Influence of vane/blade spacing and cold-gas injection on vane and blade heat-flux distributions for the Teledyne 702 HP turbine stage

In order to have higher power to weight ratio and higher efficiency gas turbine engines, turbine inlet temperatures continue to rise. State-of-the-art turbine inlet temperatures now exceed the turbine rotor material capability. Accordingly, one of the best methods to protect turbine airfoil surfaces is to use film cooling on the airfoil external surfaces. In general, sizable amounts of expensive cooling flow delivered from the core compressor are used to cool the high temperature surfaces. That sizable cooling flow, on the order of 20% of the compressor core flow, adversely impacts the overall engine performance and hence the engine power density. With better understanding of the cooling flow and accurate prediction of the heat transfer distribution on airfoil surfaces, heat transfer designers can have a more efficient design to reduce the cooling flow needed for high temperature components and improve turbine efficiency. This in turn lowers the overall specific fuel consumption (SFC) for the engine. Accurate prediction of rotor metal temperature is also critical for calculations of cyclic thermal stress, oxidation, and component life. The utilization of three-dimensional computational fluid dynamics (3D CFD) codes for turbomachinery aerodynamic design and analysis is now a routine practice in the gas turbine industry. The accurate heat-transfer and metal-temperature prediction capability of any CFD code, however, remains challenging. This difficulty is primarily due to the complex flow environment of the high-pressure turbine, which features high speed rotating flow, coupling of internal and external unsteady flows, and film-cooled, heat transfer enhancement schemes. In this study, conjugate heat transfer (CHT) simulations are performed on a high-pressure cooled turbine stage, and the heat flux results at mid span are compared to experimental data obtained at The Ohio State University Gas Turbine Laboratory (OSUGTL). Due to the large difference in time scales between fluid and solid, the fluid domain is simulated as steady state while the solid domain is simulated as transient in CHT simulation. This paper compares the unsteady and transient results of the heat flux on a high-pressure cooled turbine rotor with measurements obtained at OSUGTL.

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Time-Resolved Heat Transfer and Surface Pressure Measurements for a Fully Cooled Transonic Turbine Stage

Journal of Turbomachinery ◽

10.1115/1.4029950 ◽

2015 ◽

Vol 137 (9) ◽

Cited By ~ 3

Author(s):

Jeremy B. Nickol ◽

Randall M. Mathison ◽

Malak F. Malak ◽

Rajiv Rana ◽

Jong S. Liu

Keyword(s):

Heat Transfer ◽

Heat Flux ◽

Flow Field ◽

Surface Pressure ◽

High Frequency ◽

Gas Turbines ◽

Pressure Measurements ◽

Turbine Stage ◽

Cooling Flow ◽

Blade Cooling

The flow field in axial gas turbines is driven by strong unsteady interactions between stationary and moving components. While time-averaged measurements can highlight many important flow features, developing a deeper understanding of the complicated flows present in high-speed turbomachinery requires time-accurate measurements that capture this unsteady behavior. Toward this end, time-accurate measurements are presented for a fully cooled transonic high-pressure turbine stage operating at design-corrected conditions. The turbine is run in a short-duration blowdown facility with uniform, radial, and hot streak vane-inlet temperature profiles as well as various amounts of cooling flow. High-frequency response surface pressure and heat-flux instrumentation installed in the rotating blade row, stator vane row, and stationary outer shroud provide detailed measurements of the flow behavior for this stage. Previous papers have reported the time-averaged results from this experiment, but this paper focuses on the strong unsteady phenomena that are observed. Heat-flux measurements from double-sided heat-flux gauges (HFGs) cover three spanwise locations on the blade pressure and suction surfaces. In addition, there are two instrumented blades with the cooling holes blocked to isolate the effect of just blade cooling. The stage can be run with the vane and blade cooling flow either on or off. High-frequency pressure measurements provide a picture of the unsteady aerodynamics on the vane and blade airfoil surfaces, as well as inside the serpentine coolant supply passages of the blade. A time-accurate computational fluid dynamics (CFD) simulation is also run to predict the blade surface pressure and heat-flux, and comparisons between prediction and measurement are given. It is found that unsteady variations in heat-flux and pressure are stronger at low to midspan and weaker at high span, likely due to the impact of secondary flows such as the tip leakage flow. Away from the tip, it is seen that the unsteady fluctuations in pressure and heat-flux are mostly in phase with each other on the suction side, but there is some deviation on the pressure side. The flow field is ultimately shown to be highly three-dimensional, as the movement of high heat transfer regions can be traced in both the chord and spanwise directions. These measurements provide a unique picture of the unsteady flow physics of a rotating turbine, and efforts to better understand and model these time-varying flows have the potential to change the way we think about even the time-averaged flow characteristics.

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Experimental Study of Artificial Ground Freezing by Natural Cold Gas Injection

Applied Sciences ◽

10.3390/app10176055 ◽

2020 ◽

Vol 10 (17) ◽

pp. 6055

Author(s):

Zhao Liu ◽

Youhong Sun ◽

Bingge Wang ◽

Qiang Li

Keyword(s):

Phase Transition ◽

Gas Flow ◽

Gas Injection ◽

Freezing Rate ◽

New Method ◽

Cooling Stage ◽

Ground Freezing ◽

Artificial Ground Freezing ◽

Time Required ◽

Cold Gas

The application of conventional artificial ground freezing (AGF) has two disadvantages: low freezing rate and small frozen range. In this study, a new method with natural cold gas injection was proposed, whereby the shallow soils and water can be frozen rapidly due to the effect of the heat convection. Cold gas from −15 °C to −10 °C, in the winter of northeast China, was injected into the laboratory-scale sand pipe; evolution of the induced frozen front and water migration were studied, and then, the feasibility of the new method was analyzed. According to the evolution of the induced frozen front, the freezing process was divided into an initial cooling stage, phase transition stage, and subcooled stage. The results showed that the increase of initial water content at the beginning of the experiments had little effect on the time required for completing the initial cooling stage, while the time required for the phase transition would increase in nearly the same proportion. In addition, the increase of the cold gas flow rate could not only strengthen the cooling rate of the initial cooling stage but also shorten the phase transition time; thereby, the freezing rate was increased. The freezing rate could reach 0.18–0.61 cm/min in the direction of cold gas flow, and compared to the conventional AGF (months are required for approximately 1 m), the freezing efficiency was greatly improved.

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Tangential Gas Injection Into a Supersonic Stream

2010 14th International Heat Transfer Conference, Volume 2 ◽

10.1115/ihtc14-22922 ◽

2010 ◽

Author(s):

Eduard B. Vasilevskiy ◽

Ivan V. Ezhov ◽

Andrey V. Novikov

Keyword(s):

Heat Flux ◽

Free Stream ◽

Numerical Study ◽

Gas Injection ◽

Angle Of Attack ◽

Supersonic Stream ◽

Stagnation Temperature ◽

Mass Rate ◽

Free Stream Mach Number ◽

Injection Effect

An experimental and numerical study of a tangential gas injection effect on a flow pattern and heat flux was carried out. The cooling gas (air) was injected in the flow (air) through the tangential axis-symmetric slot on the spherically blunted cylinder streamlined longitudinally. Experiments were conducted in TsAGI shock wind tunnel at free-stream Mach number M∞ = 6, Reynolds number Re∞, Rw = 0.76×106 (calculated for free-stream parameters and bluntness radius Rw = 37.5 mm), cylinder angle of attack α = 0…30°, slot width h* = hk/Rw = 0–0.021, free-stream stagnation temperature T0 = 710 K, pressure behind the normal shock ps = 0.5 bar. The mass rate of the injected gas G* = gj/πρ∞u∞rw2 = 0…0.12. It is shown, that maximum of the heat flux toward the sphere surface could be sufficiently decreased. For example, for coolant mass rate G* = 0.03 and angle of attack α = 0 the heat flux maximum is reduced by factor of two.

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Fully-Cooled Single Stage HP Transonic Turbine: Part II—Influence of Cooling Mass Flow Changes and Inlet Temperature Profiles on Blade and Shroud Heat-Transfer

Volume 4: Heat Transfer, Parts A and B ◽

10.1115/gt2010-23445 ◽

2010 ◽

Cited By ~ 1

Author(s):

C. W. Haldeman ◽

M. G. Dunn ◽

R. M. Mathison

Keyword(s):

Heat Transfer ◽

Heat Flux ◽

Temperature Profile ◽

Mass Flow ◽

Stanton Number ◽

Inlet Temperature ◽

Temperature Profiles ◽

Turbine Stage ◽

Flow Rates ◽

Cooling Effects

A fully cooled transonic HP turbine stage is utilized to investigate the combined effects of turbine stage cooling variation and vane inlet temperature profile on heat transfer to the blades with the stage operating at the proper design corrected conditions. For this series of experiments, both the vane row and the blade row were fully cooled. The matrix of experimental conditions included varying the cooling flow rates and the vane inlet temperature profiles to observe the overall effect on airfoil heat-transfer. The data presented in Part I focused on the aerodynamics of the fully cooled turbine for a subset of the cases investigating two vane inlet temperature profiles (uniform and radial), and three different cooling levels (none, nominal and high) for the high Reynolds number condition. This part of the paper focuses on the time-average heat-flux measurements on the blade and shroud region for the same cooling mass flow rates and vane inlet temperature profiles. The cooling effects are shown to be small and are centered primarily on the suction side of the airfoil. This relatively small influence is due to the ratio of the cooling gas to metal temperature being closer to 1 than the design value would dictate. The vane inlet temperature profile effects are more dominant, and using a Net Stanton Number Reduction Factor to compare the cases, an effect on the order of about 0.25 is demonstrated. This effect is due primarily to the change in the reference temperature used for the Stanton number calculation. The differences due to profile effects are small, but observable towards the trailing edge of both the blade and rotor shroud. This data set forms an excellent baseline for heat-flux calculations, as the variation in the main input conditions are well documented and do not produce large changes in the heat-flux. It provides insight into the flow physics of an actual engine and guidelines about proper normalization of variables for a cooled turbine stage, supporting further development of computational heat-flux modeling techniques.

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Time-Resolved Heat Transfer and Surface Pressure Measurements for a Fully-Cooled Transonic Turbine Stage

Volume 5B: Heat Transfer ◽

10.1115/gt2014-26407 ◽

2014 ◽

Author(s):

Jeremy B. Nickol ◽

Randall M. Mathison ◽

Malak F. Malak ◽

Rajiv Rana ◽

Jong S. Liu

Keyword(s):

Heat Transfer ◽

Heat Flux ◽

Flow Field ◽

Surface Pressure ◽

High Frequency ◽

Gas Turbines ◽

Pressure Measurements ◽

Turbine Stage ◽

Cooling Flow ◽

Blade Cooling

The flow field in axial gas turbines is driven by strong unsteady interactions between stationary and moving components. While time-averaged measurements can highlight many important flow features, developing a deeper understanding of the complicated flows present in high-speed turbomachinery requires time-accurate measurements that capture this unsteady behavior. Towards this end, time-accurate measurements are presented for a fully cooled transonic high-pressure turbine stage operating at design-corrected conditions. The turbine is run in a short-duration blowdown facility with uniform, radial, and hot streak vane-inlet temperature profiles as well as various amounts of cooling flow. High frequency response surface-pressure and heat-flux instrumentation installed in the rotating blade row, stator vane row, and stationary outer shroud provide detailed measurements of the flow behavior for this stage. Previous papers by Haldeman et al. [1, 2] have reported the time-averaged results from this experiment, but this paper focuses on the strong unsteady phenomena that are observed. Heat-flux measurements from double-sided heat-flux gauges cover three span-wise locations on the blade pressure and suction surfaces. In addition, there are two instrumented blades with the cooling holes blocked to isolate the effect of just blade cooling. The stage can be run with the vane and blade cooling flow either on or off. High-frequency pressure measurements provide a picture of the unsteady aerodynamics on the vane and blade airfoil surfaces, as well as inside the serpentine coolant supply passages of the blade. A time-accurate CFD simulation is also run to predict the blade surface pressure and heat-flux, and comparisons between prediction and measurement are given. It is found that unsteady variations in heat-flux and pressure are stronger at low to mid-span and weaker at high span, likely due to the impact of secondary flows such as the tip leakage flow. Away from the tip, it is seen that the unsteady fluctuations in pressure and heat-flux are mostly in-phase with each other on the suction side, but there is some deviation on the pressure side. The flow field is ultimately shown to be highly three-dimensional, as the movement of high heat transfer regions can be traced in both the chord and span-wise directions. These measurements provide a unique picture of the unsteady flow physics of a rotating turbine, and efforts to better understand and model these time-varying flows have the potential to change the way we think about even the time-averaged flow characteristics.

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Volt-Ampere and Thermal Features of a Direct-Current Dual-Jet Plasma Generator With a Cold Gas Injection

IEEE Transactions on Plasma Science ◽

10.1109/tps.2010.2061238 ◽

2010 ◽

Vol 38 (10) ◽

pp. 2906-2913 ◽

Cited By ~ 6

Author(s):

Zhe Wang ◽

Gui-Qing Wu ◽

Nan Ge ◽

He-Ping Li ◽

Cheng-Yu Bao

Keyword(s):

Direct Current ◽

Gas Injection ◽

Plasma Generator ◽

Thermal Features ◽

Jet Plasma ◽

Cold Gas

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Time-Averaged Heat Flux for a Recessed Tip, Lip, and Platform of a Transonic Turbine Blade

Volume 3: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/2000-gt-0197 ◽

2000 ◽

Cited By ~ 11

Author(s):

M. G. Dunn ◽

C. W. Haldeman

Keyword(s):

Heat Flux ◽

Pressure Ratio ◽

Design Flow ◽

Turbine Stage ◽

Suction Surface ◽

Surface Measurements ◽

Blade Tip ◽

Flux Level ◽

Transonic Turbine ◽

Exit Mach Number

The results of an experimental research program determining the blade platform heat-flux level and the influence of blade tip recess on the tip region heat transfer for a full-scale rotating turbine stage at transonic vane exit conditions are described. The turbine used for these measurements was the Allison VBI stage operating in the closed vane position (vane exit Mach number _ 1.1). The stage was operated at the design flow function, total to static pressure ratio, and corrected speed. Measurements were obtained at several locations on the platform and in the blade tip region. The tip region consists of the bottom of the recess, the lip region (on both the pressure and suction surface sides of the recess), and the 90% span location on the blade suction surface. Measurements were obtained for three vane/blade spacings; 20%, 40%, and 60% of vane axial chord and for a single value of the tip gap (the distance between the top of the lip and the stationary shroud) equal to 0.0012-m (0.046-in) or 2.27% of blade height.

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Time-Averaged Heat Flux for a Recessed Tip, Lip, and Platform of a Transonic Turbine Blade

Journal of Turbomachinery ◽

10.1115/1.1311285 ◽

2000 ◽

Vol 122 (4) ◽

pp. 692-698 ◽

Cited By ~ 37

Author(s):

M. G. Dunn ◽

C. W. Haldeman

Keyword(s):

Heat Flux ◽

Pressure Ratio ◽

Design Flow ◽

Turbine Stage ◽

Suction Surface ◽

Surface Measurements ◽

Blade Tip ◽

Flux Level ◽

Transonic Turbine ◽

Exit Mach Number

The results of an experimental research program determining the blade platform heat-flux level and the influence of blade tip recess on the tip region heat transfer for a full-scale rotating turbine stage at transonic vane exit conditions are described. The turbine used for these measurements was the Allison VBI stage operating in the closed vane position (vane exit Mach number≈1.1). The stage was operated at the design flow function, total to static pressure ratio, and corrected speed. Measurements were obtained at several locations on the platform and in the blade tip region. The tip region consists of the bottom of the recess, the lip region (on both the pressure and suction surface sides of the recess), and the 90 percent span location on the blade suction surface. Measurements were obtained for three vane/blade spacings; 20, 40, and 60 percent of vane axial chord and for a single value of the tip gap (the distance between the top of the lip and the stationary shroud) equal to 0.0012 m (0.046 in) or 2.27 percent of blade height. [S0889-504X(00)00604-8]

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