Randall M. Mathison

Aerodynamic and Heat Flux Measurements in a Single-Stage Fully Cooled Turbine—Part II: Experimental Results

This paper describes the experimental approach utilized to perform experiments using a fully cooled rotating turbine stage to obtain film effectiveness measurements. Significant changes to the previous experimental apparatus were implemented to meet the experimental objectives. The modifications include the development of a synchronized blowdown facility to provide cooling gas to the turbine stage, installation of a heat exchanger capable of generating a uniform or patterned inlet temperature profile, novel utilization of temperature and pressure instrumentation, and development of robust double-sided heat flux gauges. With these modifications, time-averaged and time-accurate measurements of temperature, pressure, surface heat flux, and film effectiveness can be made over a wide range of operational parameters, duplicating the nondimensional parameters necessary to simulate engine conditions. Data from low Reynolds number experiments are presented to demonstrate that all appropriate scaling parameters can be satisfied and that the new components have operated correctly. Along with airfoil surface heat transfer and pressure data, temperature and pressure data from inside the coolant plenums of the vane and rotating blade airfoils are presented. Pressure measurements obtained inside the vane and blade plenum chambers illustrate passing of the wakes and shocks as a result of vane/blade interaction. Part II of this paper (Haldeman, C. W., Mathison, R. M., Dunn, M. G., Southworth, S. A., Harral, J. W., and Heltland, G., 2008, ASME J. Turbomach., 130(2), p. 021016) presents data from the low Reynolds number cooling experiments and compares these measurements to CFD predictions generated using the Numeca FINE/Turbo package at multiple spans on the vanes and blades.

Aerodynamics and Heat Transfer for a Cooled One and One-Half Stage High-Pressure Turbine—Part I: Vane Inlet Temperature Profile Generation and Migration

As controlled laboratory experiments using full-stage turbines are expanded to replicate more of the complicated flow features associated with real engines, it is important to understand the influence of the vane inlet temperature profile on the high-pressure vane and blade heat transfer as well as its interaction with film cooling. The temperature distribution of the incoming fluid governs not only the input conditions to the boundary layer but also the overall fluid migration. Both of these mechanisms have a strong influence on surface heat flux and therefore component life predictions. To better understand the role of the inlet temperature profile, an electrically heated combustor emulator capable of generating uniform, radial, or hot streak temperature profiles at the high-pressure turbine vane inlet has been designed, constructed, and operated over a wide range of conditions. The device is shown to introduce a negligible pressure distortion while generating the inlet temperature conditions for a stage-and-a-half turbine operating at design-corrected conditions. For the measurements described here, the vane is fully cooled and the rotor purge flow is active, but the blades are uncooled. Detailed temperature measurements are obtained at rake locations upstream and downstream of the turbine stage as well as at the leading edge and platform of the blade in order to characterize the inlet temperature profile and its migration. The use of miniature butt-welded thermocouples at the leading edge and on the platform (protruding into the flow) on a rotating blade is a novel method of mapping a temperature profile. These measurements show that the reduction in fluid temperature due to cooling is similar in magnitude for both uniform and radial vane inlet temperature profiles.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part I: Influence of Cooling Variation

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design-corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film-cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the uncooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton number distribution on the blade.

History of Short-Duration Measurement Programs Related to Gas Turbine Heat Transfer, Aerodynamics, and Aeroperformance at Calspan and OSU

Short-duration facilities have been used for the past thirty-five years to obtain measurements of heat transfer, aerodynamic loading, vibratory response, film-cooling influence, purge flow migration, and aeroperformance for full-stage high-pressure turbines operating at design corrected conditions of flow function, corrected speed, and stage pressure ratio. This paper traces the development of experimental techniques now in use at The Ohio State University (OSU) Gas Turbine Laboratory (GTL) from initial work in this area at the Cornell Aeronautical Laboratory (CAL, later to become Calspan) in 1975 through to the present. It is intended to summarize the wide range of research that can be performed with a short-duration facility and highlight the types of measurements that are possible.Beginning with heat-flux measurements for the vane and blade of a Garrett TFE 731-2 HP turbine stage with vane pressure-surface slot cooling, the challenge of each experimental program has been to provide data to aid turbine designers in understanding the relevant flow physics and help drive the advancement of predictive techniques. Through many different programs, this has involved collaborators at a variety of companies and experiments performed with turbine stages from Garrett, Allison, Teledyne, Pratt and Whitney, General Electric Aviation, Rocketdyne, Westinghouse, and Honeywell. The Vane/Blade Interaction measurement and CFD program, which ran from the early eighties until 2000, provided a particularly good example of what can be achieved when experimentalists and computational specialists collaborate closely. Before conclusion of this program in 2000, the heat-flux and pressure measurements made for this transonic turbine operated with and without vane trailing edge cooling flow were analyzed and compared to predictive codes in conjunction with engineers at Allison, United Technologies Research Center, Pratt and Whitney, and GE Aviation in jointly published papers.When the group moved to OSU in 1995 along with the facility used at Calspan, refined techniques were needed to meet new research challenges such as investigating blade damping and forced response, measuring aeroperformance for different configurations, and preparing for advanced cooling experiments that introduced complicating features of an actual engine to further challenge computational predictions. This required conversion of the test-gas heating method from a shock-tunnel approach to a blowdown approach using a combustor emulator to also create inlet temperature profiles, the development of instrumentation techniques to work with a thin-walled airfoil with backside cooling, and the adoption of experimental techniques that could be used to successfully operate fully cooled turbine stages (vane row cooled, blade row cooled, and proper cavity purge flow provided). Further, it was necessary to develop techniques for measuring the aeroperformance of these fully cooled machines.Copyright

Design, Construction, and Operation of a Combustor Emulator for Short-Duration High-Pressure Turbine Experiments

Short-duration facilities have proven to be very adept at creating uniform inlet conditions for full-scale rotating turbine rigs to create experimental data sets used for design code verification. However, it has been well known that the inlet to the high-pressure turbine stage in an engine is anything but uniform, and as codes have improved, it has become clear that reproducing these inlet conditions is critical to proper modeling. To aid in this work, a combustor emulator has been designed and constructed that allows the user to create a variety of temperature profiles (radial, circumferential, pure hot streak) that are used to modify the inlet conditions to a high-pressure turbine stage operating at corrected engine conditions. This emulator differs from past designs in that it allows for changes in the number of circumferential “hot-streaks”, the circumferential location of these hotstreaks, as well as their radial locations. These are usually chosen to mimic the number of combustor nozzles in a real engine (between 15 and 20) so that one can model the clocking interactions between the hot spots and the HPT vane. The data presented in this paper shows the different types of temperature profiles that can be obtained with this device and compares the predicted design characteristics of the temperature profiles to the experimentally observed properties.

Time-Averaged and Time-Accurate Aerodynamic Effects of Forward Rotor Cavity Purge Flow for a High-Pressure Turbine—Part I: Analytical and Experimental Comparisons

The effect of rotor purge flow on the unsteady aerodynamics of a high-pressure turbine stage operating at design corrected conditions has been investigated, both experimentally and computationally. The experimental configuration consisted of a single-stage high-pressure turbine with a modern film-cooling configuration on the vane airfoil and the inner and outer end wall surfaces. Purge flow was introduced into the cavity located between the high-pressure vane and the high-pressure disk. The high-pressure blades and the downstream low-pressure turbine nozzle row were not cooled. All of the hardware featured an aerodynamic design typical of a commercial high-pressure ratio turbine and the flow path geometry was representative of the actual engine hardware. In addition to instrumentation in the main flow path, the stationary and rotating seals of the purge flow cavity were instrumented with high frequency response flush-mounted pressure transducers and miniature thermocouples in order to measure the flow field parameters above and below the angel wing.Predictions of the time-dependent flow field in the turbine flow path were obtained using FINE/Turbo, a three-dimensional Reynolds-averaged Navier–Stokes computational fluid dynamics CFD code that had the capability to perform both a steady and unsteady analysis. The steady and unsteady flow fields throughout the turbine were predicted using a three blade-row computational model that incorporated the purge flow cavity between the high-pressure vane and disk. The predictions were performed in an effort to mimic the design process with no adjustment of boundary conditions to better match the experimental data. The time-accurate predictions were generated using the harmonic method. Part I of this paper concentrates on the comparison of the time-averaged and time-accurate predictions with measurements in and around the purge flow cavity. The degree of agreement between the measured and predicted parameters is described in detail, providing confidence in the predictions for the flow field analysis that will be provided in Part II.

Comparison of Harmonic and Time Marching Unsteady Computational Fluid Dynamics Solutions With Measurements for a Single-Stage High-Pressure Turbine

The unsteady aerodynamics of a single-stage high-pressure turbine has been the subject of a study involving detailed measurements and computations. Data and predictions for this experiment have been presented previously, but the current study compares predictions obtained using the nonlinear harmonic simulation method to results obtained using a time-marching simulation with phase-lag boundary conditions. The experimental configuration consisted of a single-stage high-pressure turbine (HPT) and the adjacent, downstream, low-pressure turbine nozzle row (LPV) with an aerodynamic design that is typical to that of a commercial high-pressure ratio HPT and LPV. The flow path geometry was equivalent to engine hardware and operated at the proper design-corrected conditions to match cruise conditions. The high-pressure vane and blade were uncooled for these comparisons. All three blade rows are instrumented with flush-mounted, high-frequency response pressure transducers on the airfoil surfaces and the inner and outer flow path surfaces, which include the rotating blade platform and the stationary shroud above the rotating blade. Predictions of the time-dependent flow field for the turbine flow path were obtained using a three-dimensional, Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) code. Using a two blade row computational model of the turbine flow path, the unsteady surface pressure for the high-pressure vane and rotor was calculated using both unsteady methods. The two sets of predictions are then compared to the measurements looking at both time-averaged and time-accurate results, which show good correlation between the two methods and the measurements. This paper concentrates on the similarities and differences between the two unsteady methods, and how the predictions compare with the measurements since the faster harmonic solution could allow turbomachinery designers to incorporate unsteady calculations in the design process without sacrificing accuracy when compared to the phase-lag method.

Aerodynamics and Heat Transfer for a Cooled One and One-Half Stage High-Pressure Turbine: Part I—Vane Inlet Temperature Profile Generation and Migration

As controlled laboratory experiments using full-stage turbines are expanded to replicate more of the complicated flow features associated with real engines, it is important to understand the influence of the vane inlet temperature profile on the high-pressure vane and blade heat transfer as well as its interaction with film cooling. The temperature distribution of the incoming fluid governs not only the input conditions to the boundary layer but overall fluid migration. Both of these mechanisms have a strong influence on surface heat flux and therefore component life predictions. To better understand the role of the inlet temperature profile, an electrically heated combustor emulator capable of generating uniform, radial, or hot-streak temperature profiles at the high-pressure turbine vane inlet has been designed, constructed, and operated over a wide range of conditions. The device is shown to introduce a negligible pressure distortion while generating the inlet temperature conditions for a stage-and-a-half turbine operating at design-corrected conditions. For the measurements described here, the vane is fully cooled and the rotor purge flow is active but the blades are un-cooled. Detailed temperature measurements are obtained at rake locations upstream and downstream of the turbine stage as well as at the leading edge and platform of the blade in order to characterize the inlet temperature profile and its migration. The use of miniature butt-welded thermocouples at the leading edge and on the platform (protruding into the flow) on a rotating blade is a novel method of mapping temperature profile. These measurements show that the reduction in fluid temperature due to cooling is similar in magnitude for both a uniform and radial vane inlet temperature profile.Copyright

Aerodynamic and Heat Flux Measurements in a Single Stage Fully Cooled Turbine: Part I — Experimental Approach

This paper presents measurements and the companion CFD predictions for a fully cooled, high-work single stage HP turbine operating in a short-duration blowdown rig. Part I of this paper presented the experimental approach, and Part II focuses on the results of the measurements and demonstrates how these results compare to predictions made using the Numeca FINE/Turbo CFD package. The measurements are presented in both time-averaged and time-accurate formats. The results include the heat transfer at multiple spans on the vane, blade, and rotor shroud as well as flow path measurements of total temperature and total pressure. Surface pressure measurements are available on the vane at midspan, and on the blade at 50% and 90% spans as well as the rotor shroud. In addition, temperature and pressure measurements obtained inside the coolant cavities of both the vanes and blades are presented. Time-averaged values for the surface pressure on the vane and blade are compared to steady CFD predictions. Additional comparisons will be made between the heat transfer on cooled blades and uncooled blades with identical surface geometry. This, along with measurements of adiabatic wall temperature, will provide a basis for analyzing the effectiveness of the film-cooling scheme at a number of locations.Copyright

Measurement of Air Film Damping Effectiveness

Air film damping systems have attracted considerable interest within the gas turbine industry because of their effectiveness at controlling modes of vibration without environmental limitations. Though still in the early stages of development, air film dampers have promise for improving the high cycle fatigue characteristics of solid gas turbine airfoils. This study used experimental methods to compare the vibrational response of a solid flat plate with the response of an identically sized plate that incorporated an air film damper. It also investigated the influence of elevated pressures on air film damping effectiveness, the impact of the damper on the various vibration modes, and the relative strain levels of the air film cover plate to the solid backing. The results show that the air film damper is very effective in controlling the two-stripe mode for which it was designed. Increasing the surrounding air pressure makes the damper more effective and shifts the resonant frequencies lower.