Acoustic Minimization of Rotor-Stator Interaction Noise of Axial Fans by Trailing-Edge Blowing

2014 ◽  
Author(s):  
Michael Kohlhaas ◽  
Thomas H. Carolus
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Rotor Blades ◽  
The Novel ◽  
Optimization Strategy ◽  
Trailing Edge ◽  
Rotor Stator Interaction ◽  
Axial Fans ◽  
Novel Strategy

The work deals with tonal noise of axial turbo fan stages by rotor-stator interaction. The objective of this study is the reduction of this noise by injection of a secondary mass flow at the rotor blades’ trailing edge (trailing edge blowing, TEB). A literature survey suggested that at least the tonal rotor/stator interaction sound can be reduced considerably by perfect blade wake filling which usually is achieved by blowing air through slots in the blade trailing edge region. Our own studies proved that this traditional trailing edge blowing strategy is problematic. Hence, in this paper a novel strategy of trailing edge blowing is described. The key idea is the combination of experimental trailing edge blowing with an evolutionary optimization algorithm. Aiming directly at a minimum of far field sound pressure level we identified spanwise trailing edge flowing distributions that reduced the fundamental tone at blade passing frequency (BPF) by 1.4 dB, and at its first harmonic by 21.4 dB (the latter corresponding to a complete elimination). The required blowing mass flow rate is 2% of the mass flow rate through the stage and hence relatively moderate. 3D hot wire measurement revealed that the acoustically relevant upwash velocity fluctuations as seen by the stator are not only caused by the wakes from the rotor blades but also by vortex structures at the rotor’s hub and tip. Traditional trailing edge blowing primarily aims at eliminating the rotor blade wakes. By applying the novel optimization strategy even some (but not all) of those secondary flow structures could be put out of action which eventually led to the reduction of the spectral components observed.

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.

Design of a Highly Efficient Low-Noise Fan for Ultra-High Bypass Engines

2006 ◽  
Author(s):  
Burak Kaplan ◽  
Eberhard Nicke ◽  
Christian Voss
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Low Noise ◽  
Test Rig ◽  
Noise Emission ◽  
Rotor Stator Interaction ◽  
Funded Project ◽  
Bypass Ratio

In order to achieve an environmental-friendly engine i.e. with more efficiency and less noise emission, a geared ultra-high bypass ratio fan test rig has been designed within the EU-funded project SILENCE®. Engine cycle requirements were untypical in terms of mass flow rate, pressure ratio and BPR. In order to reach the desired mass flow rate and simultaneously to avoid a strong interaction of the shocks with the boundary layer an S-shape leading edged rotor with forward sweep close to casing has been designed. Specific blade numbers for rotor and stator has been used to minimize the rotor-stator interaction noise. For the same purpose a backward swept bypass stator has been designed. There are two stators in the core duct in order to bring the flow to zero swirl which is a necessity for test rig measurements. The main design loop includes blade shape and the flow path optimization as well as the computation of stress distribution in all blades and the rotor disc. The fan is being manufactured from titanium because of its specific aeroelastic properties and stators are made of steel. The rig is scheduled to be tested in 2006 for its aerodynamic and aeroacoustic performance.

2D Parametric Study Using CFD of a Circulation Control Inlet Guide Vane

2007 ◽  
Author(s):  
H. E. Hill ◽  
W. F. Ng ◽  
P. P. Vlachos ◽  
S. A. Guillot ◽  
D. Car
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Circulation Control ◽  
Trailing Edge ◽  
Pressure Losses ◽  
Supply Pressure ◽  
Edge Radius

Circulation control inlet guide vanes (IGVs) may provide significant benefits over current IGVs that employ mechanical means for flow turning. This paper presents the results of a two-dimensional computational study on a circulation control IGV that takes advantage of the Coanda effect for flow vectoring. The IGV in this study is an uncambered airfoil that alters circulation around itself by means of a Coanda jet that exhausts along the IGV’s trailing edge surface. The IGV is designed for an axial inlet flow at a Mach number of 0.54 and an exit flow angle of 11 degrees. These conditions were selected to match the operating conditions of the 90% span section of the IGV of the TESCOM compressor rig at the Compressor Aero Research Laboratory (CARL) located at Wright-Patterson AFB, the hardware that is being used as the baseline in this study. The goal of the optimization was to determine the optimal jet height, trailing edge radius, and supply pressure that would meet the design criteria while minimizing the mass flow rate and pressure losses. The optimal geometry that was able to meet the design requirements had a jet height of h/Cn = 0.0057 and a trailing edge Radius R/Cn = 0.16. This geometry needed a jet to inflow total pressure ratio of 1.8 to meet the exit turning angle requirement. At this supply pressure ratio the mass flow rate required by the flow control system was 0.71 percent of the total mass flow rate through the engine. The optimal circulation control IGV had slightly lower pressure losses when compared with a reference cambered IGV.

Efficiency Improvement of Industrial Fans if Operating at Off-Design Modes

2019 ◽  
Vol 7 (3) ◽  
pp. 43-51
Author(s):  
Глеб Замолодчиков ◽  
Gleb Zamolodchikov ◽  
Р. Тумашев ◽  
R. Tumashev ◽  
Н. Щеголев ◽  
...  
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
High Efficiency ◽  
Operating Mode ◽  
Rotor Blades ◽  
Aerodynamic Load ◽  
Axial Fan ◽  
Performance Factor ◽  
Wide Range

This paper’s aim is enhancement of efficiency for fans adjusting by turn of rotor blades. A high load axial fan and a fan with decreased rotor’s pitch chord ratio by reduction of blades number were investigated. Have been performed tests of the fan with design characteristics as follows: theoretical head coefficient Ht = 0,3, mass flow rate Ca = 0,4, hub’s relative diameter ν = 0.6, and with blades, graded on the law of permanent circulation. The area of effective adjustment was estimated by the performance factor value η* ≥ 0,8. When changing the stagger angles in a wide range from 26° to 70°, the area of highly economical work was in variation ranges 0,26–0,78 for the mass flow rate Ca , and 0,24–0,5 for the theoretical head coefficient Ht accordingly. Tests of fans with a reduced blades number in the rotor (12 instead of 16 for the original fan) has showed that under the same stagger angles the fan’s high-efficiency operating mode is approximately in the same range of Ca variation at slightly reduced values of theoretical head coefficient. Maximal performance factor has increased on 2.5%. Decreasing the number of rotary blades, simplifying the turning mechanism and reducing the weight are possible in the design of fans with increased values of aerodynamic load coefficients.

Numerical and Experimental Investigation of Turning Flow Effects on Innovative Pin Fin Arrangements for Trailing Edge Cooling Configurations

2010 ◽  
Author(s):  
C. Bianchini ◽  
B. Facchini ◽  
F. Simonetti ◽  
L. Tarchi ◽  
S. Zecchi
Keyword(s):  
Heat Transfer ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Surface Flow ◽  
Point Of View ◽  
Experimental Point ◽  
Outlet Section ◽  
Trailing Edge ◽  
Tip Mass

The effect of the array configuration of circular pin-fins is investigated from a numerical and experimental point of view reproducing a typical cooling scheme of a real high pressure aeroengine blade. The airstream enters the domain of interest radially from the hub inlet and exits axially from the trailing edge (TE) outlet section. More than one hundred turbulators are inserted in the wedge shaped TE duct to enhance the heat transfer: a reference array implementing 7 rows of staggered pins is compared with an innovative pentagonal arrangement. Investigations were made considering real engine flow conditions: both numerical calculations and experimental measurements were performed fixing Re = 18000 and Ma = 0.3 in the TE throat section. The effect of the tip mass flow rate was also taken into account, investigating 0% and 25% of the TE mass flow rate. The experimental activity was aimed at obtaining detailed heat transfer coefficient maps over the internal pressure side (PS) surface by means of the transient technique with thermochromic liquid crystals. Particle Image Velocimetry measurements were performed and surface flow visualizations were made by means of the oil & dye technique on the PS surface. Steady-state RANS simulations were performed with two different CFD codes: the commercial software Ansys CFX® 11.0 and an in-house solver based on the opensource toolbox OpenFOAM®, to compare the performance and predictive capabilities. Turbulence was modeled by means of the k–ω SST model with an hybrid near wall treatment allowing strong clustering of the wall of interest as well as quite coarse refinement on the other viscous surfaces.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade: Analysis of the Adiabatic Film Cooling Effectiveness

2009 ◽  
Author(s):  
Axel Dannhauer
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Pressure Side ◽  
Cooling Efficiency ◽  
Ir Thermography ◽  
Trailing Edge ◽  
Film Cooling Efficiency

Within a European research project experimental studies were performed concerning the determination of the film cooling efficiency on the pressure side of trailing edges of high pressure turbine blades. The experiments were carried out at the linear cascade wind tunnel (EGG) of the German Aerospace Center (DLR), Go¨ttingen. The thermodynamic investigations were performed using the same cascade geometries and trailing edge configurations as for the aerodynamic measurements. Two different trailing edge geometries with coolant ejection were investigated. The first configuration was equipped with a pressure side cutback while for the second configuration the pressure side film cooling was realized by a row of cylindrical holes. The determination of the surface temperatures was done by using a combination of IR-thermography and thermocouples. Preliminary studies showed the feasibility to use metallic surfaces of the suction side of the adjacent blade as a mirror for IR-thermography. Thus it is possible to observe the pressure side near the trailing edge of interest by means of an infrared camera. The camera was mounted outside of the cascade’s free stream ensuring no influence to the aerodynamic boundary conditions. Up to seven flush mounted thermocouples on each side of the trailing edge were used for an in-situ calibration of the infrared pictures and thermal loss calculations. The distributions and averaged values of the film cooling efficiency are in agreement with aerodynamic measurements [9]. The results for the cutback configuration with 0.5% mass flow rate ejected show an accumulation of coolant just behind the coolant slot which is caused by a vortex in the dead region of the cutback. In case of 1.0% mass flow rate a refilling of this region with coolant is indicated. For higher mass flow rates the distributions of the film cooling efficiency looses it’s homogeneity due to flow separations on some ribs of the pin fin array inside of the slot. For the configuration with pressure side bleeding the best coverage could be obtained applying 1.0% mass flow rate.

A detailed investigation of a variable nozzle turbine with novel forepart rotation guide vane

2018 ◽  
Vol 233 (4) ◽  
pp. 994-1007
Author(s):  
Dengfeng Yang ◽  
Ce Yang ◽  
Dazhong Lao ◽  
Tao Zeng
Keyword(s):  
Flow Rate ◽  
Guide Vane ◽  
Forced Response ◽  
Rotor Blades ◽  
Leakage Flow ◽  
The Novel ◽  
Leakage Flow Rate ◽  
Rotor Stator Interaction ◽  
Stage Performance ◽  
Work First

One of the disadvantages of a variable nozzle turbine in practical application is the stage performance degradation due to nozzle endwall leakage flow at small nozzle openings. Aiming at restricting the nozzle leakage flow rate to improve turbine stage performance, a novel forepart rotation guide vane has been proposed and numerically studied in present work. First, the numerical results of baseline turbine were validated by experimental data to ensure the accuracy of numerical methods. Then steady and unsteady simulations were performed on both baseline and forepart rotation guide vane turbines to demonstrate the effectiveness of the novel vane and to study the characteristics of nozzle leakage flow, respectively. Results indicate that there is up to 13.5% peak efficiency improvement that has been achieved at 10% nozzle opening with the forepart rotation guide vane design; besides, rotor–stator interaction for forepart rotation guide vane is also mitigated due to the reduced nozzle leakage flow rate, thus the intensity of loading fluctuation on rotor blades is weakened significantly, which is beneficial to improve rotor blade forced response.

scholarly journals Experimental Analysis of the Global Performance and the Flow Through a High-Bypass Turbofan in Windmilling Conditions

2015 ◽  
Vol 137 (5) ◽  
Author(s):  
Nicolás García Rosa ◽  
Guillaume Dufour ◽  
Roger Barènes ◽  
Gérard Lavergne
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Loss ◽  
Severe Case ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Radial Profiles ◽  
Flow Through ◽  
Resistive Loads

A detailed study of the air flow through the fan stage of a high-bypass, geared turbofan in windmilling conditions is proposed, to address the key performance issues of this severe case of off-design operation. Experiments are conducted in the turbofan test rig of ISAE, specifically suited to reproduce windmilling operation in an ambient ground setup. The engine is equipped with conventional measurements and radial profiles of flow quantities are measured using directional five-hole probes to characterize the flow across the fan stage and derive windmilling performance parameters. These results bring experimental evidence of the findings of the literature that both the fan rotor and stator operate under severe off-design angle-of-attack, leading to flow separation and stagnation pressure loss. The fan rotor operates in a mixed fashion: spanwise, the inner sections of the rotor blades add work to the flow while the outer sections extract work and generate a pressure loss. The overall work is negative, revealing the resistive loads on the fan, caused by the bearing friction and work exchange in the different components of the fan shaft. The parametric study shows that the fan rotational speed is proportional to the mass flow rate, but the fan rotor inlet and outlet relative flow angles, as well as the fan load profile, remain constant, for different values of mass flow rate. Estimations of engine bypass ratio have been done, yielding values higher than six times the design value. The comprehensive database that was built will allow the validation of 3D Reynolds-averaged Navier–Stokes (RANS) simulations to provide a better understanding of the internal losses in windmilling conditions.

Experimental/Numerical Investigation on the Effects of Trailing-Edge Cooling Hole Blockage on Heat Transfer in a Trailing-Edge Cooling Channel

2013 ◽  
Author(s):  
M. E. Taslim ◽  
X. Huang
Keyword(s):  
Heat Transfer ◽  
Mass Flow Rate ◽  
Turbulence Model ◽  
Mass Flow ◽  
Flow Rate ◽  
Transfer Coefficients ◽  
Trailing Edge ◽  
Edge Channel ◽  
Heat Transfer Coefficients ◽  
Near Wall

Hot and harsh environments, sometimes experienced by gas turbine airfoils, can create undesirable effects such as clogging of the cooling holes. Clogging of the cooling holes along the trailing edge of an airfoil on the tip side and its effects on the heat transfer coefficients in the cooling cavity around the clogged holes is the main focus of this investigation. Local and average heat transfer coefficients were measured in a test section simulating a rib-roughened trailing edge cooling cavity of a turbine airfoil. The rig was made up of two adjacent channels, each with a trapezoidal cross sectional area. The first channel supplied the cooling air to the trailing-edge channel through a row of racetrack-shaped slots on the partition wall between the two channels. Eleven cross-over jets, issued from these slots entered the trailing-edge channel, impinged on eleven radial ribs and exited from a second row of race-track shaped slots on the opposite wall that simulated the cooling holes along the trailing edge of the airfoil. Tests were run for the baseline case with all exit holes open and for cases in which 2, 3 and 4 exit holes on the airfoil tip side were clogged. All tests were run for two cross-over jet angles. The first set of tests were run for zero angle between the jet axis and the trailing-edge channel centerline. The jets were then tilted towards the ribs by five degrees. Results of the two set of tests for a range of jet Reynolds number from 10,000 to 35,000 were compared. The numerical models contained the entire trailing-edge and supply channels with all slots and ribs to simulate exactly the tested geometries. They were meshed with all-hexa structured mesh of high near-wall concentration. A pressure-correction based, multi-block, multi-grid, unstructured/adaptive commercial software was used in this investigation. The realizable k – ε turbulence model in combination with enhanced wall treatment approach for the near wall regions were used for turbulence closure. Boundary conditions identical to those of the experiments were applied and several turbulence model results were compared. The numerical analyses also provided the share of each cross-over and each exit hole from the total flow for different geometries. The major conclusions of this study were: a) Clogging of the exit holes near the airfoil tip alters the distribution of the coolant mass flow rate through the crossover holes and changes the flow structure. Depending on the number of clogged exit holes (from 3 to 6, out of 12), the tip-end crossover hole experienced from 35% to 49% reductions in its mass flow rate while the root-end crossover hole, under the same conditions, experienced an increase of the same magnitude in its mass flow rate, b) up to 64% reduction in heat transfer coefficients on the tip-end surface areas around the clogged holes were observed which might have devastating effects on the airfoil life. At the same time, a gain in heat transfer coefficient of up 40% was observed around the root-end due to increased crossover flows, c) Numerical heat transfer results with the use of the realizable k – ε turbulence model in combination with enhanced wall treatment approach for the near wall regions were generally in a reasonable agreement with the test results. The overall difference between the CFD and test results was about 10%.

scholarly journals An active control method for reducing sonic boom of supersonic aircraft

2021 ◽  
Vol 39 (3) ◽  
pp. 566-575
Author(s):  
Liuqing Ye ◽  
Zhengyin Ye ◽  
Boping Ma
Keyword(s):  
Mass Flow Rate ◽  
Active Control ◽  
Mass Flow ◽  
Flow Rate ◽  
Control Method ◽  
Leading Edge ◽  
Sonic Boom ◽  
Trailing Edge ◽  
Supersonic Aircraft ◽  
Active Control Method

Sonic boom reduction has been an urgent need to develop the future supersonic transport, because of the heavy damages of the noise pollution. This paper provides an active control method for the supersonic aircraft to reduce the sonic boom, wherein a suction slot near the leading edge and an injection slot near the trailing edge on the airfoil suction surface are opened, and the mass flow sucked in near the leading edge is equal to the mass flow injected near the trailing edge. The diamond and 566 airfoils are adopted as the baseline airfoil to verify the capability of the active control method, and the effects of the suction and injection location, the mass flow rate and the attack angle on the ground boom signature, the maximum overpressure, the drag coefficients and the ratio of lift to drag are studied in detail. The results show that the proposed active control method can significantly reduce the sonic boom, and the reduction of the sonic boom intensity is more sensitive to the injection near the trailing edge than the suction near the leading edge. Applying this active control method to the diamond (NACA0008) airfoil, when the mass flow rate is 6.5 kg/s(7.5 kg/s), the value of maximum positive overpressure is decreased by 12.87%(12.85%), the value of maximum negative overpressure is decreased by 33.83%(56.77%) and the drag coefficient is decreased by 9.50%(10.96%). It can be seen that the method proposed in this paper has great benefits in the reduction of sonic boom and provides a useful reference for designing a new generation of lower sonic boom supersonic aircraft.

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