Total Pressure Loss Coefficient Research Articles

Aerodynamic performance of profiled endwalls with upstream slot purge flow in a linear turbine cascade having pressure side separation

In aeroengines, purge flow directly fed from the compressor (which bypasses the combustor) is introduced through the disk space between blade rows to prevent the hot ingress. Higher quantity of purge gas fed through the wheel space can provide additional thermal protection to the passage endwall and blade surfaces. However, the interaction of purge flow with the mainstream flow leads to higher secondary losses. Secondary losses inside a turbine blade passage can be reduced effectively by endwall contouring. This paper presents computational investigation on the influence of non-axisymmetric endwall contouring over endwall secondary flow modification in the presence of purge flow with the pressure side bubble (PSB). The experimental analysis was conducted for the base case without purge and base case with purge (BCP) configurations having flat endwalls. The total pressure loss coefficient and exit yaw angle deviation were measured with the help of a five-hole pressure probe. Static pressure distribution over the blade midspan was obtained by 16 channel Scanivalve. Aerodynamic performances of three different profiled endwalls are numerically analyzed and are compared against the BCP configuration. The effects of different contoured endwall geometries on endwall static pressure distribution and secondary kinetic energy were also discussed. Analysis shows that in the first contoured endwall configuration (EC1), the formation of stagnation zones at a contour valley close to the suction surface causes the exit total pressure loss coefficient to increase. The shifting of the contour valley near to the pressure surface (EC2 configuration) has resulted in local acceleration of the diverted pressure side leg of the horseshoe vortex over the hump toward the end of the passage. In the third configuration (EC3 configuration), reduced valley depth and optimum hump height have effectively redistributed the endwall pitchwise pressure gradient. The increased static pressure coefficient at the endwall near to the pressure surface has eliminated the PSB formation. In addition, computational results of unsteady Reynolds averaged Navier–Stokes simulations are obtained for analyzing transient behavior of PSB, with more emphasis on its migration on the pressure surface and transport across the blade passage. The additional work done by the mainstream fluid to transport the low momentum PSB fluid has caused higher aerodynamic penalty at the blade exit region. In this viewpoint, the implementation of contoured endwalls has shown beneficial effects by eliminating the PSB and related secondary vortices. At 27% of axial chord downstream of the blade trailing edge, a 4.1% reduction in the total pressure loss coefficient was achieved with endwall contouring.

Multi-Objective Optimization of Cascade Blade Profile Based on Reinforcement Learning

The multi-objective optimization of compressor cascade rotor blade is important for aero engine design. Many conventional approaches are thus proposed; however, they lack a methodology for utilizing existing design data/experiences to guide actual design. Therefore, the conventional methods require and consume large computational resources due to their need for large numbers of stochastic cases for determining optimization direction in the design space of problem. This paper proposed a Reinforcement Learning method as a new approach for compressor blade multi-objective optimization. By using Deep Deterministic Policy Gradient (DDPG), the approach modifies the blade profile as an intelligent designer according to the design policy: it learns the design experience of cascade blade as accumulated knowledge from interaction with the computation-based environment; the design policy can thus be updated. The accumulated computational data is therefore transformed into design experience and policies, which are directly applied to the cascade optimization, and the good-performance profiles can be thus approached. In a case study provided in this paper, the proposed approach is applied on a blade profile, which is thus optimized in terms of total pressure loss and laminar flow area. Compared with the initial profile, the total pressure loss coefficient is reduced by 3.59%, and the relative laminar flow area at the suction surface is improved by 25.4%.

Aerothermal and aerodynamic performance of turbine blade squealer tip under the influence of guide vane passing wake

In this paper, the performance of the turbine blade squealer tip has been studied detailed aimed to highlight the impact of the upstream guide vane passing wake. The first stage of GE-E3 high-pressure turbine has been employed to perform the three-dimensional simulation and the computational domain has been scaled based on the domain scaling method. Boundary conditions are consistent with operating conditions of the annular cascade testing. Circumferential averaged and realistic non-uniform interface conditions have been used to obtain steady and unsteady characteristics respectively. The validation of the turbulent model and mesh independent test has been conducted detailed in previous work. Three squealer tips, including two widths and heights, have been designed and investigated to understand its influence. Results show that the aerothermal performance of the squealer tip is remarkably influenced by the upstream passing wake. Although steady and time-averaged results have a good agreement, the variation of instantaneous heat transfer coefficient (HTC) would be over 30%, especially on the Cavity Floor region. Changing the geometry of the squealer also has different impacts on both steady and unsteady performance. The unsteady aerodynamic has relatively small fluctuation within 10%, and the distribution of steady and time-averaged leakage flow as well as total pressure loss coefficient still have a satisfactory agreement.

Blowing Ratio Effect on Film Cooling Performance for a Row Holes Installed in Different Trench Configurations

In the present study, a numerical analysis was performed to evaluate the performance of cooling hole embedded in different trenched designs (triangular trench, semi-cylindrical trench and corrugated trench) in improving the film cooling efficiency over a flat plate. These concepts are compared to the rectangular trenched and the traditional cylindrical hole. The commercial software ANSYS CFX 18 was used to conduct a series of required numerical calculations. The centerline and laterally averaged film cooling effectiveness and total pressure loss coefficient for the five cases are analyzed at three blowing ratios, M=0.5, M=1.0 and M=1.5. Results show a uniform coverage is obtained by hole installed in trench. The main result obtained in this paper that the cooling hole with corrugated trench enhance the film cooling effectiveness with less total pressure loss. The main result of this study reveals that the jet installed in the trenches yield a better film cooling effectiveness especially at higher blowing ratios (M≥1).

Numerical investigation and loss estimation of high-pressure turbine cascade flow with contoured endwall and incoming wakes

With the aim of suppressing the irreversible losses, the secondary flow at endwall and incoming vortices interactions are assumed to be the most concerned physics in gas turbine for modern aero-engine and military vehicles. However, the combining effect of contoured endwall and incoming periodic wakes were seldom discussed previously. In the present work, the research subject is the high-pressure turbine cascade with non-axisymmetric endwall and incoming wakes induced by moving bar fulfilled with periodic boundary conditions in pitchwise direction. Firstly, the aerodynamics and thermodynamics in the turbine cascade with secondary flow control technique, which denotes to non-axisymmetric endwall, were experimentally and numerically investigated. For what concerns the profile aerodynamic loadings, total pressure loss coefficients and vortices coefficient were surveyed to quantify the influence of interaction between incoming wakes and secondary flow at contoured endwall. The cumulative difference of entropy production rate associated with dissipation is then employed to evaluate the loss under different inflow condition. Combined with dynamic mode decomposition (DMD) method, the energy loss features associated with incoming wakes at different passing frequencies was captured. The results indicated that the flow separation at the suction wall with the axial position of 70%∼80% Cax upstream trailing edge was suppressed in the cases with incoming wakes for high pressure turbine cascade. And stronger incoming wakes tends show better suppression effect on the flow separation. However, when combined with contoured endwall, the negative influence of non-axisymmetric endwall at the middle passage was enlarged. And the general aerodynamics of the cascade flow may not be necessarily improved in term of energy loss characteristics with the increase of passing frequency of incoming wakes. The present work preserves guiding significance for the design and flow control of high pressure turbine cascade.

Pressure drop analysis and aerodynamic design of compressor L-inlet duct

Pressure drop inside the compressor L-inlet duct (CLID) has a great impact on the intake loss, which directly affects the operating cost of the power plant. Generally, every 500 Pa inlet pressure drop is equivalent to 0.5% intake loss at standard conditions, which can contribute to higher than 1% of gas turbine power loss in terms of fuel flow. Hence, the pressure drop should be kept as low as possible. Pressure drop that occurs within the CLID mainly depends on the structural parameters, however, no general design guidelines were found in open literature. Moreover, a reasonable intake duct aerodynamic design is the fundamental solution to the problem of compressor inlet distortion. Therefore, in this paper, the fluid dynamics of CLID are revealed by using both the experimental and numerical methods. Then, the effects of typical parameters on the pressure drop and total pressure loss coefficient are investigated. Furthermore, the conceptual design of CLID is proposed on the basis of distortion flow analysis and 1D intake volute design. The results show that larger CLID width (a/D4) and depth (b/D4) result in larger flow separation at the contraction, thereby leading to smaller total pressure drop and total pressure loss coefficient. However, the opposite change occurs at the cone diameter (D0/D4). Besides, the conceptual design of CLID is proposed by indirectly establishing the relationship between the outlet flow angle and inlet structural parameters.

The effect of rectangulargurney flap on wells turbine performance

Wells turbine is very adept at converting pneumatic power from ocean waves into mechanical energy. However, such turbines incur from low performance, limited operating range and low efficiency. The present study introduces a method to enhance the power produced from Wells turbine by implementing rectangular Gurney flap (GF). The Gurney flap is placed on both pressure and suction sides of the trailing edge (TE) and perpendicular to the chord line without change chord length. GF increases the lift coefficient by altering the Kutta condition at the TE. The turbine performance is herein evaluated by solving numerically 3D incompressible Reynolds Averaged Navier–Stokes equations (RANS) by using ANSYS Fluent 19.0. The performance is demonstrated according to the flow coefficient, torque coefficient, total pressure loss coefficient, and the turbine efficiency, as well as the velocity and pressure fields around the turbine blades. The validation of the present work was achieved using previous experimental work and CFD work by using SST k – ω turbulence model. The present work showed that the Wells turbine augmented with GF increases the torque coefficient by 41.98%. Besides, the stall has been delayed compared with the conventional Wells turbine.

Performance improvement of low-pressure cylinder in high back pressure steam turbine for direct heating

It is the general retrofit to remove the last stages in the low-pressure cylinder of steam turbine for high back pressure direct heating, which can avoid the exergy loss caused by the traditional extracting steam heating mode of combined heat and power. However, the flow field of low-pressure cylinder will change. In this study, the last stage flow field of low pressure cylinder of such high back pressure turbine, as well as the temperature distribution, is analyzed and optimized by setting the diversion ring. Low-pressure cylinder of the 300 MW steam turbine is selected as the research object. The optimal scheme is obtained after analyzing vortex distribution, static pressure recover coefficient and total pressure loss coefficient in flow field under different heating conditions. The results show that adding appropriate diversion ring bringing about vortex effectively reduced, leading to a sharp increase of 7.84% in efficiency of low-pressure cylinder and 0.63in statistic pressure recover efficient. It is also contributing to an apparently decrease of 0.95 in total pressure loss efficient. All results indicate that the aerodynamic performance of low-pressure cylinder has obvious promotion. Under high back pressure heating condition, the structure of low-pressure cylinder of steam turbine changed.

Loss Mechanism of Low-Pressure Turbine Secondary Flows Due to Different Incoming Boundary Layers

Abstract In high bypass ratio engines, the flow exits the interturbine duct (ITD) and enters the low-pressure (LP) turbine. This paper aims to understand the effects of the boundary layer at the exit of ITD on the endwall secondary flows and loss of the first blade row in a low-pressure turbine. From the Navier–Stokes equations, the loss is decomposed into the parts generated by the mean vortex as well as turbulence theoretically. The result of computational fluid dynamics (CFD) shows that the incoming boundary layer from the ITD increases the total pressure loss coefficient by 14% compared to the case with uniform inlet condition. Although the distribution of the secondary vortices is strongly affected by the inlet boundary layer, the loss generated by the mean vortex within the blade passage is hardly affected. The analysis based on the turbulent dissipation shows that the dominant factor leading to the loss increase is the turbulent dissipation downstream of the blade trailing edge (TE) near the hub. The mixing process of the wake and the strong counter-rotating vortex pair (CVP) increases the turbulent dissipation significantly. It is also found that a simplified incoming boundary layer defined by the Prandtl's one-seventh power law can not reproduce the complex effects of the incoming boundary layer from the ITD.

Influence of leading edge with real manufacturing error on aerodynamic performance of high subsonic compressor cascades

To investigate the influence of real leading-edge manufacturing error on aerodynamic performance of high subsonic compressor blades, a family of leading-edge manufacturing error data were obtained from measured compressor cascades. Considering the limited samples, the leading-edge angle and leading-edge radius distribution forms were evaluated by Shapiro-Wilk test and quantile–quantile plot. Their statistical characteristics provided can be introduced to later related researches. The parameterization design method B-spline and Bezier are adopted to create geometry models with manufacturing error based on leading-edge angle and leading-edge radius. The influence of real manufacturing error is quantified and analyzed by self-developed non-intrusive polynomial chaos and Sobol’ indices. The mechanism of leading-edge manufacturing error on aerodynamic performance is discussed. The results show that the total pressure loss coefficient is sensitive to the leading-edge manufacturing error compared with the static pressure ratio, especially at high incidence. Specifically, manufacturing error of the leading edge will influence the local flow acceleration and subsequently cause fluctuation of the downstream flow. The aerodynamic performance is sensitive to the manufacturing error of leading-edge radius at the design and negative incidences, while it is sensitive to the manufacturing error of leading-edge angle under the operation conditions with high incidences.

An aerothermal analysis of the effects of tip gap height and cavity depth of a gas turbine blade

To investigate the effects of tip gap height and the squealer cavity depth on aerothermal performance of a gas turbine rotor, two tip gap height and eight squealer cavity depth are investigated under stage environment. For the aerodynamic performance of blade tips, when the tip gap height is increased from 1.0% span to 1.5% span, the tip leakage flow rate is increased by about 1.0% total mass flow and the total pressure loss coefficient is increased by about 0.03. With the increase of cavity depth, the tip leakage flow rate is reduced but the vortexes loss in the cavity and passage vortex is enlarged, so the total pressure loss coefficient is reduced at first and then is increased. For the film cooling effectiveness of blade tips, the average film cooling effectiveness of tip is improved at a large tip gap height and a small cavity depth. With the increase of cavity depth, the area needed cooling protection is increased; the relative location between the film hole and the separation line is changed, resulting in the change of film area of the squealer tip. Thus, the film cooling effectiveness of the squealer tip is reduced with fluctuation when the cavity depth is increased. In all, the influence of the tip gap height and cavity depth on tip leakage flow rate, total pressure loss coefficient and film cooling effectiveness is conflicting. Determining an optimal cavity depth in which more film holes located on the separation line for a fixed tip gap height is important to obtain better aerodynamic performance and film cooling performance.

Effects of casing angle on the performance of parallel hub axial annular diffuser

AbstractIn this paper, the effects of non-swirling and swirling flow on the performance of parallel hub axial annular diffuser has been investigated. The study was conducted on a fully developed swirling flow and non-swirling flow to predict the separation of the flow from the wall. Three different annular diffusers were used with casing wall angles of 3°, 6°, and 9°. Furthermore, various swirl angles (0–25°) at the inlet of diffusers have been investigated to analyze the performance across the length. It was found that parallel hub axial annular diffuser performance increases up to a certain length as the inlet swirl angle increases. However, the performance also improves as the diffuser area ratio (AR) increases. The performance is evaluated based on the static pressure recovery coefficient (Cp) and the total pressure loss coefficient (CTL). The highest possible pressure recovery is achieved by the 12° swirl angle with a casing angle of 6°.

Effects of casing angle on the performance of parallel hub axial annular diffuser

Abstract In this paper, the effects of non-swirling and swirling flow on the performance of parallel hub axial annular diffuser has been investigated. The study was conducted on a fully developed swirling flow and non-swirling flow to predict the separation of the flow from the wall. Three different annular diffusers were used with casing wall angles of 3°, 6°, and 9°. Furthermore, various swirl angles (0–25°) at the inlet of diffusers have been investigated to analyze the performance across the length. It was found that parallel hub axial annular diffuser performance increases up to a certain length as the inlet swirl angle increases. However, the performance also improves as the diffuser area ratio (AR) increases. The performance is evaluated based on the static pressure recovery coefficient (Cp) and the total pressure loss coefficient (CTL). The highest possible pressure recovery is achieved by the 12° swirl angle with a casing angle of 6°.

The genetic algorithm-radial basis function neural network to quickly predict aerodynamic performance of compressors

In this paper, compressor aerodynamic performance has been predicted based on throughflow theory, combined with a surrogate model, which is a combination of the Genetic Algorithm (GA) and generalized Radial Basis function (RBF) neural network. And the predicting results have been compared with those from the traditional models and spanwise mixing model, which still widely be used to predict the aerodynamic performance. We first predicted the deviation angle and total-pressure loss coefficient (TPLC) by the surrogate model, and then using these two intermediate variables connected the model with throughflow theory. The pressure ratio and efficiency, representing the compressors’ total performance parameters, are predicted and compared with experimental data. In order to increase the accuracy of prediction, a data augmentation method based on the piecewise cubic Hermite interpolation (PCHIP) algorithm is introduced to enlarge the training database. At the same time, considering the vast differences of deviation angle and loss in different working conditions as well as aerodynamic and geometric differences of rotor and stator, the database and the network should be split into six components based on the choke, the normal and the stall conditions as well as rotor and stator. Then, the performance curves of pressure ratio and efficiency can be determined by an iteration process. The predicting results are compared with experimental data, which shows that the surrogate model matches experiments much better than those from the traditional models and spanwise mixing model.

Effect of Change in the Curvature of Flow and Performance Characteristics of Straight and S-shaped Diffuser

The effect of the change in the curvature of the stream and the execution attributes of straight walled and S-shaped rectangular diffusers having a center-length of 500 mm and an area ratio of 1.5 is studied in this paper. Straight and S-shaped diffusers with a square inlet and a rectangular outlet have been fabricated using an alternate turn angle. The cross-flow velocity, static pressure and total pressure variation along the length, at different sections of the diffusers, are measured using pressure tapings and normal three-hole probe. The low performance is due to the generation of secondary flows as a result of the geometrical curvature and the additional losses are due to the surface roughness of the diffusers. The total pressure loss coefficient for all the diffusers is nearly identical and seems to be independent of the angle of turn. In this paper, results were shown in the form of pressure recovery coefficient, coefficient of total pressure loss, average cross-flow velocity and flow deviation. Furthermore, pressure and velocity contours were developed for the entire study of all four designs of the diffuser.

Numerical study on the smoke extraction efficiency and the improvement through a smoke reservoir in the naturally ventilated tunnel with vertical shaft

Plug-holing of the smoke layer is an unfavorable phenomenon for the smoke extraction efficiency in the naturally ventilated tunnel with vertical shaft. In this work, theoretical analysis and large eddy simulation (LES) are carried out to investigate how to prevent the occurrence of the plug-holing phenomenon and improve smoke extraction efficiency. The results show that the occurrence of plug-holing is influenced by the height, cross-sectional area and total pressure loss coefficient of the vertical shaft, as well as the thickness of the smoke layer. A modified dimensionless number Fr* is proposed to represent the proportion of smoke in the exhausted mass flow. Fr* varies as the −1/4 power of the cross-sectional area of the shaft, −1/2 power of the height of the shaft, and increases linearly with the thickness of the smoke layer. The critical value of Fr* for predicting whether the plug-holing phenomenon occurs is determined to be 0.34 by combining the results of the previous small-scale experiment. In addition, a smoke reservoir is designed under the shaft to increase the thickness of the smoke layer, which is proven to be effective in preventing the occurrence of the plug-holing phenomenon and improving the smoke extraction efficiency of the vertical shaft. The findings of this paper contribute to the understandings of smoke extraction for naturally ventilated tunnel with vertical shaft, and the results are also useful to the fire protection engineering in tunnels.

Pressure loss reduction in ventilation ducts by shape optimization of the removable profiled components

Reduction of pressure loss in ventilation systems can be achieved by changing the shape of the airflow region and smoothing the boundary surfaces. The installation of special profiling components inside a straight shaped part of ventilation duct is the simplest way to reduce the coefficient of local resistance and pressure loss. For the best result under given conditions, the shape optimization problem must be formulated and solved. In this paper, we use the open source code software and numerical methods to study the optimal-shaped drag reduction components in elbow ventilation ducts. Computational Fluid Dynamics is used to predict the flow fields and the drag reduction effect. After the validation of flow model against existing experimental results, a detailed study has been conducted to shape optimization of removable profiled components. The optimization problem is parameterized by a small number of structural variables, which are the coordinates of the control points of the splines that form the contours of the flow region. Total pressure loss coefficient is selected as the objective function for optimization. A genetic algorithm is used as an optimization method. The results reveal that the removable profiled drag reduction component can reduce the resistance in ducts by 60%-80%.

Design of a Sector Cascade Applied in the Middle Stage of a Compressor Test Rig

A compressor test platform was designed in the purpose of easy assembly and cost-saving tests. New design concepts were firstly used on the test platform and iterative revisions were conducted to verify the effects, which decrease the risk of applying to the full annulus directly. The platform was a sector cascade that can be inserted into an otherwise full stator annulus, with a set of exchangeable endwall casing and blades manufactured by 3-D printing. The platform can create an operating condition which is closer to engine-realistic flow conditions than traditional cascade tests. The flow field of the prototype stator was tested in detail at the operating point and showed some flow defects in the tip region, and then three design plans were conducted to experimentally investigate which design concept could improve the blade tip flow. The concept of casing endwall profiling was the most effective one which markedly depressed the separation of the tip region and decreased the total pressure loss coefficient. The test results show that the design of the test platform was successful and promising, which could be used to conduct more researches on the flow mechanism of the middle stage.

Effects of blade surface roughness on compressor performance and tonal noise emission in a marine diesel engine turbocharger

A numerical study was conducted to investigate the effects of blade surface roughness on compressor performance and tonal noise emission. The equivalent sand-grain roughness model was used to account for blade surface roughness, and a hybrid method that combines computational fluid dynamics and boundary element method was used to predict compressor performance and tonal noise. The numerical approach was validated against experimental data for a baseline compressor. Nine different cases with different blade surface roughness were studied in this paper, the global performance was analyzed under compressor design speed, and the tonal noise level was predicted under the design condition. The results indicate that compressor total-to-total pressure ratio and isentropic efficiency were gradually decreased with the increasing blade surface roughness. Besides, the blade total pressure loss coefficient and the efficiency loss coefficient were also increased. It was found that the reverse flow at the leading edge of compressor rotor blades reduced blade loading. The pressure fluctuation at the leading edge showed that the peak of pressure fluctuations increased as the blade surface roughness was increased. The sound pressure level at blade-passing frequency shows a significant change with variation in blade surface roughness, which results in an increased total noise level. Furthermore, it was shown that the blade surface roughness had nearly no influence on acoustic directivity, but the sound pressure level increased with the increase in roughness, especially at blade-passing frequency.

VIGV Angle Optimization for Subsonic Axial Flow Compressor

Long-term performance of an axial-flow compressor by the scheduling of inlet guide vanes at off-design conditions has been studied in this paper. The compressors are used in various industries and in aviation sectors for different operations at different climatic conditions; due to diverse climatic conditions the compressor is unable to give good performances as expected. At the design conditions, the results show that the variations in total pressure loss coefficient, volume flow rate, pressure ratio, and others like thermodynamics, aerodynamic properties are different at different stagger angles of 15°, 30° and 45°. BladeGen tools were used to design the inlet guide vanes for the investigations. The performance parameters of the axial flow compressor were analyzed by using ANSYS CFX and were validated using the analytical method. The objective of this study is to minimize the total pressure loss coefficient and improve the aerodynamic characteristics of an inlet guide vanes at different stagger angles and hence reduce the overall fuel consumption. The outcomes of this work give an improved insight into the efficient use of a VIGV in axial flow compressor.