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.

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.

scholarly journals Experimental Investigation of a Forward Swept Rotor in a Multistage Fan with Inlet Distortion

2011 ◽  
Vol 2011 ◽  
pp. 1-11 ◽  
Author(s):  
Aspi R. Wadia
Keyword(s):  
Experimental Investigation ◽  
Flow Distortion ◽  
Two Stage ◽  
Guide Vanes ◽  
Stall Margin ◽  
Second Stage ◽  
Inlet Distortion ◽  
Inlet Guide Vanes ◽  
Current Production ◽  
Stage 1

Previous studies of transonic swept rotors in single stage fans have demonstrated the potential of significant improvements in both efficiency and stall margin with forward swept blading. This paper extends the assessment of the payoff derived from forward sweep to multistage configurations. The experimental investigation compare two builds of an advanced two-stage fan configuration tested alternately with a radial and a forward swept stage 1 blade. In the two-stage evaluations, the testing was extended to include the effect on inlet flow distortion. While 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, this configuration did demonstrate superior front stage efficiency and tolerance to inlet distortion. Having obtained already low distortion sensitivity with the radial rotor 1 configuration relative to current production military fan standards, the sensitivity to inlet distortion was halved with the forward swept rotor 1 configuration. In the case of the 180-degree one-per-rev distortion pattern, the two-stage configuration was evaluated both with and without inlet guide vanes (IGVs). The presence of the inlet guide vanes had a profound impact in lowering the two-stage fan's sensitivity with inlet distortion.

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.

Forward Swept Rotor Studies in Multistage Fans With Inlet Distortion

2002 ◽  
Author(s):  
A. R. Wadia ◽  
P. N. Szucs ◽  
D. W. Crall ◽  
D. C. Rabe
Keyword(s):  
Pressure Ratio ◽  
Flow Distortion ◽  
Two Stage ◽  
Good Efficiency ◽  
Guide Vanes ◽  
Stall Margin ◽  
High Pressure Ratio ◽  
Inlet Distortion ◽  
Inlet Guide Vanes ◽  
Current Production

Previous experimental and analytical studies conducted to compare the performance of transonic swept rotors in single stage fans have demonstrated the potential of significant improvements in both efficiency and stall margin with forward swept blading. This paper extends the assessment of the payoff derived from forward sweep with respect to aerodynamic performance and stability to multistage configurations. The experimental investigation compares, on a back-to-back test basis, two builds of an advanced good efficiency, high pressure ratio, two-stage fan configuration tested alternately with a radial and a forward swept stage 1 blade. In the two-stage evaluations, the testing was extended to include the effect on inlet flow distortion. While 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, this configuration did demonstrate superior front stage efficiency and tolerance to inlet distortion. Having obtained an already low distortion sensitivity with the radial rotor 1 configuration relative to current production military fan standards, the sensitivity to inlet distortion was halved with the forward swept rotor 1 configuration. In the case of the 180-degree one-per-rev distortion pattern, the two-stage configuration was evaluated both with and without inlet guide vanes (IGVs). The presence of the inlet guide vanes had a profound impact in lowering the two stage fan’s sensitivity with inlet distortion.

A-10 Wing Leading Edge Effects on Engine Stability: Part 1 — Analysis and Evaluation of Wing Leading Edge Configurations

2015 ◽  
Author(s):  
Grant T. Patterson ◽  
Brian A. Binkley ◽  
Jerome C. Jenkins
Keyword(s):  
Pressure Ratio ◽  
Leading Edge ◽  
Vortex Generators ◽  
Analysis And Evaluation ◽  
Inlet Distortion ◽  
Engine Stability ◽  
Overall Performance ◽  
Tools And Techniques ◽  
Technical Basis ◽  
Wing Stall

The A-10 aircraft has fuselage mounted engines with inlets just above the rear of the wing. The A-10 employs a deployable slat system to delay wing stall directly in front of the engines. Wing stall can lead to high inlet distortion and ultimately engine stall for this aircraft. To enhance overall performance of the A-10 Close Air Support Aircraft, wing leading-edge designs that do not employ slats were considered. Fifteen potential wing leading-edge proposals including drooped wings, wings with fences, wings with vortex generators, an optimized slat and a specially designed wing were evaluated through test and analysis for replacing the A-10 slat system. The performance of the wing leading-edge candidates were characterized by their inlet engine distortion effect on the loss of stability pressure ratio (ΔPRS) on the TF-34 engine fan and compressor. The drooped wings or “droops” were designated by the amount of droop in a percent of chord. Droops tested were 3, 5, 7, 10, and 10-5% twisted (5% outboard, 10% inboard). The 7, 10, and 10-5% droops were tested with outboard fences. The 10% droop and designed wing were tested with vortex generators. The paper discusses the previous work and technical basis for selecting the wing leading edge candidates, the analysis tools and techniques, the test and analysis of the candidate configurations, the overall effectiveness of the best candidate.

scholarly journals Efficacy of Two-Stage Treatment of Onychomycosis Using a Long-Pulsed Nd:YAG 1064-nm Laser

2019 ◽  
Vol 2019 ◽  
pp. 1-6
Author(s):  
Shan Zhong ◽  
Guan-ting Lin ◽  
Jun-ying Zhao
Keyword(s):  
Laser Treatment ◽  
Clinical Efficacy ◽  
Nail Plate ◽  
Efficacy Rate ◽  
Treatment Schedule ◽  
Two Stage ◽  
Second Stage ◽  
Controlled Treatment ◽  
Stage 1

Objectives. Onychomycosis is a fungal infection of the nail bed, nail matrix, and nail plate. Lasers have recently been studied as future clinical options for onychomycosis. We sought to evaluate the efficacy of the long-pulsed Nd:YAG 1064-nm laser on onychomycosis using a two-stage treatment. Methods. A total of 100 affected nails (88 toenails and 12 fingernails from 22 patients) were treated using a long-pulsed Nd:YAG 1064-nm laser. The self-controlled treatment schedule included the first stage (1 session per week for 8 weeks) and the second stage (1 session every 4 weeks for 16 weeks). Patients were followed up with for 12 weeks after the last laser treatment. Direct microscopy, cultures, and clinical assessments were performed at various time points. Results. Of the nails studied, 84% were infected by Trichophyton rubrum, while others were coinfected with T. rubrum and other fugal strains. The mycological clearance rate and the clinical efficacy rate of the nails were 29% and 21% after the first stage, 69% and 35% after the second stage, and 67% and 39% during follow-up, respectively. The second stage of laser treatment significantly improved the mycological clearance and clinical efficacy. Conclusion. Long-pulsed Nd:YAG 1064-nm laser two-stage treatment was effective for onychomycosis, with significantly improved mycological clearance and clinical efficacy. This trial is registered with ChiCTR 1900021669.

Effect of Inlet Distortion Features on Transonic Fan Rotor Stall

2017 ◽  
Author(s):  
James H. Page ◽  
Paul Hield ◽  
Paul G. Tucker
Keyword(s):  
Static Pressure ◽  
Main Body ◽  
Jet Engine ◽  
Stall Margin ◽  
Inlet Distortion ◽  
Factors Influencing ◽  
Transonic Fan ◽  
Passenger Aircraft ◽  
New Evidence ◽  
Transonic Rotor

The effect of inlet distortion from curved intake ducts on jet engine fan stability is an important consideration for next generation passenger aircraft such as the boundary layer ingestion (BLI) Silent Aircraft. Highly complex inlet flows which occur can significantly affect fan stability. Future aircraft designs are likely to feature more severe inlet distortion, pressing the need to understand the important factors influencing design. This paper presents the findings from a large CFD investigation into which aspects of inlet distortion cause the most significant reductions in stall margin and, therefore, which flow patterns should be targeted by mitigating technology. The study considers the following aspects of distortion commonly observed in intakes: steady vortical distortion due to secondary flow, unsteady vortical distortion due to vortex shedding and mixing, static pressure distortion due to curved streamlines, and low momentum endwall flow due to thickened boundary layers or separation. Unsteady CFD was used to determine the stall points of a multipassage transonic rotor geometry with each of the inlet distortion patterns applied. Interesting new evidence is provided which suggests that low momentum flow in the tip region, rather than distortion in the main body of the flow, leads to damaging instability.

RECOGNITION OF PARTIAL PLANAR SHAPES IN LIMITED MEMORY ENVIRONMENTS

1990 ◽  
Vol 04 (04) ◽  
pp. 603-628 ◽  
Author(s):  
SANTANU CHAUDHURY ◽  
S. SUBRAMANIAN ◽  
GUTURU PARTHASARATHY
Keyword(s):  
Heuristic Search ◽  
Recognition Algorithm ◽  
Two Stage ◽  
Second Stage ◽  
Limited Memory ◽  
Occluded Objects ◽  
Planar Shapes ◽  
Problem Instances ◽  
Stage Matching ◽  
Partially Occluded

Industrial vision systems should be capable of recognising noisy objects, partially occluded objects and randomly located and/or oriented objects. This paper considers the problem of recognition of partially occluded planar shapes using contour segment-based features. None of the techniques suggested in the literature for solving the above problem guarantee reliable results for problem instances which require memory in excess of what is available. In this paper, a heuristic search-based recognition algorithm is presented, which guarantees reliable recognition results even when memory is limited. This algorithm identifies an object, the maximum portion of whose contour is visible in a conglomerate of objects. For increasing efficiency of the method, a two-stage recognition scheme has been designed. In the first phase, a relevant subset of the known model shapes is chosen and in the second stage, matching between the unknown shape and elements of the relevant subset is attempted using the above approach. The technique is general in the sense that it can be used with any kind of contour features. To evaluate the efficiency of the method, experimentation was carried out using polygonal approximations of the object contours. Results are cited for establishing the effectiveness of the approach.

Effect of inlet distortion on the performance of axial transonic contra-rotating compressor

2016 ◽  
Vol 232 (1) ◽  
pp. 42-54 ◽  
Author(s):  
Hanru Liu ◽  
Yangang Wang ◽  
Songchuan Xian ◽  
Wenbin Hu
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Inlet Distortion ◽  
Flow Mixing ◽  
Inflow Conditions ◽  
Total Pressure Ratio

The present paper numerically conducted full-annulus investigation on the effects of circumferential total pressure inlet distortion on the performance and flow field of the axial transonic counter-rotating compressor. Results reveal that the inlet distortion both deteriorates the performance of the upstream and downstream rotors resulting in reduction of total pressure ratio, efficiency and stall margin of the transonic contra-rotating compressor. Regarding the development of distortion inside compressor, the downstream rotor reinforces the air-flow mixing effects and, thus, attenuates the distortion intensity significantly. Under the distorted inflow conditions, the detached shockwave at the leading edge of downstream rotor interacts with the tip leakage flow and causes the blockage of the blades passage, which is one important reason for the transonic contra-rotating compressor stall.

Aerodynamic Design and Testing of an Axial Flow Compressor for the GT24/26 Gas Turbines

2013 ◽  
Author(s):  
Damir Novak ◽  
Michael Loetzerich ◽  
Matthias Boese
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Leading Edge ◽  
Advanced Technology ◽  
Aerodynamic Design ◽  
3D Design ◽  
Axial Flow Compressor ◽  
Stage Matching ◽  
Flow Compressor

A 22-stage axial flow compressor with a pressure ratio 35:1 has been designed, built and successfully tested for a heavy-duty gas turbine application. Advanced technology and aero engine design tools have been used. The compressor has been designed using an “arbitrary” airfoil blading including 3D design features, like leading edge re-camber, lean, sweep and flowpath contouring. The compressor performance and part load behavior have been improved by accurate stage matching based on whole compressor 3D analyses. The new compressor has been tested in a scaled down rig and validated in the Alstom Test Power Plant (ATPP).The compressor met all design objectives and demonstrated excellent performance. This paper describes the aerodynamic design and test results.

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