Aerothermal Characteristics Research Articles

Analysis of Aerothermal Characteristics of Surface Microstructures

Recent advances in manufacturing technologies, such as additive manufacturing (AM), have raised the potential of choosing surface finish pattern as a design parameter. Hence, understanding and prediction of aerothermal effects of machined microstructures (machined roughness) would be of great interest. So far, however, roughness has been largely considered as a stochastic attribute and empirically modeled. A relevant question is: if and how would shape of the machined roughness elements matter at such fine scales? In this paper, a systematic computational study has been carried out on the aerothermal impact of some discrete microstructures. Two shapes of configurations are considered: hemispherical and rectangular elements for a Reynolds number range typical for such structures (Re < 5000). Several validation cases are studied as well as the turbulence modeling and grid sensitivities are examined to ensure the consistency of the results. Furthermore, large eddy simulation (LES) analyses are performed to contrast the behavior in a well-established turbulent to a transitional flow regime. The results reveal a distinctive common flow pattern change (from an “open separation” to a “reattached separation”) associated with a drastic change of drag correlation from a low to a high loss regime. The results indicate a clear dependence of drag and heat transfer characteristics on the element pattern and orientation relative to the flow. The distinctive performance correlations with Reynolds number can be affected considerably by the element shape, for both a transitional and a turbulent flow regime. The results also consistently illustrate that conventional empirical stochastic roughness parameters would be unable to predict these trends.

Cooling Injection Effect on a Transonic Squealer Tip—Part I: Experimental Heat Transfer Results and CFD Validation

Recent studies have demonstrated that the aerothermal characteristics of turbine rotor blade tip under a transonic condition are qualitatively different from those under a low-speed subsonic condition. The cooling injection adds further complexity to the over-tip-leakage (OTL) transonic flow behavior and aerothermal performance, particularly for commonly studied shroudless tip configurations such as a squealer tip. However there has been no published experimental study of a cooled transonic squealer. The present study investigates the effect of cooling injection on a transonic squealer through a closely combined experimental and CFD effort. Part I of this two-part paper presents the first of the kind tip cooling experimental data obtained in a transonic linear cascade environment (exit Mach number 0.95). Transient thermal measurements are carried out for an uncooled squealer tip and six cooling configurations with different locations and numbers of discrete holes. High-resolution distributions of heat transfer coefficient and cooling effectiveness are obtained. ansysFluent is employed to perform numerical simulations for all the experimental cases. The mesh and turbulence modeling dependence is first evaluated before further computational studies are carried out. Both the experimental and computational results consistently illustrate strong interactions between the OTL flow and cooling injection. When the cooling injection (even with a relatively small amount) is introduced, distinctive series of stripes in surface heat transfer coefficient are observed with an opposite trend in the chordwise variations on the squealer cavity floor and on the suction surface rim. Both experimental and CFD results have also consistently shown interesting signatures of the strong OTL flow–cooling interactions in terms of the net heat flux reduction distribution in areas seemingly unreachable by the coolant. Further examinations and analyses of the related flow physics and underlining vortical flow structures will be presented in Part II.

Optimization of a high pressure turbine blade tip cavity with conjugate heat transfer analysis

The current study aims to understand the aero-thermal performance of a cooled cavity tip in a single stage transonic turbine. The squealer tip of the uncooled turbine blade was reduced to an aerodynamic loss with suppressing leakage flow. However, the aerodynamic loss study of the cooled turbine blade tip is rare. It is necessary to study the tip cavity of the cooled turbine blade. Depth, front blend radius and aft blend radius of the cavity were set as design variables, and 30 cases were chosen using design of experiments. These cases were calculated with conjugate heat transfer method. Approximation model was made using the Kriging method, and tip cavity shape was optimized with multidisciplinary design optimization. Average total pressure loss behind the trailing edge and cooling effectiveness of blade tip surface were set to the objective function. The aerodynamic optimization model decreased 1.6 % of total pressure loss, the heat transfer optimization model increased 1.3 % point of cooling effectiveness and aero-thermal optimization model were found. Volume of tip cavity becomes larger when three design variables are grown. Amount of tip leakage flow and its distribution over the tip region increases and total pressure loss and cooling effectiveness increase. In terms of heat transfer, blade tip without cavity is advantageous. Total pressure loss coefficient, however, also increases over 5 %. To improve both aero-thermal characteristics of cooled blade tip, the design using the multidisciplinary design optimization is recommended.

On the correspondence between flow structures and convective heat transfer augmentation for multiple jet impingement

The correspondence between local fluid flow structures and convective heat transfer is a fundamental aspect that is not yet fully understood for multiple jet impingement. Therefore, flow field and heat transfer experiments are separately performed investigating mutual–jet interactions exposed in a self-gained crossflow. The measurements are taken in two narrow impingement channels with different cross-sectional areas and a single exit design. Hence, a gradually increased crossflow momentum is developed from the spent air of the upstream jets. Particle image velocimetry (PIV) and liquid crystal thermography (LCT) are used in order to investigate the aerothermal characteristics of the channel with high spatial resolution. The PIV measurements are taken at planes normal to the target wall and along the centreline of the jets, providing quantitative flow visualisation of jet and crossflow interactions. Spatially resolved heat transfer coefficient distributions on the target plate are evaluated with transient techniques and a multi-layer of thermochromic liquid crystals. The results are analysed aiming to provide a better understanding about the impact of near-wall flow structures on the convective heat transfer augmentation for these complex flow phenomena.

Numerical analysis on effects of inlet pressure and temperature non-uniformities on aero-thermal performance of a HP turbine

Unsteady numerical studies were conducted on the effects of combustor exit total pressure non-uniformity and hot streak (HS) on the aero-thermal characteristics of a high-pressure (HP) turbine stage. Firstly, four total pressure profiles including uniform, turbulent boundary layer, realistic engine combustor exit and radial reverse total pressure (higher total pressure near the endwall) were applied to reveal the vortex structures and heat transfer characteristics of the turbine. Then the four total pressure profiles coupled with inlet HS were specified to study the HS migration mechanisms. The results indicated that the turbulent boundary layer total pressure intensifies the passage vortices (PVs) while the higher total pressure near the endwall induces counter rotating vortices (CRVs), in both the stator and rotor passages. The heat transfer on the vane and blade surfaces was affected by the vortices and the extent of the impact directly depends on the intensity of the vortices, which is determined by the gradient of the total pressure. The heat transfer coefficients (HTCs) on the vane and blade pressure surface are increased by 104% and 20–30%, respectively, due to the CRV effects. The increase of tip leakage flow for the higher total pressure near the tip enhances the tip pressure side HTC by 4% comparing with the uniform case. The PVs and CRVs from the stator passage persistently exist in the rotor passage near the suction side endwall while have negligible influence on the rotor heat transfer. The migration paths of the hot streak are significantly affected by the PVs and CRVs in the stator passage, and the rotor blade surface temperature is dominated by both of the redistribution of HS in the vane passage and the vortices in the rotor passage. The PVs extend the HS radial migration and increase the vane and blade surface temperature. While the CRVs circulate the hot fluid away from the vane surface and increase the HS attenuation, thus decrease the vane and blade surface temperature significantly. The total pressure non-uniformity also decreases the turbine adiabatic efficiency by 0.6% at uniform inlet temperature condition.

Aerothermal characteristics of a transonic tip flow in a turbine cascade with tip clearance variations

A significant portion of flows over a modern high-pressure turbine blade tip is transonic, and the transonic tip leakage flows lead to significant aerodynamic losses and high heat loads onto the blade tips. This paper aims to develop a deeper understanding of the transonic tip leakage flow physics and its influence on the loss mechanism and blade-tip heat transfer. Three-dimensional (3D) Reynolds-averaged Navier-Stokes (RANS) calculations were performed using the ANSYS CFX 14.5 numerical prediction code, adopting the SST k-ω turbulence model to investigate the sensitivity of aerothermal performance of transonic tip flows to tip clearances in a RT27a turbine cascade. The transonic tip leakage flow pattern within the tip gap, the flowfield downstream of the cascade, and the blade tip heat transfer distribution are studied. The numerical results give a reasonable agreement with the experimental data. The tip aerodynamics and surface heat transfer variations with tip clearances are opposite between leading and trailing edge regions of blade tips, and tip clearance effects on the former are relatively small. As the tip clearance increases, the shock wave reflections are delayed but more evident, and it therefore leads to reduced leakage massflow density and decreased heat loads on the rear part of blade tips. Despite this, since the leakage flow near the leading edge of blade tips remains subsonic resulting in increased leakage mass flowrate and tip heat transfer, the leakage losses and overtip heat loads are increased with the increasing tip clearance.

Aero-thermal analysis of lifting body configurations in hypersonic flow

The aero-thermal analysis of a hypersonic vehicle is of fundamental interest for designing its thermal protection system. The aero-thermal environment predictions over several critical regions of the hypothesized lifting body vehicle, including the stagnation region of the nose-cap, cylindrically swept leading edges, fuselage-upper, and fuselage-lower surfaces, are discussed. The drag (Λ=70°) and temperature (Λ=80°) minimized sweepback angles are considered in the configuration design of the two hypothesized lifting body shape hypersonic vehicles. The main aim of the present study is to analyze and compare the aero-thermal characteristics of these two lifting body configurations at same heat capacity. Accordingly, a Computational Fluid Dynamics simulation has been carried out at Mach number (M∞=7), H=35km altitude with zero Angle of Attack. Finally, the material selection for thermal protection system based on these predictions and current methodology is described.

An Experimental and Numerical Study on the Aerothermal Characteristics of a Ribbed Transonic Squealer-Tip Turbine Blade With Purge Flow

Detailed heat transfer coefficient (HTC) and film cooling effectiveness (Eta) distribution on a squealer-tipped first stage rotor blade were measured using an infrared technique. The blade tip design, obtained from the Solar Turbines, Inc., gas turbine, consists of double purge hole exits and four ribs within the squealer cavity, with a bleeder exit port on the pressure side close to the trailing edge. The tests were carried out in a transient linear transonic wind tunnel facility under land-based engine representative Mach/Reynolds number. Measurements were taken at an inlet turbulent intensity of Tu = 12%, with exit Mach numbers of 0.85 (Reexit = 9.75 × 105) and 1.0 (Reexit = 1.15 × 106) with the Reynolds number based on the blade axial chord and the cascade exit velocity. The tip clearance was fixed at 1% (based on engine blade span) with a purge flow blowing ratio, BR = 1.0. At each test condition, an accompanying numerical study was performed using Reynolds-averaged Navier–Stokes (RANS) equations solver ansys fluent to further understand the tip flow characteristics. The results showed that the tip purge flow has a blocking effect on the leakage flow path. Furthermore, the ribs significantly altered the flow (and consequently heat transfer) characteristics within the squealer-tip cavity resulting in a significant reduction in film cooling effectiveness. This was attributed to increased coolant–leakage flow mixing due to increased recirculation within the squealer cavity. Overall, the peak HTC on the cavity floor increased with exit Mach/Reynolds number.

Aerothermal characteristics of bleed slot in hypersonic flows

Two types of flow configurations with bleed in two-dimensional hypersonic flows are numerically examined to investigate their aerodynamic thermal loads and related flow structures at choked conditions. One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer interactions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.

Numerical Investigation of Aerothermal Characteristics of the Blade Tip and Near-Tip Regions of a Transonic Turbine Blade

In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the exit Mach number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at exit Mach numbers of Mexit = 0.85 (Reexit = 9.75 × 105) and Mexit = 1.0 (Reexit = 1.15 × 106) with high freestream turbulence (Tu = 12%). The exit Reynolds number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with exit Mach number due to high turbulence generation in the tip gap and flow reattachment. While increase in exit Mach number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the near-tip surfaces affected by leakage vortex. Increase in exit Mach number was found to also induce strong flow relaminarization on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in exit Mach number but significant increase in local heat transfer was noted in this region.

Film Cooling Performance in a Transonic High-pressure Vane: Decoupled Simulation and Conjugate Heat Transfer Analysis

The continuous demand for increased performance and reliability of gas turbines leads to the improvement of prediction tools. Having regard to the effects of heat transfer on the residual life of gas turbine components, it is necessary to achieve a high level of accuracy in the evaluation of thermal loads. Computational fluid dynamics is able to provide reliable data in a limited lapse of time. In this paper, the numerical analysis of the cooled vane of the MT1 high-pressure turbine stage is presented. A grid dependence analysis based on the evaluation of the aero-thermal characteristics of the vane has been performed. Turbulence is modeled using the kT-kL-ω method whose performance in this kind of configuration is rarely debated in the scientific literature. Model parameters have been tuned to match the experimental data. The final objective of the present activity is to assess the capability of numerical methods to deal with an annular, transonic high-pressure vane with a realistic film cooling configuration. Adiabatic effectiveness, heat transfer coefficient and net heat flux reduction distributions have been evaluated, the latter providing relevant information on the performance of the cooling system. The coupled fluid-solid simulation of the cooled configuration has also been performed to evaluate the impact of conjugate heat transfer on the prediction of thermal loads. Results show a non-negligible difference in the wall temperature evaluation between the decoupled and the coupled approach, mainly caused by the heat conduction in the solid.

Note: Thermal imaging enhancement algorithm for gas turbine aerothermal characterization

An algorithm was developed to convert radiation intensity images acquired using a black and white CCD camera to thermal images without requiring knowledge of incident background radiation. This unique infrared (IR) thermography method was developed to determine aerothermal characteristics of advanced cooling concepts for gas turbine cooling application. Compared to IR imaging systems traditionally used for gas turbine temperature monitoring, the system developed for the current study is relatively inexpensive and does not require calibration with surface mounted thermocouples.

Aerodynamic and Aerothermal Performance of Power-law Shaped Leading Edge of Hypersonic Waveriders

A numerical study is presented on three- dimensional power-law shaped leading edge for waveriders in hypersonic flow. This work is motivated by interest in researching the flow field properties of power-law shaped leading edge as a possible candidate for blunting leading edges of waveriders. Both aerodynamic and aerothermal characteristics are calculated for a waverider forebody in hypersonic flow. Comparisons are made between power-law shaped leading edges and round leading edge. For the flow conditions considered, the power-law leading edges of waveriders can provide small drag, and at the same time it can bring enough bluntness for the necessity of thermal protection.

Film Cooling Extraction Effects on the Aero-Thermal Characteristics of Rib Roughened Cooling Channel Flow

The present study is geared towards quantifying the effects of film cooling holes on turbine internal cooling passages. In this regard, tests are conducted in a generic stationary model, with evenly distributed rib-type perturbators at 90 deg, constituting a passage blockage ratio of H/Dh = 0.3 and pitch-to-height ratio of P/H = 10. The 1/3H diameter surface-perpendicular film cooling holes are employed at a distance of 5/3H downstream of the preceding rib. Through liquid crystal thermometry measurements, the aero-thermal effects of a change in suction ratio are contrasted for various configurations (Re = 40,000 SR = 0–6), and compared with the analogous aerodynamic literature, enabling heat transfer distributions to be associated with distinct flow structures. At increased suction ratio, the size of the separation bubble downstream of the rib is observed to diminish, triggering globally an earlier reattachment; in addition to low-momentum hot fluid extraction via film cooling suction. Hence, in the presence of active flow extraction, higher overall heat transfer characteristics are observed throughout the channel. Moreover, the findings are generalized via friction factor and Nusselt number correlations, along with an analytical 20-pitch passage model. SR ∼ 3.5 is observed to provide favorable characteristics of pitch-to-pitch uniform suction ratio, lack of hot fluid ingestion and to sustain the highest passage averaged heat transfer.

Controlling the secondary flows near endwall boundary layer fences in a 90° turning duct using approximate optimization method

This paper presents the optimization of an endwall boundary layer fence which improves the aerothermal characteristics of a gas tur- bine passage. The fence in a gas turbine passage effectively reduces secondary flow loss, which is generated in the cascade of the turbine. The turbine passage was simulated by a 90o turning duct (ReD=360,000). The main purpose of the present investigation was to focus on finding a boundary layer fence with minimum total pressure loss in the passage and heat transfer coefficient on the endwall of the duct. An approximate optimization method was used for the investigation to secure the computational efficiency. Design variables were the fence shape factors (fence thickness, fence height, and fence width) as well as the radial position on the passage endwall. Results indi- cated that a significant improvement in aerodynamic performance can be achieved through the application of an optimized boundary layer fence. The optimized design reduced the mass-weighted average total pressure loss in the duct by 8.6% relative to the duct without the fence. The area-weighted average maximum Nusselt number on the endwall of the optimized design was decreased by 2.5% in com- parison with the duct without the fence.

Experimental and Numerical Investigation of Aerothermal Characteristics of Hypersonic Intermediate Experimental Vehicle

Cp = pressure coefficient h = altitude, km k = transition tripping device height, mm L = vehicle length, m M = Mach number P = pressure, pa q = convective heat flux, kW=m Re = Reynolds number T = temperature, k X = distance along vehicle forebody running from nose, m Y = spanwise vehicle coordinate, m = angle of attack, deg = angle of side slip, deg = surface control deflection angle, deg = emissivity coefficient

The Computational Modeling of Transitional Flow through a Transonic Linear Turbine: Comparative Performance of Various Turbulence Models

Aero-thermal characteristics of transitional flow through a highly loaded transonic linear turbine are studied at the design incidence using computational methods. The three-dimensional compressible turbulent flow through a turbine inlet guide vane is simulated using finite-volume method-based fluid flow solutions using both Reynolds-averaged Navier Stokes (RANS) equations-based turbulence models, and a dynamic large eddy simulation (LES) approach which is based on a Smagorinsky-Lilly subgrid scale (SGS) model. The calculations are made for exit isentropic Mach numbers ranging from 0.7 to 1.1, and for Reynolds numbers ranging from 5 × 105 to 2 × 106. The numerical results obtained from different turbulence model simulations are compared with each other, and with the experimental data available in the open literature in terms of integrated flow parameters, such as pressure and heat transfer coefficients for varying exit Mach and Reynolds numbers. The present numerical approach indicates a guide line that the present LES approach can be a reliable choice for turbomachinery flows.

E108 高圧タービン前縁フィルム冷却における冷却孔形状の効果(OS6 高温・高効率発電),動力エネルギーシステム部門20周年,次の20年への新展開)

This paper deals with the experimental and numerical studies of aero-thermal characteristics of film cooling for turbine blade leading edge having three differently hole geometries. The hole geometries used in this study are called typeA and typeB. The purpose of this study is to sort out the most suitable hole geometry. Mainstream Reynolds number based on the cylinder diameter was 86,000. The effect of blowing rates (BR) was studied of various blowing ratios of 1.0 and 2.0. Transient techniques using thermochromic liquid crystal is used to measure film effectiveness and heat transfer coefficient. In addition, Detailed investigation of film cooling for leading edge is carried out using CFD simulations. RANS applying a Shear Stress Transport turbulence model is addressed to solve thermal convection. Results show that the typeA provide better film cooling performance on BR=1.0 than typeB.

Turbine Airfoil Aerothermal Characteristics in Future Coal—Gas-Based Power Generation Systems

Most promising operating cycles being developed for future coal– gas-based systems are hydrogen-fired cycle and oxy–fuel cycle. Both cycles will likely have turbine working fluids significantly different from those of conventional air-based gas turbines. The oxy–fuel cycle, with steam and CO2 as a primary working fluid in the turbine section, will have a turbine inlet temperature target at approximately 1750 o C, significantly higher than the current level of utility turbine systems. Described in this paper is a CFD-based simulation of the transport phenomena around the gas side of a turbine airfoil under realistic operating conditions of future coal–gas-based systems. The relatively high concentration of steam in the oxy–fuel turbine leads to approximately 40% higher heat transfer coefficient on the airfoil external surface than its hydrogen-fired counterpart. This suggests that advances in cooling technology and thermal barrier coatings (TBC) are critical for the developments of future coal-based turbine systems. To further explore this issue, a comparative study on the internal cooling effectiveness between a double-wall or skin-cooled arrangement and an equivalent serpentine-cooled configuration is performed. The contribution of thermal barrier coatings (TBC) toward overall thermal protection for turbine airfoil cooled under these two different cooling configurations is also evaluated.

Characteristics of effective industrial packings for evaporative cooling of circulating water in cooling towers

Results are presented for comparative tests of the aerothermal characteristics of basic types of regular packings on a test bench developed by the Vedeneev All-Russia Research Institute of Hydraulic Engineering. These packings are widely used in draft sections and similar cooling towers employed in the chemical and other branches of industry. It is demonstrated that the drop-film type of packing is most effective.