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.

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.

Redesign of a Compressor Stage for a High-Performance Electric Supercharger in a Heavily Downsized Engine

2017 ◽  
Vol 140 (4) ◽  
Author(s):  
Peng Wang ◽  
Mehrdad Zangeneh ◽  
Bryn Richards ◽  
Kevin Gray ◽  
James Tran ◽  
...  
Keyword(s):  
Design Method ◽  
Pressure Ratio ◽  
Compressor Stage ◽  
Impeller Blade ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
Surge Margin ◽  
Stable Operating ◽  
Inverse Design Method ◽  
Ported Shroud

Engine downsizing is a modern solution for the reduction of CO2 emissions from internal combustion engines. This technology has been gaining increasing attention from industry. In order to enable a downsized engine to operate properly at low speed conditions, it is essential to have a compressor stage with very good surge margin. The ported shroud, also known as the casing treatment, is a conventional way used in turbochargers to widen the working range. However, the ported shroud works effectively only at pressure ratios higher than 3:1. At lower pressure ratio, its advantages for surge margin enhancements are very limited. The variable inlet guide vanes are also a solution to this problem. By adjusting the setting angles of variable inlet guide vanes, it is possible to shift the compressor map toward the smaller flow rates. However, this would also undermine the stage efficiency, require extra space for installing the inlet guide vanes, and add costs. The best solution is therefore to improve the design of impeller blade itself to attain high aerodynamic performances and wide operating ranges. This paper reports a recent study of using inverse design method for the redesign of a centrifugal compressor stage used in an electric supercharger, including the impeller blade and volute. The main requirements were to substantially increase the stable operating range of the compressor in order to meet the demands of the downsized engine. The three-dimensional (3D) inverse design method was used to optimize the impeller geometry and achieve higher efficiency and stable operating range. The predicted performance map shows great advantages when compared with the existing design. To validate the computational fluid dynamics (CFD) results, this new compressor stage has also been prototyped and tested. It will be shown that the CFD predictions have very good agreement with experiments and the redesigned compressor stage has improved the pressure ratio, aerodynamic efficiency, choke, and surge margins considerably.

Design and Test of a Novel Highly-Loaded Compressor

2019 ◽  
Author(s):  
Chengwu Yang ◽  
Ge Han ◽  
Shengfeng Zhao ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
...  
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Flow Angle ◽  
Stall Margin ◽  
High Pressure Ratio ◽  
Numerical Predictions ◽  
Highly Loaded

Abstract The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

Design of the High Pressure Ratio Transonic 1-1/2 Stage Fan Demonstrator Hulda

2007 ◽  
Author(s):  
Hans Ma˚rtensson ◽  
Jo¨rgen Burman ◽  
Ulf Johansson
Keyword(s):  
Guide Vane ◽  
Pressure Ratio ◽  
Inlet Guide Vane ◽  
Demonstration Program ◽  
Good Efficiency ◽  
High Pressure Ratio ◽  
Mechanical Constraints ◽  
Computational Analyses ◽  
Variable Inlet Guide Vane ◽  
First Time

As the first design in a demonstration program for future fighter engine fans a 400 mm 1-1/2 stage fan has been designed and built. A new method including mechanical constraints for designing the blades and gas path is used for the first time on a new design. The approach closely integrates CFD for performance and FE methods for the structure. By this, advanced computational analyses affect the design from the early stages. A design that is successful in achieving good efficiency based on CFD as well as reasonable aeromechanical properties based on FE is derived. The fan incorporates a front frame (FF), variable inlet guide vane (VIGV), rotor 1 (R1) and stator 1 (S1).

Investigation of Self-Recycling-Casing-Treatment (SRCT) Influence on Stability of High Pressure Ratio Centrifugal Compressor With a Volute

2011 ◽  
Author(s):  
Mingyang Yang ◽  
Ricardo Martinez-Botas ◽  
Yangjun Zhang ◽  
Xinqian Zheng ◽  
Takahiro Bamba ◽  
...  
Keyword(s):  
High Pressure ◽  
Numerical Method ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Flow Distortion ◽  
Casing Treatment ◽  
High Pressure Ratio ◽  
The Stability

Large feasible operation range is a challenge for high pressure ratio centrifugal compressor of turbocharger in vehicle engine. Self-Recycling-Casing-Treatment (SRCT) is a widely used flow control method to enlarge the range for this kind of compressor. This paper investigates the influence of symmetrical/asymmetrical SRCT (ASRCT) on the stability of a high pressure ratio centrifugal compressor by experimental testing and numerical simulation. Firstly, the performance of the compressor with/without SRCT is tested is measured investigate the influence of flow distortion on the stability of compressor as well as the numerical method validation. Then detailed flow field investigation is conducted by experimental measurement and the numerical method to unveil the reasons for stability enhancement by symmetrical/asymmetrical SRCT. Results show that static pressure distortion at impeller outlet caused by the volute can make passages be confronted with flow distortion less stable than others because of their larger positive slope of T-S pressure ratio performance at small flow rate. SRCT can depress the flow distortion and reduce the slope by non-uniform recycling flow rate at impeller inlet. Moreover, ASRCT can redistribute the recycling flow in circumferential direction according to the asymmetric geometries. When the largest recycling flow rate is imposed on the passage near the distorted static pressure, the slope will be the most effectively reduced. Therefore, the stability is effectively enhanced by the optimized recycling flow device.

Aerodynamics of a Transonic Centrifugal Compressor Impeller

2002 ◽  
Author(s):  
Seiichi Ibaraki ◽  
Tetsuya Matsuo ◽  
Hiroshi Kuma ◽  
Kunio Sumida ◽  
Toru Suita
Keyword(s):  
Shock Wave ◽  
High Pressure ◽  
Centrifugal Compressor ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Flow Distortion ◽  
Flow Measurements ◽  
High Pressure Ratio ◽  
Compressor Impeller

High pressure ratio centrifugal compressors are applied to turbochargers and turboshaft engines because of their small dimensions, high efficiency and wide operating range. Such a high pressure ratio centrifugal compressor has a transonic inlet condition accompanied with a shock wave in the inducer portion. It is generally said that extra losses are generated by interaction of the shock wave and the boundary layers on the blade surface. To improve the performance of high pressure ratio centrifugal compressor it is necessary to understand the flow phenomena. Although some research works on transonic impeller flow have been published, some unknown flow physics are still remaining. The authors designed a transonic impeller, with an inlet Mach number is about 1.3, and conducted detailed flow measurements by using Laser Doppler Velocimetry (LDV). In the result the interaction between the shock wave and tip leakage vortex at the inducer and flow distortion at the downstream of inducer were observed. The interaction of the boundary layer and the shock wave was not observed. Also computational flow analysis were conducted and compared with experimental results.

A Numerical Study on the Effect of Inter-Stage Water Injection in an Opposed-Setting-Two-Stage Centrifugal Compressor

2020 ◽  
Author(s):  
Ming Ni ◽  
Shaojuan Geng ◽  
Zuojun Wei ◽  
Xiaohua Gan
Keyword(s):  
Centrifugal Compressor ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Water Injection ◽  
Two Stage ◽  
Stall Margin ◽  
Injection Parameter ◽  
Insufficient Time ◽  
The Stability ◽  
The Right

Abstract Inter-stage water injection is currently a common method to reduce the compression work. In this paper, we evaluate the effect of this approach on an opposed-setting-two-stage centrifugal compressor which is designed for a small gas turbine numerically. The Eulerian-Lagrangian method is adopted to calculate the continuous flow field and discrete particle terms with the k-ε turbulence model. 6 water injection conditions are conducted to compare with the dry air condition. The performance curve shift to the right-up side, i.e. higher total pressure ratio, higher efficiency and larger choking mass flow. However, the larger the performance improvement gained by inter-stage injection, the worse the stability will be. Under the micro injection ratio (about 0.1%), the stall margin of the compressor is improved. Additionally, with a proper injection parameter, the loading in the diffuser can be reduced and promise a wider stall margin. Nevertheless, there still exists droplets at the outlet of the high-pressure diffuser because of the insufficient time for evaporation.

Influence of Adjustable Inlet Guide Vanes on the Performance Characteristics of a Shrouded Centrifugal Compressor

2017 ◽  
Author(s):  
Yubao Tian ◽  
Yonghong Tang ◽  
Zhiheng Wang ◽  
Guang Xi
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Air Separation ◽  
Performance Characteristics ◽  
Flow Angle ◽  
Separation Unit ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
Lag Effect ◽  
Stagger Angle

A shrouded centrifugal compressor model stage used for 120,000 m3/h oxygen production air separation unit was designed and tested at several IGV stagger angles from −15° to +60° and machine Mach number from 0.97 to 0.5. Present research works aimed to assess the influence of the adjustable IGVs and the IGV modeling on the shrouded centrifugal compressor performance characteristics and inlet flow field and to explore the effect factors of the CFD prediction accuracy and compressor stability at different IGV stagger angles. The measured results show that the model stage with 0° IGV stagger angle yields almost the same stagnation pressure ratio performance as the stage-only model but at a lower peak isentropic efficiency. With an appropriate IGV stagger angle setting ranging from −15° to +30°, the compressor stability could be efficiently enhanced. Numerical studies indicate that a large IGV hub gap may lead to a significant lag effect on the flow angle generated by the inlet guide vanes when increasing the IGV stagger angle.

Numerical and Experimental Investigations of Clocking Effect in a Two-Stage Compressor With πC = 3.7

2015 ◽  
Author(s):  
Victor Mileshin ◽  
Yaroslav Druzhinin ◽  
Alexander Stepanov ◽  
Nikolay Savin
Keyword(s):  
Experimental Data ◽  
Pressure Ratio ◽  
Pressure Fluctuations ◽  
Angular Position ◽  
Tip Clearance ◽  
Test Facility ◽  
Experimental Investigations ◽  
Aerodynamic Load ◽  
Two Stage ◽  
Stall Margin

In recent years, a number of studies in Russia and abroad was completed with the aim of decreasing pressure fluctuations and losses in blade cascades by controlling the unsteady interactions of blade rows (known as “clocking effect”) [1–4]. Tests of individual stages demonstrated that the clocking effect is responsible for 1.5–2.0% in efficiency and 50% in pressure fluctuations [5]. This paper presents the results of experimental and theoretical studies of the clocking effect on gas-dynamic characteristics of a high-loaded two-stage compressor simulating the first two stages of HPC for an advanced engine. The compressor is designed with the help of up-to-date 1D, 2D and 3D direct and inverse problem solutions and distinguished by high aerodynamic loads of stages with πk=3.7 total pressure ratio, 17% stall margin and 88% adiabatic efficiency at Ncor=88% rotational speed that was demonstrated experimentally [6]. The compressor was tested at CIAM’s C-3 test facility in the assembly with d=0.5, 0.75, 1.0-mm tip clearance in both rotors (relative clearance in first stage 4.6·10−3; 6.9·10−3; 9.2·10−3 and relative clearance in second stage 9.1·10−3; 13.7·10−3; 18.3·10−3). When tested, clocking effects were checked up for separate and simultaneous changes in clocking positions of stator and rotor blade rows. Indications of a blade tip-timing system and pressure pulsation sensors were used as experimental data. Earlier, it was shown that physics of the rotor clocking is a wake interaction which modifies the behavior of a boundary layer in Rotor 2 blades. This work studies the mechanism of rotor clocking in combination with changes in angular position of Rotor 2 blades due to interactions with Rotor 1 wakes. Tests showed that changes in the clocking position of the rotor with a multiple number of Rotor 1 and Rotor 2 blades affected the static position of Rotor 2 blades causing re-position of the blades depending on the rotor clocking-position. To confirm this result, 3D unsteady aerodynamic calculation was completed with the help of NUMECA software package simulating one of the test points. This work presents the calculated and experimental data showing that vortex wakes from Rotor 1 blades extend downstream, reach Rotor 2 and cause a variable aerodynamic load and a variable blade pitch.

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.

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