Numerical Study on Aeroelastic Instability for a Low-Speed Fan

2017 ◽  
Vol 139 (7) ◽  
Author(s):  
Kuen-Bae Lee ◽  
Mark Wilson ◽  
Mehdi Vahdati
Keyword(s):  
Performance Test ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Computational Domain ◽  
Low Speed ◽  
Acoustic Liners ◽  
Significant Difference ◽  
Fan Blade ◽  
Flutter Margin

Over recent years, engine designs have moved increasingly toward low specific thrust cycles to deliver significant specific fuel consumption (SFC) improvements. Such fan blades may be more prone to aerodynamic and aeroelastic instabilities than conventional fan blades. The aim of this paper is to analyze the flutter stability of a low-speed/low pressure ratio fan blade. By using a validated computational fluid dynamics (CFD) model (AU3D), three-dimensional unsteady simulations are performed for a modern low-speed fan rig for which extensive measured data are available. The computational domain contains a complete fan assembly with an intake duct and the downstream outlet guide vanes (OGVs), which is a whole low-pressure (LP) domain. Flutter simulations are conducted over a range of speeds to understand flutter characteristics of this blade. Only the first flap (1F) mode is considered in this work. Measured rig data obtained by using the same fan set but with two different lengths of the intake showed a significant difference in the flutter boundary for the two intakes. AU3D computations were performed for both intakes and were used to explain this difference between the two intakes, and showed that intake reflections play an important role in flutter of this blade. This observation indicates that the experiment with the long intake used for the performance test may be misleading for flutter. In the next phase of this work, two possible modifications for increasing the flutter margin of the fan blade were explored: changing the mode shape of the blade and using acoustic liners in the casing. The results show that it is possible to increase the flutter margin of the blade by either decreasing the ratio of the twisting to plunging motion in 1F mode or by introducing deep acoustic liners in the intake. The liners have to be deep enough to attenuate the flutter pressure waves and hence influence the stability. The results indicate the importance of reflection in flutter stability of the fan blade and clearly show that intake duct needs to be included in flutter study of any fan blade.

Numerical Study on Aeroelastic Instability for a Low Speed Fan

2016 ◽  
Author(s):  
Kuen-Bae Lee ◽  
Mark Wilson ◽  
Mehdi Vahdati
Keyword(s):  
Numerical Study ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Computational Domain ◽  
Low Speed ◽  
Acoustic Liners ◽  
Significant Difference ◽  
Flutter Boundary ◽  
Fan Blade ◽  
Flutter Margin

Over recent years engine designs have moved increasingly toward low specific thrust cycles to deliver significant specific fuel consumption (SFC) improvements. Such fan blades are more loaded than conventional fan blades and therefore can be more prone to aerodynamic and aeroelastic instabilities. The aim of this paper is to analyse the flutter stability of a low speed/low pressure ratio fan blade. By using a validated CFD model (AU3D), three dimensional unsteady simulations are performed for a modern low speed fan rig for which extensive measured data are available. The computational domain contains a complete fan assembly with an intake duct and the downstream OGVs (whole LP domain). Flutter simulations are conducted over a range of speeds to understand flutter characteristics of this blade. Only the 1F (first flap) mode is considered in this work. Measured rig data obtained by using the same fan set but with two different intakes showed a significant difference in the flutter boundary for the two intakes. AU3D computations were performed for both intakes and were used to explain this difference between the two intakes, and showed that intake reflections play an important role in flutter of this blade. In the next phase of this work, two possible modifications for increasing the flutter margin of the fan blade were explored: 1. Changing the mode shape of the blade 2. Using acoustic liners in the casing The results show that it is possible to increase the flutter margin of the blade by either decreasing the ratio of the twisting to plunging motion in 1F mode or by introducing deep acoustic liners in the intake. The liners have to be deep enough to attenuate the flutter pressure waves and hence influence the stability. The results indicate the importance of reflection in flutter stability of the fan blade, and clearly show that intake duct needs to be included in flutter study of any fan blade.

Numerical Analysis of Non-Radial Blading in a Low Speed and Low Pressure Turbine for Electric Turbocompounding Applications

2021 ◽  
Author(s):  
Eva Alvarez-Regueiro ◽  
Esperanza Barrera-Medrano ◽  
Ricardo Martinez-Botas ◽  
Srithar Rajoo
Keyword(s):  
Numerical Analysis ◽  
Numerical Simulations ◽  
Thermal Stresses ◽  
Waste Heat ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Turbine Rotor ◽  
Operating Conditions ◽  
Blade Angle ◽  
Low Speed

Abstract This paper presents a CFD-based numerical analysis on the potential benefits of non-radial blading turbine for low speed-low pressure applications. Electric turbocompounding is a waste heat recovery technology consisting of a turbine coupled to a generator that transforms the energy left over in the engine exhaust gases, which is typically found at low pressure, into electricity. Turbines designed to operate at low specific speed are ideal for these applications since the peak efficiency occurs at lower pressure ratios than conventional high speed turbines. The baseline design consisted of a vaneless radial fibre turbine, operating at 1.2 pressure ratio and 28,000rpm. Experimental low temperature tests were carried out with the baseline radial blading turbine at nominal, lower and higher pressure ratio operating conditions to validate numerical simulations. The baseline turbine incidence angle effect was studied and positive inlet blade angle impact was assessed in the current paper. Four different turbine rotor designs of 20, 30, 40 and 50° of positive inlet blade angle are presented, with the aim to reduce the losses associated to positive incidence, specially at midspan. The volute domain was included in all CFD calculations to take into account the volute-rotor interactions. The results obtained from numerical simulations of the modified designs were compared with those from the baseline turbine rotor at design and off-design conditions. Total-to-static efficiency improved in all the non-radial blading designs at all operating points considered, by maximum of 1.5% at design conditions and 5% at off-design conditions, particularly at low pressure ratio. As non-radial fibre blading may be susceptible to high centrifugal and thermal stresses, a structural analysis was performed to assess the feasibility of each design. Most of non-radial blading designs showed acceptable levels of stress and deformation.

scholarly journals Design Studies of Lift Fan Engines Suitable for Use in Civilian VTOL Aircraft

1972 ◽  
Author(s):  
R. J. Roelke ◽  
Steve Zigan
Keyword(s):  
Pressure Ratio ◽  
Low Pressure ◽  
Design Studies ◽  
Low Speed ◽  
Engine Design ◽  
Vtol Aircraft

Low pressure ratio fan engines are receiving increasing attention as a means to provide low speed lift for civilian VTOL transports. Two general types of fan lift engines that are being studied are integral fans and remote powered fans. Preliminary engine design studies of both types of lift fan systems have been made. This paper summarizes a portion of the results of the engine design studies, including the crucial engine requirements, and some of the characteristics of the emerging engine designs of each type.

scholarly journals The Aerodynamic Performance of an Over-the-Rotor Liner With Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

2013 ◽  
Author(s):  
Rick Bozak ◽  
Christopher Hughes ◽  
James Buckley
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Oscillating Flow ◽  
Acoustic Liners ◽  
Performance Loss ◽  
Acoustic Liner ◽  
Blade Tip ◽  
Fan Blade ◽  
The Impact

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1% which is within the repeatability of this experiment.

Numerical Study on the Effects of Blade Leading Edge Shape to the Performance of Supersonic Rotor

2003 ◽  
Author(s):  
Kicheol Park
Keyword(s):  
Rotational Speed ◽  
Pressure Loss ◽  
Total Pressure ◽  
Performance Test ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Manufacturing Cost ◽  
Blade Leading Edge

Recently, it is required to design a fan and compressor with higher stage pressure ratio while maintaining adiabatic efficiency high also. To increase the stage pressure ratio, blade rotational speed or diffusion factor should be increased. In the case of increased rotational speed, relative speed of flow at blade leading edge is well supersonic. With supersonic rotor blade, total pressure loss is mainly due to leading edge shock waves and the thickness should be thin enough to minimize this. As a result, the blade is like to be week in terms of mechanical strength and the manufacturing cost would be increased because high-precision NC machining is required. Furthermore, it is one of the biggest hurdles to maintain proper level of thickness while one making small stages. In this paper, aerodynamic performance of supersonic rotor blades with different leading edge thickness and shapes are calculated using the finite volume method. The effects of blade leading edge shape and thickness to the performance are investigated especially in terms of total pressure loss and the already known loss correlations of leading edge thickness are examined. Subsequently this will be verified by performance test on rig.

Numerical Investigations on Aerodynamic Design Criteria for Low Speed Mixed Flow Compressor

2021 ◽  
Author(s):  
Hemant Kumar ◽  
Chetan S. Mistry
Keyword(s):  
Flow Field ◽  
Flow Structure ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Cone Angle ◽  
Tip Clearance ◽  
Working Fluid ◽  
Mixed Flow ◽  
Low Speed ◽  
Flow Compressor

Abstract A surge in the small jet engine market due to aero-propulsion purposes generates a requirement to develop compact and robust high-performance compressors. Mixed flow compressors can provide a comparatively higher pressure ratio compared to axial compressors and have less frontal area than centrifugal compressors. Rapid progress in manufacturing and computational capabilities has resulted in the successful design of mixed flow compressors in recent decades. In the present study, the mixed flow compressor was designed to operate at 3,000 rpm with a small total-to-total pressure ratio of 1.03 and a mass flow rate = 1.98 kg/s to carry at low-speed testing for university-level research. Meanline design for the compressor with air as working fluid was done. The blade geometry was developed using commercial Ansys® Bladegen module. The flow domain mesh was generated by the TurboGrid module. Ansys CFX was used as a solver and post-processing tool for the present numerical study. The present work describes the detailed design procedure, overall performance, and flow field features of a low-speed mixed-flow compressor with the special requirement of axial flow exit. The parametric analysis was carried out on splitter blade placement, wrap angle (10°, 20°, 30°, and 50°), and exit cone angle (30°, 40°, 50°, 60°, and 65°), at constant tip clearance and keeping the other parameters constant to observe their effect on performance and flow structure. The use of splitter blades smoothen the flow structure along both stream-wise and span-wise direction, which minimizes flow the separation issue and thereby helping in extending the overall operating range. Comparing the flow field characteristic and performance of each parametric variable, the optimum range of design values is exhibited. The numerical observation and analysis done on parametric variations in this paper can be used for the design of such a future low-speed mixed flow compressor for different performance expectations and installation requirements.

Surface Roughness Impact on Low-Pressure Turbine Performance due to Operational Deterioration

2017 ◽  
Author(s):  
Andreas Kellersmann ◽  
Sarah Weiler ◽  
Christoph Bode ◽  
Jens Friedrichs ◽  
Jörn Städing ◽  
...  
Keyword(s):  
Surface Roughness ◽  
Laser Scanning ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Low Pressure ◽  
Laser Scanning Microscopy ◽  
Confocal Laser ◽  
Low Pressure Turbine ◽  
The Impact

The overall efficiency and operational behavior of aircraft engines are influenced by the surface finish of the airfoils. During operation, the surface roughness significantly increases due to erosion and deposition processes. The aim of this study is to analyze the influence of roughness on the aerodynamics of the low-pressure turbine of a mid-sized high bypass turbofan. In order to gain a better insight into the operational roughness structures, a sample of new, used, cleaned and reworked turbine blades and vanes are measured using the confocal laser scanning microscopy technique. The measurement results show local inhomogeneities. The roughness distributions measured are then converted into their equivalent sand grain roughness ks,eq to permit an evaluation of the impact on aerodynamic losses. The numerical study is performed using the CFD-solver TRACE which was validated before with existing data from Rig experiments. It is observed that the influence of the surface roughness on the turbine efficiency is significant at take-off but negligible at cruise. A detailed analysis on the aerodynamics at take-off shows that very rough airfoils lead to higher profile and secondary loss. Due to the higher disturbances present in flows circulating over rough walls, the transition occurs earlier and the momentum thickness increases in the turbulent boundary layer. The service-induced roughness structures cause an efficiency drop in the low pressure turbine of ηT = −0.16% compared to new parts. A gas path analysis showed that this results in an increased fuel flow of Δṁf = +0.06% and an exhaust gas temperature rise of ΔEGT = +1.2K for fixed engine pressure ratio which is equivalent to roughly 4 percent of the typical EGT margin of a fully refurbished engine. This result stresses the importance of roughness induced loss in low pressure turbines.

Surface Roughness Impact on Low-Pressure Turbine Performance Due to Operational Deterioration

2018 ◽  
Vol 140 (6) ◽  
Author(s):  
Andreas Kellersmann ◽  
Sarah Weiler ◽  
Christoph Bode ◽  
Jens Friedrichs ◽  
Jörn Städing ◽  
...  
Keyword(s):  
Surface Roughness ◽  
Laser Scanning ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Low Pressure ◽  
Laser Scanning Microscopy ◽  
Confocal Laser ◽  
Low Pressure Turbine ◽  
The Impact

The overall efficiency and operational behavior of aircraft engines are influenced by the surface finish of the airfoils. During operation, the surface roughness significantly increases due to erosion and deposition processes. The aim of this study is to analyze the influence of roughness on the aerodynamics of the low-pressure turbine (LPT) of a midsized high bypass turbofan. In order to gain a better insight into the operational roughness structures, a sample of new, used, cleaned, and reworked turbine blades and vanes are measured using the confocal laser scanning microscopy technique. The measurement results show local inhomogeneities. The roughness distributions measured are then converted into their equivalent sand grain roughness ks,eq to permit an evaluation of the impact on aerodynamic losses. The numerical study is performed using the computational fluid dynamics (CFD)-solver turbomachinery research aerodynamics computational environment (TRACE) which was validated before with the existing data from rig experiments. It is observed that the influence of the surface roughness on the turbine efficiency is significant at take-off but negligible at cruise. A detailed analysis on the aerodynamics at take-off shows that very rough airfoils lead to higher profile and secondary loss. Due to the higher disturbances present in flows circulating over rough walls, the transition occurs earlier, and the momentum thickness increases in the turbulent boundary layer. The service-induced roughness structures cause an efficiency drop in the LPT of ηT=−0.16% compared to new parts. A gas path analysis showed that this results in an increased fuel flow of Δm˙f=+0.06% and an exhaust gas temperature (EGT) rise of ΔEGT=+1.2K for fixed engine pressure ratio which is equivalent to roughly 4% of the typical EGT margin of a fully refurbished engine. This result stresses the importance of roughness-induced loss in LPTs.

scholarly journals Experimental and numerical investigation of heat dissipation from an electronic component in a closed enclosure

2018 ◽  
Vol 144 ◽  
pp. 04010
Author(s):  
Bobin Saji George ◽  
M. Ajmal ◽  
S. R. Deepu ◽  
M. Aswin ◽  
D. Ribin ◽  
...  
Keyword(s):  
Power Dissipation ◽  
Heat Dissipation ◽  
Numerical Study ◽  
Electronic Component ◽  
Thermal Design ◽  
Electronic Systems ◽  
Computational Domain ◽  
Computational Tool ◽  
Cycle Times ◽  
Design Cycle

Intensifying electronic component power dissipation levels, shortening product design cycle times, and greater than before requirement for more compact and reliable electronic systems with greater functionality, has heightened the need for thermal design tools that enable accurate solutions to be generated and quickly assessed. The present numerical study aims at developing a computational tool in OpenFOAM that can predict the heat dissipation rate and temperature profile of any electronic component in operation. A suitable computational domain with defined aspect ratio is chosen. For analyzing, “buoyant Boussinesq Simple Foam“ solver available with OpenFOAM is used. It was modified for adapting to the investigation with specified initial and boundary conditions. The experimental setup was made with the dimensions taken up for numerical study. Thermocouples were calibrated and placed in specified locations. For different heat input, the temperatures are noted down at steady state and compared with results from the numerical study.

Effect of Vaned Diffuser on a Modified Turbocharger Centrifugal Compressor Performance

2014 ◽  
Vol 663 ◽  
pp. 347-353
Author(s):  
Layth H. Jawad ◽  
Shahrir Abdullah ◽  
Zulkifli R. ◽  
Wan Mohd Faizal Wan Mahmood
Keyword(s):  
Centrifugal Compressor ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Aerodynamic Characteristics ◽  
Width Ratio ◽  
Outlet Section ◽  
Suction Surface ◽  
Vaned Diffuser ◽  
Minimum Width

A numerical study that was made in a three-dimensional flow, carried out in a modified centrifugal compressor, having vaned diffuser stage, used as an automotive turbo charger. In order to study the influence of vaned diffuser meridional outlet section with a different width ratio of the modified centrifugal compressor. Moreover, the performance of the centrifugal compressor was dependent on the proper matching between the compressor impeller along the vaned diffuser. The aerodynamic characteristics were compared under different meridional width ratio. In addition, the velocity vectors in diffuser flow passages, and the secondary flow in cross-section near the outlet of diffuser were analysed in detail under different meridional width ratio. Another aim of this research was to study and simulate the effect of vaned diffuser on the performance of a centrifugal compressor. The simulation was undertaken using commercial software so-called ANSYS CFX, to predict numerically the performance charachteristics. The results were generated from CFD and were analysed for better understanding of the fluid flow through centrifugal compressor stage and as a result of the minimum width ratio the flow in diffuser passage tends to be uniformity. Moreover, the backflow and vortex near the pressure surface disappear, and the vortex and detachment near the suction surface decrease. Conclusively, it was observed that the efficiency was increased and both the total pressure ratio and static pressure for minimum width ratio are increased.

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