Performance Analysis of the Test Results on a Two-Stage Transonic Fan

1983 ◽  
Vol 105 (1) ◽  
pp. 125-129
Author(s):  
Baoshi Chen ◽  
Tianyi Zhang
Keyword(s):  
Mach Number ◽  
Performance Analysis ◽  
Radial Distribution ◽  
Flow Separation ◽  
High Load ◽  
Test Results ◽  
Two Stage ◽  
Stall Margin ◽  
High Incidence ◽  
Transonic Fan

Test results obtained from a two-stage fan are analysed and the reasons that caused the design performance target not to be attained are presented in this paper. Addition of a partspan shroud on rotor 1 caused higher losses and changed radial distribution of parameters. Modification on the flowpath and chord length of stator 1 resulted in excessively high inlet Mach number and flow separation in the hub region. The high load and high incidence at the hub of rotor 2 caused higher losses and reduced stall margin of the fan.

Aerodynamic Design and Testing of an Imbedded Forward Swept Rotor in a Two-Stage Transonic Fan

2018 ◽  
Author(s):  
A. R. Wadia ◽  
J. D. Niedermeier ◽  
P. N. Szucs ◽  
N. G. Cormier ◽  
D. W. Crall ◽  
...  
Keyword(s):  
Leading Edge ◽  
Aerodynamic Design ◽  
Two Stage ◽  
Stall Margin ◽  
Second Stage ◽  
Inlet Distortion ◽  
Overall Performance ◽  
Stage Matching ◽  
Stage 1 ◽  
Transonic Fan

Previous experimental and analytical studies comparing the performance of transonic swept rotors in single and multistage fans have demonstrated the potential of large improvements in clean inlet performance and substantial improvements in fan sensitivity with inlet distortion with forward swept blading. A two-stage, low-aspect ratio transonic fan investigation was previously conducted in the Air Force’s Compressor Research Facility, in two builds on a back-to-back test basis, using a radial and a forward swept stage 1 blade. While the forward swept stage 1 blade configuration did demonstrate superior front stage efficiency and tolerance to inlet distortion, the common second stage among the two builds prevented the overall fan from showing clean inlet performance and stability benefits with the forward swept Rotor 1. To address this measured overall performance shortfall, this paper reports on the design of a new second stage blade tested in the same two-stage fan rig with the forward swept stage 1 blade configuration. The new second stage blade was designed with forward sweep to improve efficiency and operability while replicating the baseline radial Rotor 2’s aerodynamic design conditions within the same flow path. The design point requirements of the forward swept Rotor 2 were selected to preserve the internal stage matching with the radially stacked Rotor 2. As the new stage 2 blade had to fit within the existing radial Rotor 2’s physical envelope, the new blade was designed with forward sweep through lean only, which proved to be quite challenging from a mechanical growth and deflection view point. The first attempt to run the fan rig with the new stage 2 blade resulted in a leading-edge tip rub during a part speed stall event. However, even with this unfortunate event, fan mapping test results with clean inlet from part speed to 97.5 percent design speed showed a significant improvement in overall fan efficiency and stall margin, validating the hypothesis that in the earlier tests stage 2 was indeed the limiting stage that prevented the fan from reaching its overall performance goals. Based on this experience and the test data acquired with unstable leadingedge tip rubs during stall deflections with forward swept airfoils leaned in the direction of rotation, a process was developed to determine the acceptability criteria of such blading.

Aerodynamic Design and Testing of an Imbedded Forward Swept Rotor in a Two-Stage Transonic Fan

2019 ◽  
Vol 141 (8) ◽  
Author(s):  
A. R. Wadia ◽  
J. D. Niedermeier ◽  
P. N. Szucs ◽  
N. G. Cormier ◽  
D. W. Crall ◽  
...  
Keyword(s):  
Leading Edge ◽  
Aerodynamic Design ◽  
Two Stage ◽  
Stall Margin ◽  
Second Stage ◽  
Inlet Distortion ◽  
Overall Performance ◽  
Stage Matching ◽  
Stage 1 ◽  
Transonic Fan

Previous experimental and analytical studies comparing the performance of transonic swept rotors in single and multistage fans have demonstrated the potential of large improvements in clean inlet performance and substantial improvements in fan sensitivity with inlet distortion with forward swept blading. A two-stage, low-aspect ratio transonic fan investigation was previously conducted in the Air Force’s Compressor Research Facility, in two builds on a back-to-back test basis, using a radial and a forward swept stage 1 blade. While the forward swept stage 1 blade configuration did demonstrate superior front stage efficiency and tolerance to inlet distortion, the common second stage among the two builds prevented the overall fan from showing clean inlet performance and stability benefits with the forward swept rotor 1. To address this measured overall performance shortfall, this paper reports on the design of a new second stage blade tested in the same two-stage fan rig with the forward swept stage 1 blade configuration. The new second stage blade was designed with forward sweep to improve efficiency and operability while replicating the baseline radial rotor 2’s aerodynamic design conditions within the same flow path. The design point requirements of the forward swept rotor 2 were selected to preserve the internal stage matching with the radially stacked rotor 2. As the new stage 2 blade had to fit within the existing radial rotor 2’s physical envelope, the new blade was designed with forward sweep through lean only, which proved to be quite challenging from a mechanical growth and deflection view point. The first attempt to run the fan rig with the new stage 2 blade resulted in a leading-edge tip rub during a part speed stall event. However, even with this unfortunate event, fan mapping test results with clean inlet from part speed to 97.5% design speed showed a significant improvement in overall fan efficiency and stall margin, validating the hypothesis that in the earlier tests stage 2 was indeed the limiting stage that prevented the fan from reaching its overall performance goals. Based on this experience and the test data acquired with unstable leading-edge tip rubs during stall deflections with forward swept airfoils leaned in the direction of rotation, a process was developed to determine the acceptability criteria of such blading.

High Specific Speed, High Inducer Tip Mach Number, Centrifugal Compressor

2003 ◽  
Author(s):  
C. Rodgers
Keyword(s):  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
High Efficiency ◽  
Flow Fields ◽  
Test Results ◽  
Positive Slope ◽  
Stall Margin ◽  
Recovery Characteristic ◽  
Specific Speed

The demand for more efficient turbocharger and aviation centrifugal compressors operating at higher pressure ratios and specific speeds with extended flow ranges is focusing research efforts on the inducer and diffuser transonic flow fields. At pressure ratios above 5.0 and specific speeds of unity inducer tip relative Mach numbers exceeding 1.4 can be encountered, precipitating both increased shock losses and diminished stall margin. The results of compressor rig testing on a research 6.8 inch tip (173mm) diameter single stage centrifugal compressor operating with inducer tip relative Mach number up to 1.5 are presented. The test results reveal high efficiency combined with extended flow range. This was achieved through improved impeller stability with shroud bleed, thereby permitting the diffuser to operate stably on its positive slope recovery characteristic.

Static-pressure probes that are theoretically insensitive to pitch, yaw and Mach number

1970 ◽  
Vol 44 (3) ◽  
pp. 513-528 ◽  
Author(s):  
A. M. O. Smith ◽  
A. B. Bauer
Keyword(s):  
Mach Number ◽  
Cross Sections ◽  
Flow Separation ◽  
Potential Flow ◽  
Static Pressure ◽  
Angle Of Attack ◽  
Test Results ◽  
Cross Sectional

The idea of distributing static probe cross-sectional areas so as to render the probe insensitive to Mach number is combined here with that of using non-circular cross-sections to render probes insensitive to yaw and angle of attack. Appropriate non-circular cross-sections are described in detail, and a general means of designing blunt or slender probes to have zero sensitivity to yaw and angle of attack in potential flow is described. Four experimental probes have been tested, and test results are presented. These results show that the probes are quite insensitive to yaw and angle of attack within certain limiting angles, which are assumed to correspond to the onset of flow separation.

scholarly journals Performance analysis of hydrogen fuel cell with two-stage turbo compressor for automotive applications

2021 ◽  
Vol 7 ◽  
pp. 2635-2646
Author(s):  
Nabeel Ahsan ◽  
Ans Al Rashid ◽  
Asad A. Zaidi ◽  
Ramsha Imran ◽  
Sikandar Abdul Qadir
Keyword(s):  
Fuel Cell ◽  
Performance Analysis ◽  
Hydrogen Fuel Cell ◽  
Automotive Applications ◽  
Hydrogen Fuel ◽  
Two Stage ◽  
Turbo Compressor

Measurements of Endwall Flows in Transonic Linear Turbine Cascades: Part II—High Flow Turning

2010 ◽  
Author(s):  
F. Taremi ◽  
S. A. Sjolander ◽  
T. J. Praisner
Keyword(s):  
Mach Number ◽  
Kinetic Energy ◽  
Flow Visualization ◽  
Flow Separation ◽  
Surface Flow ◽  
Streamwise Vorticity ◽  
Vortical Structures ◽  
Turbine Cascades ◽  
Mach Numbers ◽  
Surface Flow Visualization

An experimental investigation of two low-turning (90°) transonic linear turbine cascades was presented in Part I of the paper. Part II examines two high-turning (112°) turbine cascades. The experimental results include total pressure losses, streamwise vorticity and secondary kinetic energy distributions. The measurements were made using a seven-hole pressure probe downstream of the cascades. In addition to the measurements, surface flow visualization was conducted to assist in the interpretation of the flow physics. The turbine cascades in Part II, referred to as SL1F and SL2F, have the same inlet and outlet design flow angles, but different aerodynamic loading levels: SL2F is more highly loaded than SL1F. The surface flow visualization results show evidence of small flow separation on the suction side of both airfoils. At the design conditions (outlet Mach number ≈ 0.8), SL2F exhibits stronger vortical structures and larger secondary velocities than SL1F. The two cascades, however, produce similar row losses based on the measurements at 40% axial chord lengths downstream of the trailing edge. Additional data were collected at off-design outlet Mach numbers of 0.65 and 0.91. As the Mach number is raised, the cascades become more aft-loaded. The absolute blade loadings increase, but the Zweifel coefficients decrease due to higher outlet dynamic pressures. Both profile and secondary losses decrease at higher Mach numbers; the main vortical structures and the corresponding peak losses migrate towards the endwall, and there are reductions in secondary kinetic energy and exit flow angle variations. The streamwise vorticity distributions show smaller peak vorticities associated with the passage and the counter vortices at higher exit Mach numbers. The corner vortex, on the other hand, becomes more intensified, resulting in reduction of flow overturning near the endwall. The results for SL1F and SL2F are compared and contrasted with the results for the lower turning cascades presented in Part I. The possible effects of suction-surface flow separation on profile and secondary losses are discussed in this context. The current research project is part of a larger study concerning the effects of endwall contouring on secondary losses, which will be presented in the near future.

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