Inner Workings of Shrouded and Unshrouded Transonic Fan Blades

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
A. R. Wadia ◽  
P. N. Szucs
Keyword(s):  
Boundary Layer ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Suction Side ◽  
Operating Conditions ◽  
Radial Migration ◽  
Boundary Layer Interaction ◽  
Improved Performance ◽  
Design Speed

This paper reports on the numerical assessment of the differences in aerodynamic performance between part span shrouded and unshrouded fan blades generally found in the first stage of multistage fans in low bypass ratio aircraft engines. Rotor flow fields for both blade designs were investigated at two operating conditions using a three-dimensional viscous flow analysis. Although designed to the same radius ratio, aspect ratio, and solidity, the unshrouded fan rotor had a slightly increased tip speed (+3%) and somewhat lower pressure ratio (−3.2%) due to engine cycle requirements. Even when allowing for these small differences, the analysis reveals interesting differences in the level and in the radial distribution of efficiency between these two rotors. The reason for the improved performance of the shrouded rotor in part can be attributed to the shroud blocking off the radial migration of boundary layer fluid centrifuged from the hub on the suction side. As a result, the shock boundary layer interaction seems to be improved on the shrouded blade. At the cruise condition, the efficiency is the same for both rotors. The slightly better efficiency of the shrouded blade in the outer panel is nullified by the large efficiency penalty in the vicinity of the shroud. As there is no significant radial migration of fluid in the suction side boundary layer, as indicated by the analysis at this condition relative to the design speed case, the benefit due to the shroud is greatly reduced. At this speed and at lower speeds, the shroud becomes a net additional loss for the blade. Also of interest from the numerical results is the indication that significant blade ruggedization penalties to performance can be reduced in the case of the unshrouded blade through custom tailoring of its mean camber line.

Inner Workings of Shrouded and Un-Shrouded Transonic Fan Blades

2006 ◽  
Author(s):  
A. R. Wadia ◽  
P. N. Szucs
Keyword(s):  
Boundary Layer ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Suction Side ◽  
Operating Conditions ◽  
Radial Migration ◽  
Multi Stage ◽  
Boundary Layer Interaction ◽  
Improved Performance

This paper reports on the numerical assessment of the differences in aerodynamic performance between part span shrouded and un-shrouded fan blades generally found in the first stage of multi-stage fans in low bypass ratio aircraft engines. Rotor flow fields for both blade designs were investigated at two operating conditions using a three-dimensional viscous flow analysis. Although designed to the same radius ratio, aspect ratio and solidity, the un-shrouded fan rotor had a slightly increased tip speed (+3%) and somewhat lower pressure ratio (-3.2%) due to engine cycle requirements. Even when allowing for these small differences, the analysis reveals interesting differences in the level and in the radial distribution of efficiency between these two rotors. The reason for the improved performance of the shrouded rotor in part can be attributed to the shroud blocking off the radial migration of boundary layer fluid centrifuged from the hub on the suction side. As a result, the shock boundary layer interaction seems to be improved on the shrouded blade. At the cruise condition, the efficiency is the same for both rotors. The slightly better efficiency of the shrouded blade in the outer panel is nullified by the large efficiency penalty in the vicinity of the shroud. As there is no significant radial migration of fluid in the suction side boundary layer as indicated by the analysis at this condition relative to the design speed case; the benefit due to the shroud is greatly reduced. At this speed and at lower speeds the shroud becomes a net additional loss for the blade. Also of interest from the numerical results, is the indication that significant blade ruggedization penalties to performance can be reduced in the case of the un-shrouded blade through custom tailoring of its mean camber line.

scholarly journals Inner Workings of Aerodynamic Sweep

1997 ◽  
Author(s):  
A. R. Wadia ◽  
P. N. Szucs ◽  
D. W. Crall
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Clearance ◽  
Surface Boundary ◽  
Surface Boundary Layer ◽  
Stall Margin ◽  
Design Speed ◽  
The Impact

The recent trend in using aerodynamic sweep to improve the performance of transonic blading has been one of the more significant technological evolutions for compression components in turbomachinery. This paper reports on the experimental and analytical assessment of the pay-off derived from both aft and forward sweep technology with respect to aerodynamic performance and stability. The single stage experimental investigation includes two aft-swept rotors with varying degree and type of aerodynamic sweep and one swept forward rotor. On a back-to-back test basis, the results are compared with an unswept rotor with excellent performance and adequate stall margin. Although designed to satisfy identical design speed requirements as the unswept rotor, the experimental results reveal significant variations in efficiency and stall margin with the swept rotors. At design speed, all the swept rotors demonstrated a peak stage efficiency level that was equal to that of the unswept rotor. However, the forward-swept rotor achieved the highest rotor-alone peak efficiency. At the same time, the forward-swept rotor demonstrated a significant improvement in stall margin relative to the already satisfactory level achieved by the unswept rotor. Increasing the level of aft sweep adversely affected the stall margin. A three-dimensional viscous flow analysis was used to assist in the interpretation of the data. The reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the up, was identified as the prime contributor to the enhanced performance with forward sweep. The impact of tip clearance on the performance and stability for one of the aft-swept rotors was also assessed.

scholarly journals The Effect of the Inlet Boundary Layer on the Tip Flow Field in a Transonic Fan Rotor

1998 ◽  
Author(s):  
Junji Takado ◽  
Toyotaka Sonoda ◽  
Satoshi Nakamura
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Interaction Region ◽  
Operating Conditions ◽  
Laser Doppler Velocimeter ◽  
Intensity Level ◽  
Pressure Surface ◽  
Transonic Fan

Experimental and numerical investigations have been carried out to understand the effects of the inlet boundary layer (IBL) on the tip flow field including the aerodynamic performance in a transonic fan rotor. Both the steady and the unsteady phenomena in the tip flow field have been investigated for operating conditions near peak efficiency and near stall with the two types of tip IBL. In order 10 study these phenomena, high response pressure data with Kulite transducers and laser doppler velocimeter (LDV) data have been acquired around the tip region. Furthermore, three-dimensional Navier-Stokes numerical simulations have been compared with the measured results. The results indicate that the tip IBL significantly influences the spanwise distribution of pressure ratio around the tip region and the stall characteristics including the passage shock / tip leakage vortex interaction, the blockage generation, the wake structure, and the unsteadiness of the tip flow field. In particular, at a near stall condition for the thick IBL with high turbulence intensity level, the tip diffusion level is increased due to a larger blockage, which is generated downstream of a much stronger interaction region. These phenomena are a consequence of the low momentum fluid in the tip IBL, and significantly reduce the stall margin. Furthermore, the unsteadiness drastically increases around the interaction region and around the pressure surface where the blockage migrates. These unsteady phenomena are distinctive features near stall. Downstream of the rotor, the larger and more unsteady blockage is discharged from the pressure surface side, and complicates the three-dimensional rotor exit flow field around the tip region.

Aerodynamic Performance of Large Centrifugal Compressors

1982 ◽  
Vol 104 (4) ◽  
pp. 796-804 ◽  
Author(s):  
Fumikata Kano ◽  
Noriyuki Tazawa ◽  
Yoshiteru Fukao
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Design Procedure ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
Boundary Layer Theory ◽  
Air Separation ◽  
Three Dimensional Flow ◽  
Dimensional Measurements

The aerodynamic performance of impellers and diffusers of the large centrifugal compressor were studied. A performance design procedure based on the quasi-three-dimensional flow analysis which is combined with the boundary layer theory was developed. The conditions of the boundary layer at the impeller exit and at the diffuser vane throat were calculated, and the three-dimensional measurements were carried out. This result shows that the low momentum flow is accumulated at the corner of the shroud and the blade suction side of the impeller. These results were applied to the development of a large four-stage isothermal compressor which handles the air for an air separation apparatus. This was tested in the field and showed an isothermal efficiency of 76 percent.

Advanced Non-Axisymmetric Endwall Contouring for Axial Compressors by Generating an Aerodynamic Separator—Part II: Experimental and Numerical Cascade Investigation

2009 ◽  
Author(s):  
Alexander Hergt ◽  
Christian Dorfner ◽  
Wolfgang Steinert ◽  
Eberhard Nicke ◽  
Heinz-Adolf Schreiber
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Suction Side ◽  
Significant Loss ◽  
Optimization Techniques ◽  
Loss Reduction ◽  
Axial Compressors ◽  
Boundary Layer Interaction ◽  
Flow Phenomena

Modern methods for axial compressor design are capable of shaping the blade surfaces in a three dimensional way. Linking these methods with automated optimization techniques provides a major benefit to the design process. The application of non-axisymmetric contoured endwalls is considered to be very successful in turbine rotors and vanes. Concerning axial compressors non-axisymmetric endwalls are still a field of research. This two-part paper presents the recent development of a novel endwall design. A vortex created by a nonaxisymmetric endwall groove acts as an aerodynamic separator, preventing the passage vortex from interacting with the suction side boundary layer. This major impact on the secondary flow results in a significant loss reduction by means of load redistribution, reduction of recirculation areas and suppressed corner separation. Part I of this paper deals with the endwall design and its compressor application. The resulting flow phenomena and physics are described and analysed in detail. The second paper presents the detailed experimental and numerical investigation of the developed endwall groove. The measurements carried out at the transonic cascade wind tunnel of DLR in Cologne, demonstrated a considerable influence on the cascade performance. A loss reduction and redistribution of the cascade loading were achieved at the aerodynamic design point as well as near the stall condition of the cascade. This behaviour is well predicted by the numerical simulation. The combined analysis of experimental and numerical flow patterns allows a detailed interpretation and description of the resulting flow phenomena. In this context high fidelity 3D-RANS flow simulations are required to analyse the complex blade and endwall boundary layer interaction.

Inner Workings of Aerodynamic Sweep

1998 ◽  
Vol 120 (4) ◽  
pp. 671-682 ◽  
Author(s):  
A. R. Wadia ◽  
P. N. Szucs ◽  
D. W. Crall
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Clearance ◽  
Surface Boundary ◽  
Surface Boundary Layer ◽  
Stall Margin ◽  
Design Speed ◽  
The Impact

The recent trend in using aerodynamic sweep to improve the performance of transonic blading has been one of the more significant technological evolutions for compression components in turbomachinery. This paper reports on the experimental and analytical assessment of the pay-off derived from both aft and forward sweep technology with respect to aerodynamic performance and stability. The single-stage experimental investigation includes two aft-swept rotors with varying degree and type of aerodynamic sweep and one swept forward rotor. On a back-to-back test basis, the results are compared with an unswept rotor with excellent performance and adequate stall margin. Although designed to satisfy identical design speed requirements as the unswept rotor, the experimental results reveal significant variations in efficiency and stall margin with the swept rotors. At design speed, all the swept rotors demonstrated a peak stage efficiency level that was equal to that of the unswept rotor. However, the forward-swept rotor achieved the highest rotor-alone peak efficiency. At the same time, the forward-swept rotor demonstrated a significant improvement in stall margin relative to the already satisfactory level achieved by the unswept rotor. Increasing the level of aft swept adversely affected the stall margin. A three-dimensional viscous flow analysis was used to assist in the interpretation of the data. The reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the tip, was identified as the prime contributor to the enhanced performance with forward sweep. The impact of tip clearance on the performance and stability for one of the aft-swept rotors was also assessed.

scholarly journals Design and Evaluation of a Single Passage Test Model to Obtain Turbine Airfoil Film Cooling Effectiveness Data

1995 ◽  
Author(s):  
Frederick A. Buck ◽  
Chander Prakash
Keyword(s):  
Boundary Layer ◽  
Film Cooling ◽  
Density Ratio ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Suction Side ◽  
Test Model ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Single Passage

A single passage test model has been designed to simulate the mainstream aerodynamics between two adjacent turbine airfoils and to measure the film cooling effectiveness from coolant injection on the pressure and suction sides of the airfoils. Film cooling tests were run on the model using a gas concentration/mass transfer technique with a foreign gas as the coolant to match density ratio. Aspects of the design and test are discussed including the use of a two-dimensional inviscid flow analysis to design boundary layer bleeds upstream of the pressure- and suction-side airfoil surfaces. Results of two- and three-dimensional viscous flow analyses that were used to evaluate various design features including inlet bellmouth, boundary layer bleeds, adjustable tailboards and model backpressure are presented. Aerodynamic and film cooling effectiveness test measurements made with the model will show that the model flow field can be controlled to match results from a previous thermal cascade test.

scholarly journals Aerodynamic Performance of Large Centrifugal Compressors

1982 ◽  
Author(s):  
Fumikata Kano ◽  
Noriyuki Tazawa ◽  
Yoshiteru Fukao
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Design Procedure ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
Boundary Layer Theory ◽  
Air Separation ◽  
Three Dimensional Flow ◽  
Dimensional Measurements

The aerodynamic performance of impellers and diffusers of the large centrifugal compressor were studied. A performance design procedure based on the quasi-three-dimensional flow analysis which is combined with the boundary layer theory was developed. The conditions of the boundary layer at the impeller exit and at the diffuser vane throat were calculated, and the three-dimensional measurements were carried out. This result shows that the low momentum flow is accumulated at the corner of the shroud and the blade suction side of the impeller. These results were applied to the development of a large four-stage isothermal compressor which handles the air for an air separation apparatus. This was tested in the field and showed an isothermal efficiency of 76 percent.

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