Numerical Investigation of the Influence of Real World Blade Profile Variations on the Aerodynamic Performance of Transonic Nozzle Guide Vanes

2010 ◽  
Author(s):  
R. Edwards ◽  
A. Asghar ◽  
R. Woodason ◽  
M. LaViolette ◽  
K. Goni Boulama ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Trailing Edges ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Transonic Cascade

This paper addresses the issue of aerodynamic consequences of small variations in airfoil profile. A numerical comparison of flow field and cascade pressure losses for two representative repaired profiles and a reference new vane were made. Coordinates for the three airfoil profiles were obtained from the nozzle guide vanes of refurbished turboshaft engines using 3D optical scanning and digital modeling. The repaired profiles showed differences in geometry in comparison with the new vane, particularly near the leading and trailing edges. A numerical simulation was conducted using a commercial CFD code which uses the finite element approach for solving the governing equations. The computational predictions of the aerodynamic performance were validated with experimental results obtained from a transonic cascade consisting of blades with the same airfoil profiles. A CFD analysis was performed for the cascade at subsonic inlet and transonic exit conditions. Boundary layer growth, wake formation, and shock boundary layer interactions were observed in the two-dimensional computations. The flow field showed the presence of shock waves downstream of the passage throat and near the trailing edges of the blades. A conspicuous change in flow pattern due to subtle variation in airfoil profile was observed. The calculated flow field was compared with the flow pattern visualized in the experimental test rig using the Schlieren method. The total pressure calculation for the cascade exit showed an increase in pressure loss for one of the off-design profiles. The pressure loss calculations were also compared with the multi-hole total pressure probe measurement in the transonic cascade rig.

Numerical Investigation of the Influence of Real World Blade Profile Variations on the Aerodynamic Performance of Transonic Nozzle Guide Vanes

2011 ◽  
Vol 134 (2) ◽  
Author(s):  
R. Edwards ◽  
A. Asghar ◽  
R. Woodason ◽  
M. LaViolette ◽  
K. Goni Boulama ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Guide Vanes ◽  
Trailing Edges ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

This paper addresses the issue of aerodynamic consequences of small variations in airfoil profile. A numerical comparison of flow field and cascade pressure losses for two representative repaired profiles and a reference new vane were made. Coordinates for the three airfoil profiles were obtained from the nozzle guide vanes of refurbished turboshaft engines using 3D optical scanning and digital modeling. The repaired profiles showed differences in geometry in comparison with the new vane, particularly near the leading and trailing edges. A numerical simulation was conducted using a commercial CFD code, which uses the finite volume approach for solving the governing equations. The computational predictions of the aerodynamic performance were compared with experimental results obtained from a cascade consisting of blades with the same airfoil profiles. The CFD analysis was performed for the cascade at subsonic inlet and transonic exit conditions. Boundary layer growth, wake formation, and shock boundary layer interactions were observed in the two-dimensional computations. The flow field showed the presence of shock waves downstream of the passage throat and near the trailing edges of the blades. A conspicuous change in flow pattern due to subtle variation in airfoil profile was observed. The calculated flow field was compared with the flow pattern visualized in the experimental test rig using the schlieren method. The total pressure calculation for the cascade exit showed an increase in pressure loss for one of the off-design profiles. The pressure loss calculations were also compared with the multihole total pressure probe measurement in the transonic cascade rig.

Impact of a Combination of Micro-Vortex Generator and Boundary Layer Suction on Performance in a High-Load Compressor Cascade

2018 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Chuanle Liu
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Control Methods ◽  
Compressor Cascade ◽  
Boundary Layer Suction ◽  
Micro Vortex Generator

The performance of a compressor cascade is considerably influenced by flow control methods. In this paper, the synergistic effects of combination between micro-vortex generators (MVG) and boundary layer suction (BLS) are discussed in a high-load compressor cascade. Seven cases, which are grouped by a kind of micro-vortex generator and boundary layer suction with three locations, are investigated to control secondary flow effects and enhance the aerodynamic performance of the compressor cascade. The MVG is mounted on the end-wall in front of the passage. The rectangle suction slot with three radial positions is installed on the blade suction surface near the trailing edge. The numerical results show that: at the design condition, the total pressure loss is effectively decreased as well as the static pressure coefficient increase when the combined MVG and SBL method (COM) is used, which is superior to MVG in an aerodynamic performance. At the stall condition, the induced vortex coming from MVG could mix the low-energy fluid and mainstream, which result in the reduced separation, and the total pressure loss decreased by 11.54% when the suction flow ratio is 1.5%. The total pressure loss decreases by 14.59% when the COM control methods are applied.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

scholarly journals The Effect of Inlet Boundary Layer Thickness on the Flow Within an Annular S-Shaped Duct

1998 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Mineyasu Oana
Keyword(s):  
Boundary Layer ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Vortex Pair ◽  
Separated Flow ◽  
Total Pressure Loss ◽  
Centrifugal Impeller

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the same geometry as that used to connect compressor spools on our experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with a similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e. thin and thick IBL) were used. Results showed that large differences of flow pattern were observed at the S-Shaped duct exit between two types of the IBL, though the value of “net” total pressure loss has not been remarkably changed. According to “overall” total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor as a large inlet distortion, and strongly affect the downstream compressor performance. There is a much-distorted three-dimensional flow pattern at the exit of S-Shaped duct. This means that the aerodynamic sensitivity of S-Shaped duct to the IBL thickness is very high. Therefore, sufficient carefulness is needed to design not only downstream aerodynamic component (for example centrifugal impeller) but also upstream aerodynamic component (LPC OGV).

Blowing Control in Serpentine Inlet and its Effect on Fan-Stage Performance

2020 ◽  
Author(s):  
Yubo He ◽  
Qingzhen Yang ◽  
Huicheng Yang ◽  
Saile Zhang ◽  
Haoqi Yang
Keyword(s):  
Flow Field ◽  
Mass Flow Rate ◽  
Flow Rate ◽  
Pressure Loss ◽  
Total Pressure ◽  
High Energy ◽  
Aerodynamic Performance ◽  
Blowing Ratio ◽  
Total Pressure Recovery

Abstract Serpentine inlet is widely used in military and civil aircraft due to its good stealth performance. However, it generates a high total pressure loss and swirl distortion which significantly affects the performance and the stability of the compressor. In order to improve the quality of the flow field at the aerodynamic interface plane (AIP), a flow control is required inside the serpentine inlet. The objective of this paper was to study the effectiveness of the blowing active flow control on reducing the swirl distortion and on improving the total pressure recovery at the AIP, by reducing the low-momentum flow in the serpentine inlet. The mechanism of the blowing control and the effect of the design parameters (i.e. blowing angle, blowing position and blowing flow rate) on the aerodynamic performance at the AIP were studied. The optimal solution was applied to the full flow path of the serpentine inlet and the fan-stage. The numerical results showed that the quality of the flow field at the AIP were effectively improved by blowing high-energy airflow into the boundary layer of the serpentine inlet. The blowing position had a high influence on the blowing effect, and upper wall blowing scheme obtained greater benefits than lower wall blowing scheme and combination blowing scheme. In addition, the blowing angle should be selected to avoid the high-energy air from pipes mixing with mainstream in the serpentine inlet which will result in an additional total pressure loss. When the ratio of the blowing mass flow rate to the designed mass flow rate of the serpentine inlet was about 1.5%, the swirl distortion on the AIP reached a minimum value, which then did not show a significant difference in performance with blowing ratio increased. When the upper wall blowing scheme was adopted with a blowing angle of 6 degrees and a blowing ratio of 1.5%, the AIP aerodynamic performance achieved the highest improvement, with an increase of the total pressure recovery factor by about 1%, and a decrease of the circumferential total pressure distortion and the swirling distortion by 60% and 61%, respectively. With the optimal control scheme, the area of the low-pressure region near the upper wall was remarkably reduced, and the performance of fan-stage was improved, with an increase of the pressure ratio by about 1.5%, and the efficiency of the single-stage compressor by about 3.1%, respectively.

Numerical Investigation of Dump Diffuser Combustor Performance at Uniform and Non-Uniform Inlet Conditions

2020 ◽  
Author(s):  
Heyu Wang ◽  
Kai Hong Luo
Keyword(s):  
Boundary Layer ◽  
Numerical Investigation ◽  
Layer Thickness ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Thickness ◽  
Static Pressure ◽  
Aerodynamic Performance ◽  
Aero Engine ◽  
Inlet Conditions

Abstract A numerical investigation has been conducted for an axisymmetric dump diffuser combustor, which is a simplified geometry of a typical lean-burn combustor in a modern civil aero-engine gas turbine. The aerodynamic performance of the combustor is analyzed with an emphasis on two common performance parameters: static pressure recovery and total pressure loss. The former is essential in maintaining high-pressure air flow across the liner, whereas the latter involves the specific fuel consumption of the aero-engine. At first, the effects of geometrical parameters of the dump diffuser combustor are investigated. A high diffuser angle seems to be detrimental to both static pressure recovery and total pressure loss. On the other hand, a high dump gap ratio is beneficial from the aerodynamic performance point of view. However, all these desired characteristics are subject to mechanical constraints and their implications for specific consumption. Optimum values of those parameters should exist for a given desired aerodynamics performance. The majority of previous researches, including the first part of this study, have been carried out with uniform inlet conditions due to a typical independent design cycle of each component. The effects of compressor exit conditions are usually not considered in the early stage design process. In the second part of this study, various inlet conditions representing a more realistic compressor exit condition such as inlet symmetrical and asymmetrical boundary layer thickness are investigated. The performance of an asymmetrical configuration with a thin boundary layer thickness near the outer annulus is almost comparable to that of its uniform counterpart. Findings of this study provide useful input for combustor designers to improve the combustor’s performance based on the compressor exit conditions.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2018 ◽  
Vol 140 (7) ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high-pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake's path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue three-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady stator 2 performance, time-resolved interrogations of the stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between rotor 1 and rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the stator 1 wake and rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, stator 2 experiences significantly different inlet blockage and unsteadiness from the rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of stator 2.

The Effect of Inlet Boundary Layer Thickness on the Flow Within an Annular S-Shaped Duct

1999 ◽  
Vol 121 (3) ◽  
pp. 626-634 ◽  
Author(s):  
T. Sonoda ◽  
T. Arima ◽  
M. Oana
Keyword(s):  
Boundary Layer ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Vortex Pair ◽  
Separated Flow ◽  
Total Pressure Loss ◽  
Centrifugal Impeller

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the same geometry as that used to connect compressor spools on our experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with a similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e., thin and thick IBL) were used. Results showed that large differences of flow pattern were observed at the S-shaped duct exit between two types of the IBL, though the value of “net” total pressure loss has not been remarkably changed. According to “overall” total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor as a large inlet distortion, and strongly affect the downstream compressor performance. There is a much-distorted three-dimensional flow pattern at the exit of S-shaped duct. This means that the aerodynamic sensitivity of S-shaped duct to the IBL thickness is very high. Therefore, sufficient care is needed to design not only downstream aerodynamic components (for example, centrifugal impeller) but also upstream aerodynamic components (LPC OGV).

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

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