The Effect of Variable Chord Length on Transonic Axial Rotor Performance

2002 ◽  
Vol 124 (3) ◽  
pp. 351-357 ◽  
Author(s):  
William B. Roberts ◽  
Albert Armin ◽  
George Kassaseya ◽  
Kenneth L. Suder ◽  
Scott A. Thorp ◽  
...  
Keyword(s):  
Leading Edge ◽  
Aerodynamic Performance ◽  
Chord Length ◽  
Rotor Blades ◽  
Civil Aviation ◽  
Turbofan Engines ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Significant Difference ◽  
Leading Edges

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94–96% of nominal (new) blade chord.

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

2001 ◽  
Author(s):  
William B. Roberts ◽  
Albert Armin ◽  
George Kassaseya ◽  
Kenneth L. Suder ◽  
Scott A. Thorp ◽  
...  
Keyword(s):  
Leading Edge ◽  
Aerodynamic Performance ◽  
Chord Length ◽  
Rotor Blades ◽  
Civil Aviation ◽  
Turbofan Engines ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Significant Difference ◽  
Leading Edges

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100%, 80% and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

Characterization of Shock Wave–Endwall Boundary Layer Interactions in a Transonic Compressor Rotor

1988 ◽  
Vol 110 (3) ◽  
pp. 386-392 ◽  
Author(s):  
D. C. Rabe ◽  
A. J. Wennerstrom ◽  
W. F. O’Brien
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Air Force ◽  
Leading Edge ◽  
Transonic Compressor ◽  
Laser Measurements ◽  
Compressor Rotor ◽  
Boundary Layer Interaction ◽  
Air Force Base

The passage shock wave–endwall boundary layer interaction in a transonic compressor was investigated with a laser transit anemometer. The transonic compressor used in this investigation was developed by the General Electric Company under contract to the Air Force. The compressor testing was conducted in the Compressor Research Facility at Wright-Patterson Air Force Base, OH. Laser measurements were made in two blade passages at seven axial locations from 10 percent of the axial blade chord in front of the leading edge to 30 percent of the axial blade chord into the blade passage. At three of these axial locations, laser traverses were taken at different radial immersions. A total of 27 different locations were traversed circumferentially. The measurements reveal that the endwall boundary layer in this region is separated from the core flow by what appears to be a shear layer where the passage shock wave and all ordered flow seem to end abruptly.

Stall Inception and Development Process Due to Tip Leakage Flow in Axial Compressor Rotor Blades Row

2014 ◽  
Vol 30 (3) ◽  
pp. 307-313 ◽  
Author(s):  
R. Taghavi-Zenou ◽  
S. Abbasi ◽  
S. Eslami
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Axial Compressor ◽  
Leading Edge ◽  
Rotor Blades ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Compressor Rotor

ABSTRACTThis paper deals with tip leakage flow structure in subsonic axial compressor rotor blades row under different operating conditions. Analyses are based on flow simulation utilizing computational fluid dynamic technique. Three different circumstances at near stall condition are considered in this respect. Tip leakage flow frequency spectrum was studied through surveying instantaneous static pressure signals imposed on blades surfaces. Results at the highest flow rate, close to the stall condition, showed that the tip vortex flow fluctuates with a frequency close to the blade passing frequency. In addition, pressure signals remained unchanged with time. Moreover, equal pressure fluctuations at different passages guaranteed no peripheral disturbances. Tip leakage flow frequency decreased with reduction of the mass flow rate and its structure was changing with time. Spillage of the tip leakage flow from the blade leading edge occurred without any backflow in the trailing edge region. Consequently, various flow structures were observed within every passage between two adjacent blades. Further decrease in the mass flow rate provided conditions where the spilled flow ahead of the blade leading edge together with trailing edge backflow caused spike stall to occur. This latter phenomenon was accompanied by lower frequencies and higher amplitudes of the pressure signals. Further revolution of the rotor blade row caused the spike stall to eventuate to larger stall cells, which may be led to fully developed rotating stall.

scholarly journals Low Aspect Ratio Transonic Rotors: Part 2 — Influence of Location of Maximum Thickness on Transonic Compressor Performance

1992 ◽  
Author(s):  
A. R. Wadia ◽  
C. H. Law
Keyword(s):  
Aspect Ratio ◽  
Test Data ◽  
Three Dimensional ◽  
Leading Edge ◽  
Chord Length ◽  
Design Parameters ◽  
Transonic Compressor ◽  
Low Aspect Ratio ◽  
Maximum Thickness ◽  
Design Speed

Transonic compressor rotor performance is sensitive to variations in several known design parameters. One such parameter is the chordwise location of maximum thickness. This article reports on the design and experimental evaluation of two versions of a low aspect ratio transonic rotor that had the location of the tip blade section maximum thickness moved forward in two increments from the nominal 70 percent to 55 and 40 percent chord length, respectively. The original hub characteristics were preserved and the maximum thickness location was adjusted proportionately along the span. Although designed to satisfy identical design speed requirements, the experimental results reveal significant variation in the performance of the rotors. At design speed, the rotor with its maximum thickness located at 55 percent chord length attains the highest peak efficiency amongst the three rotors but has lowest flow rollback relative to the other two versions. To focus on current ruggedization issues for transonic blading (e.g. bird, ice ingestion), detailed comparison of test data and analysis to characterize the aerodynamic flow details responsible for the measured performance differences was confined to the two rotors with the most forward location of maximum thickness. A three-dimensional viscous flow analysis was used to identify the performance enhancing features of the higher efficiency rotor and to provide guidance in the interpretation of the experimental measurements. The computational results of the viscous analysis shows that the difference in performance between the two rotors can be attributed to the higher shock losses that result from the increased leading edge “wedge angle” as the maximum thickness is moved closer to the leading edge. The test data and the three-dimensional viscous analysis also reveal that the higher efficiency rotor achieves the same static pressure rise potential and loading at a higher flow level than its lesser efficient counterpart and this is responsible for its resulting lower flow rollback and apparent loss in stall margin. Comparison of the peak efficiencies attained by the two rotors described in this article with the baseline ruggedized rotor performance presented in part 1 of this paper suggests the existence of an optimum maximum thickness location at 55 to 60 percent chord length for such low aspect ratio transonic rotors.

Investigating the effects of leading-edge tubercles on the aerodynamic performance of insect-like flapping wing

2020 ◽  
pp. 095440622094635
Author(s):  
Muhammad Bilal Anwar ◽  
Aamer Shahzad ◽  
Muhammad Nafees Mumtaz Qadri
Keyword(s):  
Leading Edge ◽  
Aerodynamic Performance ◽  
Flapping Wing ◽  
Chord Length ◽  
Humpback Whales ◽  
Power Coefficient ◽  
Pectoral Fin ◽  
Aerodynamic Efficiency ◽  
Potential Benefits ◽  
Better Than

Tubercles are small protuberances or bumps on the leading edge of humpback whale's pectoral fin. To examine the effects of leading-edge tubercles on the aerodynamic performance of a flapping wing, lift, drag, and power coefficients are obtained from numerical simulations. A revolving wing (one-degree-of-freedom azimuth rotation; rotation in a horizontal plane after an initial acceleration) with leading-edge tubercles at an angle of attack of 40° and Reynolds number of 400 is used in the present study. The reason for choosing azimuth rotation is that it resembles downstroke and upstroke of flapping motion of an insect. A rigid rectangular wing with six different combinations of wavelengths ( λ = 10% and 50% of the chord length) and amplitudes ( A = 2.5%, 5%, and 10% of the chord length) are chosen for this study. These parameters are inspired by the tubercles present at the leading edge of humpback whales' pectoral fin. It was observed that generally, tubercles degraded the aerodynamic performance of the wings in terms of lift, drag, and power coefficients. Although some of the tubercle leading-edge wings showed lower drag (2.20% lower) and lower power coefficient (2.12% lower) values than the baseline wing, none of the tubercle wing performed better than the baseline wing in terms of aerodynamic performance parameters; aerodynamic efficiency ([Formula: see text]) and power economy ([Formula: see text]). Hence, it was concluded that the tubercles are not advantageous over the straight leading-edge wing for azimuth rotating hovering insect-like motion and further investigation is required to explore its potential benefits.

scholarly journals The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

1995 ◽  
Vol 117 (4) ◽  
pp. 491-505 ◽  
Author(s):  
K. L. Suder ◽  
R. V. Chima ◽  
A. J. Strazisar ◽  
W. B. Roberts
Keyword(s):  
Surface Roughness ◽  
Surface Finish ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Blade Surface ◽  
Suction Surface ◽  
Compressor Rotor ◽  
Performance Deterioration ◽  
Design Speed

The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 rms μm (100–125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 rms μm (10–20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms pm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier–Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

IDDES Study of the Shock Induced Flow Separation in a Transonic Compressor Rotor at Near Stall Condition

2014 ◽  
Author(s):  
Ke Shi ◽  
Song Fu ◽  
Scott C. Morris
Keyword(s):  
Flow Separation ◽  
Leading Edge ◽  
Side Surface ◽  
Detached Eddy Simulation ◽  
Tip Leakage Flow ◽  
Suction Surface ◽  
High Entropy ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Blade Leading Edge

The IDDES (Improved Delayed Detached Eddy Simulation) method was applied to simulate the highly unsteady flow in a transonic compressor at near-stall operating condition. Shock induced separated flow on the casing wall in front of the blade leading edge and the flow separation induced by the shock impinging on the blade suction surface were investigated in detail with the help of IDDES. The 3D separation in the corner of blade suction side surface and the casing wall in the front portion of the blade was shown to be a vortex-like separation spiraling out from the blade suction surface, connecting to the blade suction wall and the casing wall. Blade tip leakage flow also contributed to the formation of the separation vortex. At near stall condition, shock wave was pushed forward out of the blade passage. The upstream propagation of the separation vortex and the related high entropy region were considered to be a characteristic phenomenon of the flow at near stall condition in this transonic compressor rotor. The results of the present study explain the origin and the formation of the high entropy region in a transonic compressor rotor near the blade leading edge, which can be closely related to the spike type stall inception observed by other researchers in the transonic case.

Near Stall Behavior of a Transonic Compressor Rotor With Casing Treatment

2016 ◽  
Author(s):  
Christoph Brandstetter ◽  
Felix Holzinger ◽  
Heinz-Peter Schiffer ◽  
Sina Stapelfeldt ◽  
Mehdi Vahdati
Keyword(s):  
Stability Limit ◽  
Rotating Stall ◽  
Rotor Blades ◽  
Transient Effects ◽  
Stability Range ◽  
Blade Vibration ◽  
Casing Treatment ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Engine Order

The aerodynamic and aeroelastic performance of an advanced axial slot casing treatment (CT) was investigated on a modern one and a half stage transonic compressor test rig. It is generally accepted that a well designed CT can extend the aerodynamic stability range of a compressor to lower mass flows. The extension of stall margin of the compressor rotor blades by using CT has been the subject of numerous research articles but much less attention has been paid to the behavior of the compressor in direct vicinity of the stability limit. For the compressor investigated here, two different phenomena were repeatedly observed near stall: 1) self-excited blade vibration, and 2) low engine order fluctuations developing into rotating stall. The current investigation firstly aims to identify the triggers for each of these phenomena. It then focusses on the aerodynamic and aeromechanical mechanism which lead to the formation of low engine order fluctuations shortly before stall. In order to measure the unsteady and transient effects, the system was instrumented with unsteady wall pressure transducers, a capacitive tip-timing system and strain gauges on the rotor blades. The flow structure in the blade tip region was measured via Particle Image Velocimetry underneath the CT-Cavities. Measurements showed a strong correlation between CT activity and the development of the low frequency oscillations with associated blade vibrations. Using numerical simulations, presented and validated in this paper, this correlation was attributed to an aerodynamic coupling between rotor passages through the recirculation of fluid inside the cavities.

Transient Analysis of Flow Unsteadiness and Noise Characteristics in a Centrifugal Compressor with a Novel Vaned Diffuser

2021 ◽  
Vol 11 (7) ◽  
pp. 3191
Author(s):  
Ali Zamiri ◽  
Kun Sung Park ◽  
Minsuk Choi ◽  
Jin Taek Chung
Keyword(s):  
Centrifugal Compressor ◽  
Sound Pressure Level ◽  
Sound Pressure ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Pressure Level ◽  
Stall Margin ◽  
Transonic Centrifugal Compressor ◽  
Inclination Angles ◽  
Leading Edges

The demands to apply transonic centrifugal compressor have increased in the advanced gas turbine engines. Various techniques are used to increase the aerodynamic performance of the centrifugal compressor. The effects of the inclined leading edges in diffuser vanes of a transonic centrifugal compressor on the flow-field unsteadiness and noise generation are investigated by solving the compressible, three-dimensional, transient Navier–Stokes equations. Diffuser vanes with various inclination angles of the leading edge from shroud-to-hub and hub-to-shroud are numerically modeled. The results show that the hub-to-shroud inclined leading edge improves the compressor performance (2.6%), and the proper inclination angle is effective to increase the stall margin (3.88%). In addition, in this study, the transient pressure variations and radiated noise prediction at the design operating point of the compressor are emphasized. The influences of the inclined leading edges on the pressure waves were captured in time/space domain with different convective velocities. The pressure fluctuation spectra are calculated to investigate the tonal blade passing frequency (BPF) noise, and it is shown that the applied inclination angles in the diffuser blades are effective, not only to improve the aerodynamic performance and stall margin, but also to reduce the BPF noise (7.6 dB sound pressure level reduction). Moreover, it is found that the diffuser vanes with inclination angles could suppress the separation regions and eddy structures inside the passages of the diffuser, which results in reduction of the overall sound pressure level and the broadband noise radiated from the compressor.

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