Vane Aerodynamics Research Articles

Effect of a Combined Hot-Streak and Swirl Profile on Cooled 1.5-Stage Turbine Aerodynamics: An Experimental and Computational Study

Abstract Recently developed lean-burn combustors offer reduced NOx emissions for gas turbines. The flow at exit of lean-burn combustors is dominated by hot streaks and residual swirl, which have been shown—individually—to impact turbine aerodynamic performance. Studies have shown that residual swirl at inlet to the high-pressure (HP) stage predominantly affects the vane aerodynamics, while hot streaks affect the rotor aerodynamics. Studies have also shown that these changes to the HP stage aerodynamics can affect the downstream intermediate-pressure (IP) vane aerodynamics. Yet, to date, there have been no published studies presenting experimental turbine test data with both swirl and hot streaks simultaneously present at inlet. This paper presents the first experimental and computational investigation into the effects of combined hot streaks and swirl on turbine aerodynamics. Experimental measurements were conducted in the Oxford Turbine Research Facility (OTRF), a short-duration rotating transonic facility, in which the nondimensional parameters relevant to turbine fluid mechanics and heat transfer are matched to engine conditions. The turbine under investigation is the recently commissioned LEMCOTEC turbine, which has been designed to represent modern aero-engine architectures and for robustness to lean-burn combustor-representative inlet flows. The turbine comprises an unshrouded HP stage with fully film-cooled vanes followed by low-turning IP vanes in an S-shaped duct. Two turbine inlet flows are considered. The first is uniform in total pressure, total temperature, and flow angle. The second features a nonuniform total temperature (hot streak) profile featuring strong radial and weak circumferential variation superimposed on a swirling velocity profile. This combined nonuniform profile is generated using a new combustor simulator that has recently been commissioned in the OTRF. Detailed area surveys of the flow were conducted at turbine inlet, HP rotor exit, and IP vane exit, and loading distributions were measured on the HP and IP vanes. Measurements and unsteady Reynolds-averaged Navier–Stokes (URANS) predictions suggest that the inlet temperature nonuniformity was relatively well preserved upon being convected through the turbine: the predicted root-mean-square variation in the IP vane exit total temperature field was approximately double that with uniform inlet conditions. Relatively poor comparisons between URANS and experiment highlight the challenge of accurately predicting the complex IP vane flow. In particular, small differences in exit whirl angle resulted in substantial differences in IP vane exit velocity and thus radial pressure gradient.

Impact of a Combustor–Turbine Interface on Turbine Vane Aerodynamics and Film Cooling

Abstract Optimizing the aerothermal performance of the combustor–turbine interface is an important factor in enhancing the efficiency of heavy-duty gas turbines. Also, it is a key requirement to fulfill the lifetime in this hottest area of the gas turbine. Typically transition pieces of can combustors induce a highly nonuniform swirling flow at the turbine inlet. In order to better understand the impact of the nonuniform combustor flow at the first stage vanes, a combined experimental and numerical study was carried out. The experimental facility consisted of a high-speed linear cascade with four vane passages, including an upstream transition piece, which was representative of a heavy-duty gas turbine can combustor–turbine interface geometry. The experiments were conducted at engine representative Mach numbers, and film cooling effectiveness measurements were performed at three different blowing ratios. Computational fluid dynamics (CFD) Reynolds-averaged Navier–Stokes simulations were undertaken using a commercial flow solver. The numerical model was first validated with the experimental data, using inlet traverse five-hole probe measurements, pressure taps along the airfoil perimeter, and oil flow visualization results. The investigation shows that the position of the vane relative to the combustor transition piece has a significant impact on the vane aerodynamics and also film cooling behavior. This understanding was important for a robust first vane aerothermal design of the GT36.

Numerical and Experimental Investigation of Return Channel Vane Aerodynamics With Two-Dimensional and Three-Dimensional Vanes

The present paper deals with the numerical and experimental investigation of the effect of return channel (RCH) dimensions of a centrifugal compressor stage on the aerodynamic performance. Three different return channel stages were investigated, two stages comprising three-dimensional (3D) return channel blades and one stage comprising two-dimensional (2D) RCH vanes. The analysis was performed regarding both the investigation of overall performance (stage efficiency, RCH total pressure loss coefficient) and detailed flow-field performance. For detailed experimental flow-field investigation at the stage exit, six circumferentially traversed three-hole probes were positioned downstream the return channel exit in order to get two-dimensional flow-field information. Additionally, static pressure wall measurements were taken at the hub and shroud pressure and suction side (SS) of the 2D and 3D return channel blades. The return channel system overall performance was calculated by measurements of the circumferentially averaged 1D flow field downstream the diffuser exit and downstream the stage exit. Dependent on the type of return channel blade, the numerical and experimental results show a significant effect on the flow field overall and detail performance. In general, satisfactory agreement between computational fluid dynamics (CFD)-prediction and test-rig measurements was achieved regarding overall and flow-field performance. In comparison with the measurements, the CFD-calculated stage performance (efficiency and pressure rise coefficient) of all the 3D-RCH stages was slightly overpredicted. Very good agreement between CFD and measurement results was found for the static pressure distribution on the RCH wall surfaces while small CFD-deviations occur in the measured flow angle at the stage exit, dependent on the turbulence model selected.

Effect of Low-NOX Combustor Swirl Clocking on Intermediate Turbine Duct Vane Aerodynamics With an Upstream High Pressure Turbine Stage—An Experimental and Computational Study

Flow in an intermediate turbine duct (ITD) is highly complex, influenced by the upstream turbine stage flow structures, which include tip leakage flow and nonuniformities originating from the upstream high pressure turbine (HPT) vane and rotor. The complexity of the flow structures makes predicting them using numerical methods difficult, hence there exists a need for experimental validation. To evaluate the flow through an intermediate turbine duct including a turning vane, experiments were conducted in the Oxford Turbine Research Facility (OTRF). This is a short duration high speed test facility with a 3/4 engine-sized turbine, operating at the correct nondimensional parameters for aerodynamic and heat transfer measurements. The current configuration consists of a high pressure turbine stage and a downstream duct including a turning vane, for use in a counter-rotating turbine configuration. The facility has the ability to simulate low-NOx combustor swirl at the inlet to the turbine stage. This paper presents experimental aerodynamic results taken with three different turbine stage inlet conditions: a uniform inlet flow and two low-NOx swirl profiles (different clocking positions relative to the high pressure turbine vane). To further explain the flow through the 1.5 stage turbine, results from unsteady computational fluid dynamics (CFD) are included. The effect of varying the high pressure turbine vane inlet condition on the total pressure field through the 1.5 stage turbine, the intermediate turbine duct vane loading, and intermediate turbine duct exit condition are discussed and CFD results are compared with experimental data. The different inlet conditions are found to alter the flow exiting the high pressure turbine rotor. This is seen to have local effects on the intermediate turbine duct vane. With the current stator–stator vane count of 32-24, the effect of relative clocking between the two is found to have a larger effect on the aerodynamics in the intermediate turbine duct than the change in the high pressure turbine stage inlet condition. Given the severity of the low-NOx swirl profiles, this is perhaps surprising.

Design of a Nonreacting Combustor Simulator With Swirl and Temperature Distortion With Experimental Validation

Nonuniform combustor outlet flows have been demonstrated to have significant impact on the first and second stage turbine aerothermal performance. Rich-burn combustors, which generally have pronounced temperature profiles and weak swirl profiles, primarily affect the heat load in the vane but both the heat load and aerodynamics of the rotor. Lean burn combustors, in contrast, generally have a strong swirl profile which has an additional significant impact on the vane aerodynamics which should be accounted for in the design process. There has been a move towards lean burn combustor designs to reduce NOx emissions. There is also increasing interest in fully integrated design processes which consider the impact of the combustor flow on the design of the high pressure vane and rotor aerodynamics and cooling. There are a number of current large research projects in scaled (low temperature and pressure) turbine facilities which aim to provide validation data and physical understanding to support this design philosophy. There is a small body of literature devoted to rich burn combustor simulator design but no open literature on the topic of lean burn simulator design. The particular problem is that in nonreacting, highly swirling and diffusing flows, vortex instability in the form of a precessing vortex core or vortex breakdown is unlikely to be well matched to the reacting case. In reacting combustors the flow is stabilized by heat release, but in low temperature simulators other methods for stabilizing the flow must be employed. Unsteady Reynolds-averaged Navier–Stokes and large eddy simulation have shown promise in modeling swirling flows with unstable features. These design issues form the subject of this paper.

Time-Averaged and Time-Accurate Aerodynamic Effects of Rotor Purge Flow for a Modern, One and One-Half Stage High-Pressure Turbine—Part II: Analytical Flow Field Analysis

The detailed mechanisms of purge flow interaction with the hot-gas flow path were investigated using both unsteady computationally fluid dynamics (CFD) and measurements for a turbine operating at design corrected conditions. This turbine consisted of a single-stage high-pressure turbine and the downstream, low-pressure turbine nozzle row with an aerodynamic design equivalent to actual engine hardware and typical of a commercial, high-pressure ratio, transonic turbine. The high-pressure vane airfoils and inner and outer end walls incorporated state-of-the-art film cooling, and purge flow was introduced into the cavity located between the high-pressure vane and disk. The flow field above and below the blade angel wing was characterized by both temperature and pressure measurements. Predictions of the time-dependent flow field were obtained using a three-dimensional, Reynolds-averaged Navier–Stokes CFD code and a computational model incorporating the three blade rows and the purge flow cavity. The predictions were performed to evaluate the accuracy obtained by a design style application of the code, and no adjustment of boundary conditions was made to better match the experimental data. Part I of this paper compared the predictions to the measurements in and around the purge flow cavity and demonstrated good correlation. Part II of this paper concentrates on the analytical results, looking at the primary gas path ingestion mechanism into the cavity as well as the effects of the rotor purge on the upstream vane and downstream rotor aerodynamics and thermodynamics. Ingestion into the cavity is driven by high static pressure regions downstream of the vane, high-velocity flow coming off the pressure side of the vane, and the blade bow waves. The introduction of the purge flow is seen to have an effect on the static pressure of the vane trailing edge in the lower 5% of span. In addition, the purge flow is weak enough that upon exiting the cavity, it is swept into the mainstream flow and provides no additional cooling benefits on the platform of the rotating blade.

HP Vane Aerodynamics and Heat Transfer in the Presence of Aggressive Inlet Swirl

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability. The flow at combustor exit can therefore have high residual swirl. A good deal of research concerning the flow within the combustor is available in open literature. The impact of swirl on the aerodynamic and heat transfer characteristics of an HP turbine stage is not well understood, however. A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK. The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern. At the turbine inlet plane, yaw and pitch angles of over ±40 deg have been simulated. The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the nondimensional parameters for aerodynamics and heat transfer are matched to engine conditions. The research turbine was the unshrouded MT1 design. By design, the center of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2. For the current investigation, the clocking position was such that the vortex center was aligned with the vane leading edge (every second vane). Both the aligned vane and the adjacent vane were characterized. This paper presents measurements of HP vane surface and end wall heat transfer for the two vane positions. The results are compared with measurements conducted without swirl. The vane surface pressure distributions are also presented. The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra. The aerodynamic and heat transfer characterization presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors. The findings not only have implications for the vane aerodynamic design, but also for the cooling system design.

Effects of Combustor Exit Profiles on Vane Aerodynamic Loading and Heat Transfer in a High Pressure Turbine

The flow and thermal fields exiting gas turbine combustors dictate the overall performance of the downstream turbine. The goal of this work was to investigate the effects of engine representative combustor exit profiles on high pressure turbine vane aerodynamics and heat transfer. The various profiles were produced using a nonreacting turbine inlet profile generator in the Turbine Research Facility (TRF) located at the Air Force Research Laboratory (AFRL). This paper reports how the pressure loading and heat transfer along the vane surface was affected by different turbine inlet pressure and temperature profiles at different span locations. The results indicate that the inlet total pressure profiles affected the aerodynamic loading by as much as 10%. The results also reveal that the combination of different total pressure and total temperature profiles significantly affected the vane heat transfer relative to a baseline test with uniform inlet total pressure and total temperature. Near the inner diameter endwall, the baseline heat transfer was reduced 30–40% over the majority of the vane surface. Near the outer dimeter endwall, it was found that certain inlet profiles could increase the baseline heat transfer by 10–20%, while other profiles resulted in a decrease in the baseline heat transfer by 25–35%. This study also shows the importance of knowing an accurate prediction of the local flow driving temperature when determining vane surface heat transfer.

The Influence of Large-Scale, High-Intensity Turbulence on Vane Aerodynamic Losses, Wake Growth, and the Exit Turbulence Parameters

An experimental research program was undertaken to examine the influence of large-scale high-intensity turbulence on vane exit losses, wake growth, and exit turbulence characteristics. The experiment was conducted in a four-vane linear cascade at an exit Reynolds number of 800,000 based on chord length and an exit Mach number of 0.27. Exit measurements were made for four inlet turbulence conditions including a low-turbulence case (Tu ≈ 1 percent), a grid-generated turbulence case (Tu ≈ 7.5. percent) and two levels of large-scale turbulence generated with a mock combustor (Tu ≈ 12 and 8 percent). Exit total pressure surveys were taken at two locations to quantify total pressure losses. The suction surface boundary layer was also traversed to determine losses due to boundary layer growth. Losses occurred in the core of the flow for the elevated turbulence cases. The elevated free-stream turbulence was found to have a significant effect on wake growth. Generally, the wakes subjected to elevated free-stream turbulence were broader and had smaller peak velocity deficits. Reynolds stress profiles exhibited asymmetry in peak amplitudes about the wake centerline, which are attributable to differences in the evolution of the boundary layers on the pressure and suction surfaces of the vanes. The overall level of turbulence and dissipation inside the wakes and in the free stream was determined to document the rotor inlet boundary conditions. This is useful information for assessing rotor heat transfer and aerodynamics. Eddy diffusivities and mixing lengths were estimated using X-wire measurements of turbulent shear stress. The free-stream turbulence was found to strongly affect eddy diffusivities, and thus wake mixing. At the last measuring position, the average eddy diffusivity in the wake of the high-turbulence close combustor configuration (Tu ≈ 12) was three times that of the low turbulence wake.

Theoretical Improvement of Wind Turbine Output by the Addition of Tip Devices

HE objective of this article is to present a method of analysis and some comparisons of the aerodynamic performance of wind turbines with and without tip devices. Van Holten1 presents some preliminary theoretical and experimental observations on the effect of placing tip vanes on wind turbine rotor blades. More recently, Lissaman and Walker2 have published an extensive report on the dynamic inducer. Both of these reports show that the tip vane (referred to by Lissaman and Walker as dynamic inducer) acts very much like a duct or shroud around a propeller. An analysis of the tip vane aerodynamics shows that the resulting aerodyanmic force on the tip vane should be towards the axis of rotation. This results in a circulation at the rotor tips which will increase the flow through the rotor disk. The presence of the vanes induces additional mass flow through the disk area, thus absorbing more kinetic energy from the freestream than it would without the vanes. The above observations lead us to conclude that with suitably placed tip vanes (Fig. 1) there should be an improvement in the power output from a windmill. This Note presents the results of a vortex lattice program for the rotor with tip vanes. Vortex Lattice Model Sullivan3 has presented a vortex lattice model for analyzing complex propeller configurations. This method is adapted for a theoretical analysis of the windmill rotor with tip vanes. Briefly, the votex lattice method treats the horseshoe vortex as the basic lift-producing unit within the flowfield. The bound vortex elements are placed at the quarter chord of the blade with the trailing vortices coiling into a helical vortex pattern, whose pitch is defined by the freestream velocity and the rotational velocity of the rotor blade. The induced velocity contribution at the control points, which are located at the three-quarter chord point of the blade airfoil section, is the sum of the contribution from the bound vortices and the trailing helical vortices. The contribution from the bound

Aerodynamic Effects of an Attitude Control Vane on a Tilt-Nacelle V/STOL Propulsion System

One possible technique for obtaining longitudinal control on a tilt-nacelle V/STOL aircraft is the use of a variable attitude vane assembly mounted in the propulsion system exhaust. Deflecting the vane produces large forces and moments without depending on forward speed of the aircraft. Tests are carried out in the Ames 40 by 80 ft wind tunnel on a large-scale tilt-nacelle V/STOL propulsion system with and without a variable attitude control vane assembly. Aerodynamic characteristics are analyzed in terms of nacelle aerodynamics, vane aerodynamics, and vane-induced effects on the nacelle aerodynamics. It is shown that the aerodynamic forces due to the nacelle without the vane can be a significant part of the total forces produced by the propulsion system. The control vane effectively produces large changes in pitching moment which are accompanied by significant changes in total lift and drag. The vane has a substantial effect on the propulsion-system aerodynamics. Other pertinent results are also given.