A New Concept of a Two-Dimensional Supersonic Relative Inlet Mach Number Compressor Cascade

2009 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Markus Olhofer ◽  
Toshiyuki Arima ◽  
Bernhard Sendhoff
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Optimization Method ◽  
Superior Performance ◽  
Compressor Cascade ◽  
Pressure Loss Coefficient ◽  
Supersonic Inlet ◽  
Flow Mechanisms ◽  
Improved Performance ◽  
Total Pressure Loss Coefficient

In this study, a numerical shape optimization method based on evolutionary algorithms coupled with a verified CFD solver has been applied to the ambitious target of a shock free 2-D supersonic inlet Mach number compressor cascade. The study is based on the DLR-PAV-1.5 supersonic compressor cascade designed by the pre-compression blading concept. The DLR cascade airfoil has been optimized using a verified CFD code. A superior performance of the optimized supersonic cascade with about 24% reduction of the total pressure loss coefficient compared to the original cascade has been realized. The flow mechanisms observable around the blade with improved performance and the resulting design concept are discussed in this paper.

scholarly journals Effects of Surface Roughness, Freestream Turbulence, and Incidence Angle on the Performance of a 2-D Compressor Cascade

1990 ◽  
Author(s):  
W. C. Elrod ◽  
P. I. King ◽  
E. M. Poniatowski
Keyword(s):  
Surface Roughness ◽  
Total Pressure ◽  
Incidence Angle ◽  
Two Dimensional ◽  
Freestream Turbulence ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Linear Cascade ◽  
Pressure Loss Coefficient ◽  
Total Pressure Loss Coefficient

The effects of surface roughness, freestream turbulence, and incidence angle on the performance of a two-dimensional compressor cascade were investigated. The test section consisted of seven NACA 65-A506 airfoils arranged in a linear cascade. Four different surface roughness conditions were applied to the first 25 percent chord on the suction surface of each of the five middle blades in the cascade. Freestream turbulence levels of approximately one and seven percent were used. Incidence angles of −3, zero and +3 degrees were investigated. Of the three parameters tested, freestream turbulence exerted the largest influence on blade performance. The total pressure loss coefficient increased with increased roughness and was reduced for large turbulence. Changes in flow incidence had a lesser effect on the performance of the blade.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Some Experimental Results of Two-Dimensional Compressor Cascades at Supersonic Inlet Velocities

1970 ◽  
Vol 92 (3) ◽  
pp. 267-274 ◽  
Author(s):  
H. Starken ◽  
H. J. Lichtfuss
Keyword(s):  
Mach Number ◽  
Density Ratio ◽  
Great Influence ◽  
Simple Calculation ◽  
Two Dimensional ◽  
Number Range ◽  
Pressure Loss Coefficient ◽  
Supersonic Inlet ◽  
Profile Losses ◽  
Total Pressure Loss Coefficient

The paper presents two-dimensional cascade results which have been obtained in transonic and supersonic cascade windtunnels. The upstream Mach number range is 1,0 ≤ M1 ≤ 1,4. Tests have been carried out with three different blade shapes; these are double-circular-arc and wedge profiles. The influence of solidity on the performance of these cascades has been investigated. A detailed analysis and calculation of the shock losses shows the great influence of profile shape on the total pressure loss coefficient. The profile losses are roughly constant in the investigated Mach number range. In addition, some measurements for different back pressures are presented. These results are analyzed with the aid of a simple calculation, which shows that the axial velocity-density ratio has to be considered as an important parameter in supersonic cascade measurements.

scholarly journals Investigation of Wash Fluid Preheating on the Effectiveness of Online Compressor Washing in Industrial Gas Turbines

2020 ◽  
Author(s):  
Roupa Agbadede ◽  
Biweri Kainga
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Loss Coefficient ◽  
Industrial Gas Turbines ◽  
Total Pressure Loss Coefficient

Abstract This study presents an investigation of wash fluid preheating on the effectiveness of online compressor washing in industrial gas turbines. Crude oil was uniformly applied on the compressor cascade blades surfaces using a roller brush, and carborundum particles were ingested into the tunnel to create accelerated fouled blades. Demineralized water was preheated to 500C using the heat coil provided in the tank. When fouled blades washed with preheated demineralized and the one without preheating were compared, it was observed that there was little or no difference in terms of total pressure loss coefficient and exit flow angle. However, when the fouled and washed cases were compared, there was a significant different in total pressure loss coefficient and exit flow angle.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2008 ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient ◽  
Aerodynamic Losses

The performance of suction-side gill region film cooling is investigated using the University of Utah Transonic Wind Tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77 to 1.99 similar to values present in operating gas turbine engines. Presented are local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated Aerodynamic Losses (IAL) increase anywhere from 4 to 45 percent compared to a smooth blade with no film injection. The performance of each hole type depends upon the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL aerodynamic losses are generally present at any particular wake location for the round radial RR or shaped axial SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA shaped axial holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Effect of End-Wall Vortex Generator Jet Parameters on Flow Control in High Subsonic Compressor Cascade

2018 ◽  
Author(s):  
Cong Chen ◽  
Jianyang Yu ◽  
Fu Chen
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Control Effect ◽  
Pressure Loss Coefficient ◽  
Vortex Generator Jet ◽  
End Wall ◽  
Total Pressure Loss Coefficient

In order to explore the control mechanism of vortex generator jet, which is located in the passage (PVGJ), on the separation flow, the influence of the pitch angle, skew angle, locations and jet-to-inflow ratio are studied using numerical methods in a high subsonic compressor cascade. The changing of the flow pattern is also analyzed in detail. The results show that the control effect of the end-wall vortex generator jet located in the passage is better than the leading edge one and the aerodynamic performance is effectively improved. The maximum total pressure loss coefficient decreases by 12% and the static pressure coefficient increases by 5.2% while the jet-to-inflow ratio is only 0.3%. The control effect is sensitive to the change of jet parameters. When 0 deg < β < 80 deg, 20 deg < α < 50 deg,, x < 0.5B, y < 0.15t, the vortex generation jet could acquire an ideal control effect. As the jet mass increases, the total pressure loss coefficient gradually reduces. The VGJ prevent separation mainly by bringing high momentum fluid into the near wall region and by promoting momentum transport through turbulent mixing in previous studies. Both the LVGJ and PVGJ mainly take advantage of jet vortex to prevent the cross flow from interacting with the suction side boundary layer.

scholarly journals Accurate 2-D Modelling of Transonic Compressor Cascade Aerodynamics

Aerospace ◽  
2019 ◽  
Vol 6 (5) ◽  
pp. 57 ◽  
Author(s):  
Tommaso Piovesan ◽  
Andrea Magrini ◽  
Ernesto Benini
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Transonic Compressor ◽  
Pressure Loss Coefficient ◽  
Total Pressure Loss Coefficient

Modern aeronautic fans are characterised by a transonic flow regime near the blade tip. Transonic cascades enable higher pressure ratios by a complex system of shockwaves arising across the blade passage, which has to be correctly reproduced in order to predict the performance and the operative range. In this paper, we present an accurate two-dimensional numerical modelling of the ARL-SL19 transonic compressor cascade. A large series of data from experimental tests in supersonic wind tunnel facilities has been used to validate a computational fluid dynamic model, in which the choice of turbulence closure resulted critical for an accurate reproduction of shockwave-boundary layer interaction. The model has been subsequently employed to carry out a parametric study in order to assess the influence of main flow variables (inlet Mach number, static pressure ratio) and geometric parameters (solidity) on the shockwave pattern and exit status. The main objectives of the present work are to perform a parametric study for investigating the effects of the abovementioned variables on the cascade performance, in terms of total-pressure loss coefficient, and on the shockwave pattern and to provide a quite large series of data useful for a preliminary design of a transonic compressor rotor section. After deriving the relation between inlet and exit quantities, peculiar to transonic compressors, exit Mach number, mean exit flow angle and total-pressure loss coefficient have been examined for a variety of boundary conditions and parametrically linked to inlet variables. Flow visualisation has been used to describe the shock-wave pattern as a function of the static pressure ratio. Finally, the influence of cascade solidity has been examined, showing a potential reduction of total-pressure loss coefficient by employing a higher solidity, due to a significant modification of shockwave system across the cascade.

Effects of Riblets on the Loss Behavior of a Highly Loaded Compressor Cascade

2002 ◽  
Author(s):  
Matthias Boese ◽  
Leonhard Fottner
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Pressure Loss Coefficient ◽  
Flow Phenomena ◽  
Total Pressure Loss Coefficient ◽  
Highly Loaded

An experimental investigation of the influence of riblet surface structures on the loss behavior of a highly loaded compressor cascade V103-180 featuring a large chord length for high spatial resolution of the flow phenomena was performed. The cascade experiments were carried out at the High Speed Cascade Wind Tunnel of the University of the Armed Forces Munich in order to simulate realistic Mach and Reynolds numbers. The riblets used for the first investigation are of symmetric v-groove type with heights of 0.0762, 0.1143 and 0.1524 mm, respectively [1]. With two total pressure probes simultaneously traversed over one pitch behind the center airfoil, the local total pressure difference between the structured and the smooth blade is determined. From these measurements, the total pressure loss coefficient can be evaluated. For a better understanding of the flow phenomena, the profile pressure distribution is measured for the smooth and the structured blade. Boundary layer calculations were performed in order to optimise the riblet size for the design conditions of the compressor cascade. Resulting from the measurements an optimised riblet configuration (size and shape) has been manufactured and transferred to the cascade. Further flow measurements have been performed in order to evaluate the total pressure loss coefficient. Additional insight into the flow phenomena of the boundary layer has been achieved by laser-two-focus measurements. The experimental results indicate that the riblets mainly influence the suction side boundary layer behaviour. The ideal dimensionless groove height is obtained h+ = 9 leading to a reduction of the loss coefficient of 6–8%. Values of h+ &gt; 20 cause an increase of the loss coefficient due to the development of a turbulent boundary layer separation.

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.

The Influence of Different Rim Seal Geometries on Hot-Gas Ingestion and Total Pressure Loss in a Low-Pressure Turbine

2010 ◽  
Author(s):  
P. Schuler ◽  
W. Kurz ◽  
K. Dullenkopf ◽  
H.-J. Bauer
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Gas Flow ◽  
Low Pressure ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Low Pressure Turbine ◽  
Hot Gas ◽  
Flow Through ◽  
Total Pressure Loss Coefficient

In order to prevent hot-gas ingestion into the rotating turbo machine’s inside, rim seals are used in the cavities located between stator- and rotor-disc. The sealing flow ejected through the rim seal interacts with the boundary layer of the main gas flow, thus playing a significant role in the formation of secondary flows which are a major contributor to aerodynamic losses in turbine passages. Investigations performed in the EU project MAGPI concentrate on the interaction between the sealing flow and the main gas flow and in particular on the influence of different rim seal geometries regarding the loss-mechanism in a low-pressure turbine passage. Within the CFD work reported in this paper static simulations of one typical low-pressure turbine passage were conducted containing two different rim seal geometries, respectively. The sealing flow through the rim seal had an azimuthal velocity component and its rate has been varied between 0–1% of the main gas flow. The modular design of the computational domain provided the easy exchange of the rim seal geometry without remeshing the main gas flow. This allowed assessing the appearing effects only to the change of rim seal geometry. The results of this work agree with well-known secondary flow phenomena inside a turbine passage and reveal the impact of the different rim seal geometries on hot-gas ingestion and aerodynamic losses quantified by a total pressure loss coefficient along the turbine blade. While the simple axial gap geometry suffers considerable hot-gas ingestion upstream the blade leading edge, the compound geometry implying an axial overlapping presents a more promising prevention against hot-gas ingestion. Furthermore, the effect of rim seals on the turbine passage flow field has been identified applying adequate flow visualisation techniques. As a result of the favourable conduction of sealing flow through the compound geometry, the boundary layer is less lifted by the ejected sealing flow, thus resulting in a comparatively reduced total pressure loss coefficient over the turbine blade.

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