Numerical Investigations on the Aerodynamic Performance and Endwall Cooling Characteristics of Turbine During Acceleration Process With Lagging Effects

In the process of turbine acceleration, due to the influence of compressor and complex secondary air system, the change process of coolant purge flow is relatively lagging behind that of mainstream flow and rotational speed. The lagging egress of coolant flow influence the aerodynamic performance and endwall cooling effectiveness of turbine acceleration process. The flow field and aerothermal performance of two-stage axial turbines combined with rim seal structures and coolant purge flow lagging effects in the turbine acceleration process was numerically investigated using Unsteady Reynolds-Averaged Navier-Stokes (URANS) via SST turbulence model. The effects of lagging coolant purge flow across the rim seal on the turbine aerodynamics and endwall cooling effectiveness were analyzed. The obtained results show that the turbine aerodynamic efficiency obtains the maximum value when the coolant purge flow lagging time equals to half the acceleration time at the same rotational speed after the end of lagging times. The total-to-total efficiency for the second stage is more sensitive to lagging times. The turbine output power is almost un-changed due to combination of additional work capacity and aerodynamic loss with the introduction of coolant. The turbine endwalls have the maximum averaged cooling effectiveness in the turbine acceleration process without consideration of the coolant purge flow lagging time. And endwall cooling effectiveness decreases with the increase of coolant purge flow lagging time at the same rotational speed and mainstream flow conditions. The detailed flow field of two-stage turbine considering interaction between the coolant purge flow and mainstream was also discussed. The present work provides the reference for the match design between the turbine mainstream flow and secondary air flow system.

Endwall Heat Transfer and Cooling Performance of a Transonic Turbine Vane With Upstream Injection Flow

Flow fields near the turbine vane endwall region are very complicated due to the presence of highly three-dimensional passage vortices and endwall secondary flows. This makes it challenging for the endwall to be effectively cooled by employing traditional endwall cooling methods, such as impingement cooling combined with local film cooling inside the vane passage. One effective endwall cooling scheme: coolant injection flow through discrete holes upstream of the vane leading edge on the endwall, has been considered by many gas turbine companies. The present paper focuses on endwall film cooling effectiveness evaluation with upstream coolant injection through discrete holes. Detailed experimental and numerical studies on endwall heat transfer and cooling performance with coolant injection flow through upstream discrete holes is presented in this paper. High resolution heat transfer coefficient (HTC) and adiabatic film cooling effectiveness values were measured using a transient infrared thermography technique on an axisymmetric contoured endwall. The endwall tested was a scaled up inner endwall of an industrial transonic turbine vane with double-row discrete cylindrical film cooling holes located 0.39Cx upstream of the vane leading edge. The tests were performed in a transonic linear cascade blow-down wind tunnel facility. Conditions were representative of a land-based power generation turbine with exit Mach number of 0.85 corresponding to exit Reynolds number of 1.5 × 106, based on exit condition and axial chord length. A high turbulence level of 16% with an integral length scale of 3.6%P was generated using inlet turbulence grid to reproduce the typical turbulence conditions in real turbine. Low temperature air was used to simulate the typical coolant-to-mainstream condition by controlling two parameters of the upstream coolant injection flow: mass flow rate to determine the coolant-to-mainstream blowing ratio (BR = 2.5, 3.5), and gas temperature to determine the density ratio (DR = 1.2). To highlight the interactions between the upstream coolant flow and the passage secondary flow combined with the influence on the endwall heat transfer and cooling performance, a comparison of CFD predictions to experimental results was performed by solving steady-state Reynolds-Averaged Navier-Stokes (RANS) using the commercial CFD solver ANSYS Fluent v.15. A detailed numerical method validation was performed for four different Reynolds-averaged turbulence models. The Realizable κ-ϵ model was validated to be suitable to obtain reliable numerical solution. The influences of a wide range of coolant-to-mainstream blowing ratios (BR = 1.0, 1.5, 1.9, 2.5, 3.0, 3.5) were numerically studied. Complex interactions between coolant injections and secondary flows in vane passage were presented and discussed. Results indicate that for lower values of BR, the endwall coolant coverage from the upstream double-row discrete holes is strongly controlled by the passage secondary flow, thus the cooling effectiveness is very poor. As the BR increases, the strong secondary flow in vane passage can be suppressed by the coolant injections and begin to be almost eliminated when BR increases to a critical value (BR = 2.5 – 3.0). Beyond the critical BR, most of the injected coolant begins to lift off from the endwall and penetrate significantly into the mainstream flow, yielding inefficient endwall cooling performance.

Turbine Vane Endwall Film Cooling Comparison From Five Film-Hole Design Patterns and Three Upstream Injection Angles

The effects of upstream injection angle on film cooling effectiveness of a turbine vane end wall with various endwall film-hole designs were examined by applying pressure-sensitive paint (PSP) measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create an additional controllable upstream film to cool the vane end wall. Three potential injection angles were studied: 30 deg, 40 deg, and 50 deg. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, midchord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio (MFR) 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream injection angle.

Endwall Film Cooling Performance for a First-Stage Guide Vane With Upstream Combustor Walls and Inlet Injection

Effects of an upstream combustor wall on turbine nozzle endwall film cooling performance are numerically examined in a linear cascade in this paper. Film cooling is by two rows of cooling holes at 20% of the axial chord length upstream of the vane leading edge (LE) plane. The combustor walls are modeled as flat plates with square trailing edges (TE) positioned upstream of the endwall film cooling holes. A combustor wall is in line with the LE of every second vane. The influence of the combustor wall, when shifted in the axial and tangential directions, is investigated to determine effects on passage endwall cooling for three representative film cooling blowing ratios. The results show how shed vortices from the combustor wall greatly alter the flow field near the cooling holes and inside the vane passage. Film cooling distribution patterns, particularly in the entry region and along the pressure side of the passage, are affected. The combustor wall leads to an imbalance in film cooling distribution over the endwalls for adjacent vane passages. Results show a larger effect of tangential shift of the combustor wall on endwall cooling effectiveness than the effect of an equal axial shift. The study provides guidance regarding design of combustor-to-turbine transition ducts.

Turbine Vane Endwall Film Cooling Comparison From Five Film-Hole Design Patterns and Three Upstream Leakage Injection Angles

The effects of upstream leakage injection angle on film cooling effectiveness of a turbine vane endwall with various endwall film-hole designs were examined by applying PSP measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create additional controllable upstream leakage flow to cool the vane endwall. Three potential leakage injection angles were studied: 30°, 40°, and 50°. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, mid-chord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream leakage injection angle.

Film Cooling Effectiveness Comparison on Turbine Vane Endwall With Cluster Configurations Using PSP Measurement Technique

Film cooling performance comparison on a turbine vane endwall with two different cluster configurations incorporating upstream inlet leakage flow was studied by applying PSP measurement technique. Experiments were conducted in a blow-down wind tunnel cascade facility at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000, based on axial chord length. A freestream turbulence level was generated as high as 19% with an integral length scale of 1.7 cm at the inlet of vanes’ leading edge plane. The results of each design and the comparison are presented in two-dimensional adiabatic film cooling effectiveness distributions on the endwall surface as well as the corresponding laterally averaged distributions. The focus of this study is to compare two different endwall cooling designs by the parametric effect of coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR) on film cooling effectiveness. Initially, increased MFR for the endwall cooling and upstream leakage levels up the local film cooling effectiveness and yields improved coverage on the endwall. However, the endwall and upstream leakage demonstrate different behavior, while highest MFR does not provide any improvement as endwall cooling is suffered from the jet lift-off, highest MFR from the upstream leakage shows a monotonic increasing effectiveness on the endwall. On the other hand, the DR effect on effectiveness reveals different trend at a different design. Under a given amount of coolant supply and a fixed number of cooling hole, the comparison results really emphasize the importance of the cooling hole arrangement as one of the proposed cluster design can properly cover the whole endwall area.