat transfer predictions are shown to be in fair agreement with the pitchline data but sensitive to assumptions regarding free-stream turbulence, uncertainties in inlet conditions, and variations in the applied surface Mach number distribution. Finally, full 3D heat transfer computations for the blade are discussed, with an emphasis on capturing overall 3D flow effects with sufficient accuracy for practical turbine airfoil cooling design. The results suggest that advanced analysis techniques, including 3D CFD, can be used effectively to compute not only mean values of the surface heat transfer coefficients, but also to quantify the uncertainties in the predictions and to identify the dominant sources of those uncertainties. This is an important step in creating a robust turbine cooling design process that accounts for environmental, manufacturing, and modeling variations.Future papers will present experimental and computational results for the blade tip and shroud, the vane endwalls, and TBC-coated airfoils.Copyright

Preliminary Design Optimization of Impingement Cooled Turbine Airfoils

Generally, the objective of the turbine cooling design process is to minimize the amount of cooling flow required to achieve component life goals, including creep rupture, low-cycle (thermal-mechanical) fatigue, and material oxidation. High-pressure turbine airfoils in particular are subject to large variations in external gas temperatures and heat transfer coefficients, and high aero-mechanical loads. The efficient distribution of internal and film cooling flows, the management of internal coolant temperatures, and the optimal arrangement of cooling features to control thermal gradients are major factors in controlling thermal stress. The objective of this paper is to describe a process for conceptual and preliminary turbine airfoil cooling design based on the assembly of primitive cooling elements to define an overall “realizable” 3D airfoil cooling structure. This approach allows the evaluation of multiple cooling configurations, and potentially an improved prediction of cooling flows, early in the design process. This conceptual design method can be used to quickly generate models for fabrication and testing in a rapid prototyping laboratory. The design process is outlined for the simplified case of impingement cooling in a prototypical airfoil shape to illustrate some of the key design parameters and procedures. A future paper will address improvements in automation and more complex geometries. The paper also presents predictions based on a parametric impingement cooling model for the local surface heat transfer coefficient distribution. The model is easily incorporated into the conceptual design method. It is particularly useful for determining axial and radial thermal gradients the airfoil wall induced by sparse impingement arrays, for which the use of average heat transfer coefficients give unsatisfactory results. The model was calibrated using data from past published literature and more recently obtained experimental data from the GE Global Research Center.Copyright

Experimental Investigation of Vane Clocking in a One and 1/2 Stage High Pressure Turbine

Aerodynamic measurements were acquired on a modern single-stage, transonic, high-pressure turbine with the adjacent low-pressure turbine vane row (a typical civilian one and one-half stage turbine rig) to observe the effects of low-pressure turbine vane clocking on overall turbine performance. The turbine rig (loosely referred to in this paper as the stage) was operated at design corrected conditions using the Ohio State University Gas Turbine Laboratory Turbine Test Facility (TTF). The research program utilized uncooled hardware in which all three airfoils were heavily instrumented at multiple spans to develop a full clocking dataset. The low-pressure turbine vane row (LPTV) was clocked relative to the high-pressure turbine vane row (HPTV). Various methods were used to evaluate the influence of clocking on the aeroperformance (efficiency) and the aerodynamics (pressure loading) of the LPTV, including time-resolved and time-averaged measurements. A change in overall efficiency of approximately 2–3% due to clocking effects is demonstrated and could be observed using a variety of independent methods. Maximum efficiency is obtained when the time-average surface pressures are highest on the LPTV and the time-resolved surface pressure (both in the time domain and frequency domain) show the least amount of variation. The overall effect is obtained by integrating over the entire airfoil, as the three-dimensional effects on the LPTV surface are significant. This experimental data set validates several computational research efforts that suggested wake migration is the primary reason for the perceived effectiveness of vane clocking. The suggestion that wake migration is the dominate mechanism in generating the clocking effect is also consistent with anecdotal evidence that fully cooled engine rigs do not see a great deal of clocking effect. This is consistent since the additional disturbances induced by the cooling flows and/or the combustor make it extremely difficult to find an alignment for the LPTV given the strong 3D nature of modern high-pressure turbine flows.Copyright

Aerodynamic and Heat-Flux Measurements With Predictions on a Modern One and 1/2 Stage High Pressure Transonic Turbine

Aerodynamic and heat-transfer measurements were acquired using a modern stage and 1/2 high-pressure turbine operating at design corrected conditions and pressure ratio. These measurements were performed using the Ohio State University Gas Turbine Laboratory Turbine Test Facility (TTF). The research program utilized an uncooled turbine stage for which all three airfoils are heavily instrumented at multiple spans to develop a full database at different Reynolds numbers for code validation and flow-physics modeling. The pressure data, once normalized by the inlet conditions, was insensitive to the Reynolds number. The heat-flux data for the high-pressure stage suggests turbulent flow over most of the operating conditions and is Reynolds number sensitive. However, the heat-flux data does not scale according to flat plat theory for most of the airfoil surfaces. Several different predictions have been done using a variety of design and research codes. In this work, comparisons are made between industrial codes and an older code called UNSFLO-2D initially published in the late 1980’s. The comparisons show that the UNSFLO-2D results at midspan are comparable to the modern codes for the time-resolved and time-averaged pressure data, which is remarkable given the vast differences in the processing required. UNSFLO-2D models the entropy generated around the airfoil surfaces using the full Navier-Stokes equations, but propagates the entropy invisicidly downstream to the next blade row, dramatically reducing the computational power required. The accuracy of UNSFLO-2D suggests that this type of approach may be far more useful in creating time-accurate design tools, than trying to utilize full time-accurate Navier-stokes codes which are often currently used as research codes in the engine community, but have yet to be fully integrated into the design system due to their complexity and significant processor requirements.Copyright