Endwall Surface Research Articles

Effect of profiled endwall on nozzle guide vane with coolant purge flow

The aerothermal performance of a gas turbine could be improved by applying endwall profiling to the nozzle guide vane. In this study, various thermal performance factors of flat and profiled endwalls with purge coolant flow were investigated using adequate experimental methods. Each experiment was conducted under same exit Reynolds number (300,000) for the flat and profiled endwalls. Because secondary vortices in the nozzle passage were weakened by the endwall profiling, the upstream purge coolant covered the endwall surface wider and further with the profiled endwall. Moreover, the heat transfer coefficient in upstream region could be reduced with the weakened secondary vortex. The experimental results indicated that the area-averaged film cooling effectiveness on the profiled endwall was 67.7 % higher and the area-averaged Sherwood number was 18.3 % lower than those on the flat endwall. Although the downstream region of the profiled endwall exhibited a negative cooling performance, the overall area-averaged net heat flux reduction was 169 % higher than that on the flat endwall. Thus, if an additional cooling system was applied to the downstream region of the passage, the profiled endwall could effectively improve not only the aerodynamic performance but also the cooling performance of the nozzle guide vane.

An Investigation of Secondary Cooling Effect on the Endwall Surface With Vane Trailing Edge Coolant Injection

Abstract Considering the adequate use of coolant, the secondary cooling effect of trailing edge coolant injection on the downstream endwall surface is investigated in this study. Distributions of adiabatic cooling effectiveness on the endwall surface are obtained through pressure-sensitive paint (PSP) technique in a linear cascade at four mass flowrate ratios (MFR = 1%, 2%, 3%, and 4%) and two density ratios (DR = 1.0 and 1.5). The configurations of trailing edge pressure-side cutback (PC) and central cutback (CC) with three compound angles (β = 15 deg, 30 deg, and 45 deg) are implemented in this study. Reynolds-averaged Navier–Stokes simulations are performed to present the flow field. Results show that the geometry of the cutback slot has a significant effect on the endwall cooling performance. As the compound angle increases, the coolant coverage is expanded in width. A higher density ratio leads to a decrease in the area of coolant coverage, while the distribution is more uniform. Generally, with a higher mass flow ratio, the coolant coverage is greatly improved in area and value, which almost covers the entire downstream endwall surface.

Aerothermal Performance of Slashface Leakage With Double Surface Angle Under Realistic Swirling Inflow

Abstract The turbine vanes being manufactured and assembled piece by piece leads to an unavoidable gap on the vane endwall called slashface. Usually, the coolant leakage introduced from the slashface with a single surface angle aims to prevent hot gas ingression and protect the vane endwall. However, the cooling effectiveness is greatly reduced when considering the swirling flow from the combustor. In the current work, against strong swirling inflow, a new solution to improve endwall film cooling performance through applying the slashface leakage with the double surface angle is carried out. Through numerical method, the effects of the location of the angle transition region (D/L = 0%, 25%, 50%, 75%, 100%) and mass flow ratio (MFR = 0.5%, 1.0%, 1.5%) of the leakage on the turbine vane endwall aerothermal and film cooling performance are investigated. The results indicated that the coverage area of the coolant on the endwall was obviously enlarged when the slashface leakage with the double surface angle was employed. The film cooling level on both the endwall and suction side surface of the vane increased with increasing MFR. Moreover, the endwall thermal load was reduced. The designs of D/L = 25% and 50% are recommended for their high endwall film cooling level and low endwall thermal load. This paper provides turbine designers with a new idea to increase endwall film cooling performance by the slashface leakage with the double surface angle when considering aggressive swirling inflow.

Experimental and numerical investigations of cooling characteristics on endwall partitioned film cooling with shaped holes

The growing demand for high efficiency in gas turbines requires elevated turbine inlet temperatures, underscoring the critical necessity of implementing high-efficiency film cooling technology on the turbine endwall. However, the traditional cooling layout encounters challenges in the under-cooled areas on the turbine endwall. Therefore, a partitioned film cooling layout with shaped holes on the endwall was proposed based on the heat transfer and flow characteristics of the endwall and the cooling properties of the shaped holes. Based on the heat transfer coefficient distribution on the endwall surface, the endwall was divided into four regions, arranged with different shaped holes for cooling. Numerical and experimental methods were used to study the aerodynamic and cooling characteristics of the partitioned cooling layout. The results indicated that the partitioned cooling layout's cooling performance was significantly improved compared with that of the conventional film hole arrangement. Additionally, it was verified that adjusting the compound angle of the shaped holes further provided a uniform film coverage. More importantly, by adjusting the mass flow of the coolant, the optimized blowing ratios of different regions on the partitioned film cooling layout were explored. The optimized blowing ratio scheme significantly enhanced the film cooling effectiveness due to the varying sensitivity of different shaped holes to the blowing ratio. The results of this work hold significance for advancing film cooling effectiveness and turbine performance.

EFFECTS OF CONTINUOUS AND FRACTURED RIBS PLACED AT THE ENDWALL SURFACE UPSTREAM OF THE TURBINE-VANE LEADING-EDGE

Total pressure loss, chaotic and heavy flow downstream of the turbine passage exit area are a few effects of the desired increase in inlet temperature input at the upstream part of the turbine vane passage. Huge amount of energy from the cross-flows near the middle of the vane actual chord are transported downstream by this chaotic flow which often results in aerodynamic loss and differential heat penalties. To compare the different arrangement, comparisons of the two geometries, the impact of adding various rib configurations upstream are studied for heat augmentation efficacies. The leading-edge rib(s) effects at the mid-stream of the vane passage are quantitatively examined in this work using computational fluid dynamics (CFD) technique. With the employment of Reynolds Averaged Navier-Stoke energy equations, the geometries are modelled choosing the appropriate boundary conditions. Polyhedral mesh is used with fifteen prism layer mesh at the endwall and vane surfaces to capture the flow physics close to the endwall area. Although the two configurations employed showed relative reduction in the total pressure loss coefficient, however, the fractured ribs produced superior outcomes of Nusselt number reduction of over 10% along the passage exit region. The data demonstrate a considerable difference in the overall pressure loss between the two configurations.

Experimental investigation on cooling performance and heat transfer characteristics of the ribbed endwall with upstream leakage

In this paper, the merits and demerits of endwall modification by designing micro-scale ribs onto the turbine endwall surface was examined experimentally for two aspects: cooling performance and heat transfer characteristics. The experiment was conducted on ribbed endwall with three rib directions (horizontal, vertical and tiled rib) and two rib spacings (8 mm and 10 mm) and compared with the flat endwall. The measurement of heat transfer was conducted with a high-resolution infrared thermal camera and that of cooling effectiveness (η) was conducted with pressure-sensitive paint technique. Results showed that the direction of rib and rib-induced vortex greatly affects the coverage of coolant, which the tilted and vertical ribbed endwall demonstrated positive effects on film cooling. Additionally, the heat transfer of the ribbed endwall is also connected with the layout of the rib and mainly depends on the joint influence of the rib-induced vortex and passage vortex. The Nusselt number (Nu) of the vertical rib was the smallest. In general, vertical rib with larger rib spacing shows the comprehensive performance, that an increase of 9.5% for area-averaged η than original endwall with mass flow rate of 1.0% and the least increase of Nu, which can provide guidelines for future cooling design.

A New Perspective on Internal Turbine Blade Cooling by Perforated Blockages and Pin-Fins

Abstract The cooling configuration, sequentially combining perforated blockages (forming blockage jets) and a pin-fin array inside the trailing-edge of a turbine blade has been perceived unsuitable due to the presumed inferior thermal performance (η < 1.0). In the present study, we provide a new perspective on this particular cooling configuration, based on fluidic mechanisms, newly established in a better representative setup for blockage jets aligned with pin-fins, accounting for relevant heat transfer surfaces. To this end, heat transfer on the blockage, pin-fin, and end-wall surfaces was measured at a selected Reynolds number of ReD = 26,000 using a thermochromic liquid crystal technique. Flow field mapping by particle image velocimetry and oil–dye flow visualization were supplementally performed. We demonstrate, contrary to previous studies that the thermal performance of the blockage pin-fin configuration can be e.g., η = 1.1 if the blockages and pin-fins are arranged to maximize both elements' thermofluidic advantages. Our data further suggest that unlike conventional pin-fin configurations subjected to uniform coolant stream, the blockage pin-fin configuration can offer a better performance with fewer pin-fin rows used.

Numerical investigations of secondary cooling on endwall with various cutback slot configurations

The present study numerically investigated the secondary effect of trailing edge film cooling on the downstream endwall surface with various cutback slot configurations. The factors of mass flow rate ratio (MFR = 1–4 %), compound angle (α = 15°, 30° and 45°), and trailing edge cutback position (pressure-side cutback (PC) and central cutback (CC)) are considered. The distributions of film cooling effectiveness (η) and Nusselt number (Nu) are both presented in this study. The findings reveal a significant impact of MFR and compound angle on the spanwise momentum of the coolant, which alters the endwall film cooling effectiveness and heat transfer. An increase in MFR consistently leads to a notable enhancement of endwall film cooling effectiveness and heat transfer. An increase in the compound angle enhances endwall heat transfer, while its influence on the endwall film cooling effectiveness is intricate and varies between the upstream and downstream endwall. Ultimately, there exists an optimal combination of MFR and compound angle that minimizes net heat flux ratio (NHFR). Furthermore, CC configurations exhibit superior secondary cooling on endwall and lower pressure loss compared to PC configurations, attributed to the reduced impact of the mainstream on the coolant flowing from CC. The data are helpful for a deeper insight into the secondary effect of film cooling and may provide a reference for future research.

Research on modeling method of non-axisymmetric endwall based on body-fitted coordinates

Due to the large bending of turbine blades, there is strong three-dimensional flow in cascade channel, the three-dimensional flow at the endwall of blades generates extremely complex heat transfer characteristics. The design concept of high heat load turbine blades, particularly comprehensive optimization of flow and heat transfer at the endwall, has gained increasing attention. In this paper, we have developed a three-dimensional endwall parameterized modeling platform based on Body-Fitting Coordinates and introduced a novel approach for streamwise direction and normal direction concerning endwall profile construction. The endwall modeling was optimized using Genetic Algorithm to obtain a modified one with low heat transfer level on the premise that the aerodynamic loss is not increased. The analysis indicates that the reduction of aerodynamic loss with the modified endwall is relatively marginal, whereas the heat transfer intensity on the endwall surface experiences a significant decrease. Additionally, employing an endwall model can effectively mitigate the lateral pressure gradient in the cascade channel, which exerts a pronounced influence on suppressing secondary vortex development in the cascade.

Numerical Study of Endwall Modification with Micro-Scale Ribs in a Turbine Cascade

A novel modification method, the ‘micro-scale’ rib, is proposed to expand cooling coverage for turbine endwalls. However, the introduction of the rib will inevitably affect the flow in the near-wall region. Therefore, the variation in the flow pattern for the traditional model of secondary flow needs further exploration. In this paper, to gain a clearer understanding of the micro-scale rib, the original endwall and three types of ribbed endwalls were adopted to numerically present the detailed flow, film cooling, and heat transfer information for the endwall surface and phantom cooling on the suction side (SS) of the blade. The Ansys code CFX was utilized to solve the 3D Reynolds-averaged Navier–Stokes (RANS) equations, and the SST k-ω was selected as the turbulence model after the verification. The results show that the rib-like vortex changed the flow of the coolant and had various impacts on the cooling characteristics. Although the cooling performance of the ribbed endwall improved, it also had a negative impact on heat transfer in most cases. Compared with the original, the vertical rib cases provided optimal film cooling, with increases of 26.9% and 17.4% for rib spacing values of 8 mm and 10 mm, respectively, with little difference in heat transfer (less than 1%). In addition, the horizontal rib cases presented the worse performance for both film cooling and heat transfer, which indicates that the rib layout should consider a mainstream flow direction for future designs.

Static pressure redistribution mechanism of non-axisymmetric endwall based on radial equilibrium

Non-axisymmetric endwall contouring has been proved to be an effective flow control technique in turbomachinery. Several different flow control mechanisms and qualitative design strategies have been proposed. The endwall contouring mechanism based on the flow governing equations is significant for exploring the quantitative design strategies of the non-axisymmetric endwall contouring. In this paper, the static pressure redistribution mechanism of endwall contouring was explained based on the radial equilibrium equation. A quantified expression of the static pressure redistribution mechanism was proposed. Compressor cascades were simulated using an experimentally validated numerical method to validate the static pressure redistribution mechanism. A geometric parameter named meridional curvature (Cme) is defined to quantify the concave and convex features of the endwall. Results indicate that the contoured endwall changes the streamline curvature, inducing a centrifugal acceleration. Consequently, the radial pressure gradient is reformed to maintain the radial equilibrium. The convex endwall represented by positive Cme increases the radial pressure gradient, decreasing the endwall static pressure, while the concave endwall represented by negative Cme increases the endwall static pressure. The Cme helps to establish the quantified relation between the change in the endwall radial pressure gradient and the endwall geometry. Besides, there is a great correlation between the distributions of the Cme and the change in the endwall static pressure. It can be concluded that the parameter Cme can be considered as a significant parameter to parameterize the endwall surface and to explore the quantitative design strategies of the non-axisymmetric endwall contouring.

Effect of upstream slot leakage on turbine endwall film cooling characteristics

To investigate the effect of the upstream slot leakage on the endwall film cooling characteristics, the PSP (Pressure Sensitive Paint) technology was applied to measure the turbine blade endwall film cooling characteristics of upstream leakage flow on a linear cascade, and the aerodynamic performance of the passage was numerically analyzed. The effect of the upstream slot leakage mass flow ratio (0.4%∼1.5%) and its position (0.2Cax∼0.5Cax) on the film cooling and the aerodynamic performance were analyzed. The results showed that a rising mass flow ratio of the upstream slot could enhance the film cooling effect at the turbine endwall, increase the coverage area of coolant on the endwall surface, weaken the passage vortex strength, and reduce the aerodynamic losses. While under the same mass flow ratio, reducing the distance between the slot and the blade leading edge would increase the film cooling effect, and inhibit the entrainment effect of the passage vortex on the coolant, but will not influence the leakage coverage area. The distribution of leakage flow is dominated by the secondary flow. When the slot is located on 0.3Cax in front of the passage inlet, the aerodynamic loss of the linear cascade is the minimum in all cases.

A comparative study of purge slot exit shape on the film cooling effectiveness of a gas turbine shroud

To achieve higher efficiency and performance, a modern gas turbine withstands the extremely high thermal environment. In the turbine endwall passage, a high and non-uniform heat transfer appears due to various secondary flow. The purge flow injected through the upstream slot between the stationary and rotating pars can thermally protect the endwall surface by forming air film layer on it. However, since an excessive amount of purge flow reduce the system efficiency, it is needed to effectively inject the purge flow with a limited amount. Accordingly, in the present study, the effect of purge slot exit shape on the film cooling effectiveness (hereby FCE) of a shroud surface was discussed. As nondimensional design parameters of the purge slot, L/Ws, RW/Ws, and RL/Ws were defined, representing the geometric features: the slot width (Ws), the fillet radius at the windward side (RW) and fillet radius at the leeward side (RL) of slot exit, respectively. The measurement of the FCE was performed in a five-blade linear cascade rig using the PSP (Pressure sensitive paint) technique. The results showed that the FCE increased for wider slot width at a given blowing ratio condition. In addition, the fillet radius at the windward side (RW) has a negative effect on the FCE regardless of blowing ratio condition since the purge flow was easily disturbed by the accelerated mainstream injection. On the other hand, the Case 7 with the RW/Ws = 0.9 showed 5.1 % of the improvement of FCE at a higher blowing ratio M = 0.25 compared with the Case 2 with RW/Ws = 0, which represents the sharp edge at the windward side of the slot exit.

Film Cooling on Turbine Vane Endwalls With Different Inlet Cooling Configurations: Experimental and Computational Results

Abstract A comparative study of turbine endwall film cooling resulting from three different cooling configurations in front of the passage inlet was conducted. The three inlet cooling configurations considered in this study were a conventional continuous slot, double rows of discrete film holes, and an interrupted slot with a backward-facing step. Pressure sensitive paint (PSP) was sprayed over the endwall surfaces to map adiabatic cooling effectiveness contours and five-hole probe and thermocouple measurements were implemented to assess aero-thermal fields at the passage exit. The evolution of cooling effectiveness was compared across a full range span of coolant flow rates of 0.3−1.8% in a real engine. Additionally, complex flow structures inside the passage due to interactions of coolant with mainstream flows were visualized using computational fluid dynamics (CFD) simulations to support the observed coolant coverage patterns. A side-by-side comparison of CFD simulations against experiments was made as well to evaluate the reliability of conventional turbulence modeling methods in such a complex flow. The continuous slot was found to produce the highest cooling effectiveness values while the double-row holes offered the most uniform coolant coverage, leading to spreading out of the coolant jets downstream of the passage throat and even beyond the trailing edge of the passage for coolant injection rates higher than 1.0%. Supported by the CFD-predicted flow structures and measured aerodynamic losses, the injection from the double-row holes weakened the passage secondary flows and thus caused the lowest aerodynamic loss. On the contrary, the interrupted slot enhanced the secondary flows due to the backward step, resulting in the highest flow loss.

Turbine Vane Passage Cooling Experiments With a Close-Coupled Combustor–Turbine Interface Geometry Part II: Describing the Coolant Coverage

Abstract The first-stage gas turbine vane surfaces and endwalls require aggressive cooling. This two-part paper introduces a modified design of the combustor–turbine (C–T) interface, the “close-coupled interface,” that is expected to increase cooling performance of vane passage surfaces. While Part I of the paper describes secondary flows and coolant transport in the passage, this part discusses the effects of the new C–T interface geometry on adiabatic cooling effectiveness of the endwall and vane surfaces. Compared to the traditional C–T interface, the coolant requirement is reduced for the same level of cooling effectiveness on all three surfaces for the new C–T interface design, confirming that it is an improvement over the previous design. The endwall crossflow is reduced by combustor coolant injection with the new interface leading to more pitchwise-uniform cooling of the endwall. For the pressure surface, increasing combustor coolant flowrate directly increases phantom cooling effectiveness and spreading of coolant away from the endwall. With the traditional passage vortex seen in the literature replaced by the impingement vortex of the present design, the suction surface receives less phantom cooling than does the pressure surface. However, cooling performance is still improved over that of the previous C–T interface design.

Turbine Vane Passage Cooling Experiments With a Close-Coupled Combustor-Turbine Interface Geometry—Part I: Describing the Flow

Abstract Due to the proximity of the first-stage gas turbine vanes to the combustor, coolant introduced to the combustor walls interacts with the endwall film coolant and changes the vane passage flow physics. Recent results show that combustor coolant contributes significantly to cooling the endwall and vane surfaces. In this paper, the traditional combustor-turbine interface was modified to improve overall cooling performance. The performance of this new injection cooling scheme on passage fluid dynamics and surface cooling is assessed. The first of this two-part paper reports detailed experimental tests that document secondary flows and coolant transport throughout the vane passage for four combustor coolant flowrates. The experimental facility imitates combustor coolant injection and engine-level turbulence and has a modified transition duct design, called the “close-coupled combustor-turbine interface.” The “impingement vortex” seen in previous studies with combustor cooling appears as the dominant secondary flow. It is observed in the present study over a wide range of flowrates, confirming its tie to the combustor coolant flowrate and not the combustor-turbine interface geometry. It was found, however, that the location and size of the impingement vortex are affected by coolant flowrate. Part II of this paper discusses the impact of the observed secondary flows on cooling vane passage surfaces.

Numerical and experimental investigation of quantitative relationship between secondary flow intensity and inviscid blade force in axial compressors

The secondary flow attracts wide concerns in the aeroengine compressors since it has become one of the major loss sources in modern high-performance compressors. But the research about the quantitative relationship between secondary flow and inviscid blade force needs to be more detailed. In this paper, a database of 889 three-dimensional linear cascades was built. An indicator, called Secondary Flow Intensity (SFI), was used to express the loss caused by secondary flow. The quantitative relationship between the SFI and inviscid blade force deterioration was researched. Blade oil flow and Computation Fluid Dynamics (CFD) results of some cascades were also used to cross-validate. Results suggested that all numerical cascade cases can be divided into 3 clusters by the SFI, which are called Clusters A, B and C in the order of the increasing SFI indicator. The corner stall, known as the strong corner separation, only happens when the SFI is high. Both calculations and oil flow experiments show that the SFI would stay at a low level if the vortex core at the endwall surface does not appear. The strong interaction of Kutta condition and endwall cross-flow is considered the dominant mechanism of higher secondary flow losses, rather than the secondary flow penetration depth on the suction surface. In conclusion, the inviscid blade force spanwise deterioration is strongly related to the SFI. The correlation of the SFI and spanwise inviscid blade force deterioration is given in this paper. The correlation could provide a quantitative reference for estimating secondary flow losses in the design.

Control mechanism of vortex-generator jet on turbulent separation in a highly loaded compressor cascade

The active flow control technique of vortex-generator jet (VGJ) was used to control the turbulent separation of a highly loaded compressor cascade at a high incidence. VGJ schemes on suction surface with different jet locations, skewed angles, pitch angles were numerically performed and the control mechanism of VGJ parameters on the flow field of the cascade was revealed. For avoiding the effect of endwall boundary layer, both endwall surfaces of the cascade were set to translational periodicity boundary (TPB). Results show that VGJ significantly eliminates the turbulent separation and improves the performance of cascade. The overall loss coefficient of cascade is reduced by 52.3% at most. The optimal VGJ location in this study is not separation location but 7% axial chord downstream of it, which is probably due to the ultra-high loading of the cascade. The spanwise location and intensity of jet vortices are mainly affected by skewed angle, and will increase with the skewed angle. The VGJ case with a higher pitch angle features better control effect on suction surface but increases the pitchwise scale of the low-velocity region. In addition, as pitch angle increases, the mixing loss between mainstream and jet flow increases, which brings a negative effect on cascade performance. In order to enhance the momentum transporting between mainstream flow and separation flow, counter vortex-generator jet (CVGJ) schemes were designed and studied. Though CVGJ strengthens the jet vortices near the trailing edge and achieves minimum loss near TPB sections at 150% axial chord section, the interaction of jet vortices with opposite rotating direction brings extra loss, which leads to the control effect of CVGJ is weaker than that of VGJ.

Active Flow Control in Compressor Cascades with Steady and Pulsed Jets

This paper describes a detailed experimental investigation into the impact of steady and pulsed blowing on endwall secondary structures and losses in a compressor cascade. Owing to their high configuration flexibility, injection holes are integrated in the cascade sidewalls to manage the secondary flows. Loss reductions of 3.2 and 5.72% relative to the uncontrolled case are achieved by steady blowing with straight and optimized inclined holes, respectively. Superior loss reduction of 7.85% is obtained by pulsed blowing through inclined holes. To identify the secondary flow structures near the endwall and suction surfaces, a self-developed oil visualization method and spatial-spectral analysis are performed. Experimental results show that two concentrated shedding vortices exist in the cascade corner region. Loss reduction is achieved as the blowing suppresses the dominant vortex. Pulsed blowing intensifies the acceleration effect on the boundary layer, resulting in better performance with the same injection velocity. The impact of the pulse frequency on loss generation is investigated, and it is found that the optimal frequency is close to the shedding frequency of the dominant vortex in the cascade corner region.

Influence of Compound Spherical Trenched Holes on Film Cooling Performance at the end of Combustor Simulator

The major effects of cylindrical and spherical trenched cooling holes with distance between the hole surface and the combustion chamber panel and the filler diameter on the spherical hole contact surface and the panel surface H=0.3,R=D/2=0.1, H=0.3, R=D/2=0.2 and H=0.3, R=D/2=0.3 cm at BR=3.18 on the film cooling effectiveness near the combustor end wall surface is an important subject to study in details. In this research, a three-dimensional representation of a Pratt and Whitney gas turbine engine was simulated and analysed with a commercial finite volume package FLUENT 6.2.26. The analyses were done with RANS turbulence model on internal cooling passages. The combustor simulator was combined with the interaction of two rows of dilution jets, which were staggered in the streamwise direction and aligned in the spanwise direction. In comparison with the baseline case, the application of trenched holes increased the effectiveness of film cooling up to 47% near the wall surface and an average of 35% in depth of combustor simulator .