Experimental investigations on the effects of divergent trailing edge and Gurney flaps on a supercritical airfoil

2007 ◽  
Vol 11 (2-3) ◽  
pp. 91-99 ◽  
Author(s):  
Y.C. Li ◽  
J.J. Wang ◽  
J. Hua
Keyword(s):  
Experimental Investigations ◽  
Trailing Edge ◽  
Supercritical Airfoil ◽  
Gurney Flaps

Experimental and numerical investigation into the wake of a supercritical airfoil in a compressible unsteady flow

2016 ◽  
Vol 231 (3) ◽  
pp. 419-434 ◽  
Author(s):  
B Beheshti Boroumand ◽  
M Mani ◽  
N Fallahpour
Keyword(s):  
Numerical Study ◽  
Oscillation Amplitude ◽  
Experimental Tests ◽  
Experimental Investigations ◽  
Hot Wire ◽  
Supercritical Airfoil ◽  
Signal Energy ◽  
Test Conditions ◽  
Power Spectral ◽  
Oscillation Frequencies

Experimental investigations were carried out to study the wake characteristics of a pitching supercritical airfoil at Mach number of 0.6. Flow field inside the wake was measured by a hot-wire anemometry at downstream distances from trailing edge of 0.25 and 0.5 times the chord length. All data were taken at mean incidence angles of 0 and 3°. The amplitudes of oscillation were 1 and 3° while the oscillation frequencies were 3 and 6 Hz. Output signals acquired from sensors were analyzed besides the effects of such parameters as frequency and oscillation amplitude. Moreover, a comprehensive numerical study was carried out for the same airfoil under similar experimental test conditions; then, the results of numerical simulations were analyzed and compared with those of experimental tests. Results of the present research could be summarized as: observation of hysteresis and how it is affected by frequency and amplitude variations, observation of increasing turbulence intensity by root mean square investigation and also increasing signal energy by means of power spectral density diagram for those sensors lied inside the wake, and finally, study of correlation between wake’s interior sensors and exterior ones.

scholarly journals Experimental and Numerical 3-D Investigations of the Flow Field Behind the Trailing Edge of a Cooled Turbine Guide Vane

1996 ◽  
Author(s):  
Dieter E. Bohn ◽  
Volker J. Becker ◽  
Klaus D. Behnke
Keyword(s):  
Mass Flow ◽  
Gas Flow ◽  
Guide Vane ◽  
Density Ratio ◽  
Axial Direction ◽  
Cold Test ◽  
Experimental Investigations ◽  
Trailing Edge ◽  
Hot Gas ◽  
Calculated Velocity

Experimental investigations have been carried out to analyze the hot gas flow and cooling gas flow in the direct vicinity of the trailing edge of a modern gas turbine vane, with cooling gas ejection through the trailing edge. The investigations were performed in one of the institute’s test turbines. The experimental set-up is designed to establish variable blowing ratios between cooling gas mass flow and hot gas mass flow. An alternative density ratio between hot gas and cooling gas was established by the use of CO2 instead of air as the cooling gas for cold test runs. The experimental investigations have been carried out for different radial positions. The measurement plane was located 0.2mm to 0.5mm downstream of the trailing edge (trailing edge width 1.6mm). Local regions of high anisotropic turbulence were detected in the mixing zone. For low blowing ratios, the trailing edge pressure side in the tip vicinity was found to be subjected to direct hot gas contact. The trailing edge ejection has an influence of at least one chordlength in axial direction. The experimental investigations were accompanied by 3-D Navier-Stokes computational simulations. The calculated velocity distributions were found to be quite consistent with the experimental results. The calculated flow angles differed locally from the measurements. This may be due to the turbulence model employed.

Experimental Investigations on Cooling Air Ejection at a Straight Turbine Cascade Using PIV and QLS

2011 ◽  
Author(s):  
Oliver Freund ◽  
Hans-Juergen Rehder ◽  
Philipp Schaefer ◽  
Ingo Roehle
Keyword(s):  
Measurement Techniques ◽  
Co2 Concentration ◽  
Light Sheet ◽  
Numerical Code ◽  
Experimental Investigations ◽  
Test Bed ◽  
Air Concentration ◽  
Trailing Edge ◽  
Air Ejection ◽  
Cooling Air

Due to the high turbine inlet temperatures in modern aircraft engines the adoption of several cooling techniques in the first turbine blade rows is state of the art. For this reason the influence of cooling air ejection on the main flow is in the interest of scientists. In this paper experimental and numerical investigations on the trailing edge cooling air ejection at a stator profile are presented. All measurements are performed at the Straight Cascade Wind tunnel Go¨ttingen. To verify the influence of the cooling air flow on the flow field, the velocity field is measured by Particle Image Velocimetry (PIV). The development of the cooling air concentration is analyzed by utilizing the Quantitative Light Sheet (QLS) technique. For validation purposes the QLS results are compared to CO2 concentration measurements. Both measurement techniques are in good agreement with each other. One of the most important advantages of PIV and QLS is the possibility of combining them at the same test bed due to the identical experimental setup. The experimental investigations are supported by numerical simulations based on the numerical code TRACE. Both the numerical results as well as the experimental results prove the reduction of the trailing edge shock by increasing the coolant mass flow ratio.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

2013 ◽  
Vol 444-445 ◽  
pp. 517-523
Author(s):  
Da Wei Liu ◽  
Xin Xu ◽  
Zhi Wei ◽  
De Hua Chen
Keyword(s):  
Reynolds Number ◽  
Wind Tunnel ◽  
Pressure Distribution ◽  
Low Reynolds Number ◽  
Trailing Edge ◽  
Supercritical Airfoil ◽  
Short Period ◽  
Wind Tunnel Data ◽  
Mach Numbers ◽  
Low Reynolds

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106to 30×106per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

scholarly journals Flow control over an airfoil using virtual Gurney flaps

2015 ◽  
Vol 767 ◽  
pp. 595-626 ◽  
Author(s):  
Li-Hao Feng ◽  
Kwing-So Choi ◽  
Jin-Jun Wang
Keyword(s):  
Flow Control ◽  
Shear Layer ◽  
Plasma Actuator ◽  
Recirculation Region ◽  
Dbd Plasma ◽  
Near Wake ◽  
Trailing Edge ◽  
Pressure Surface ◽  
Gurney Flaps ◽  
Gurney Flap

AbstractFlow control over a NACA 0012 airfoil is carried out using a dielectric barrier discharge (DBD) plasma actuator at the Reynolds number of 20 000. Here, the plasma actuator is placed over the pressure (lower) side of the airfoil near the trailing edge, which produces a wall jet against the free stream. This reverse flow creates a quasi-steady recirculation region, reducing the velocity over the pressure side of the airfoil. On the other hand, the air over the suction (upper) side of the airfoil is drawn by the recirculation, increasing its velocity. Measured phase-averaged vorticity and velocity fields also indicate that the recirculation region created by the plasma actuator over the pressure surface modifies the near-wake dynamics. These flow modifications around the airfoil lead to an increase in the lift coefficient, which is similar to the effect of a mechanical Gurney flap. This configuration of DBD plasma actuators, which is investigated for the first time in this study, is therefore called a virtual Gurney flap. The purpose of this investigation is to understand the mechanism of lift enhancement by virtual Gurney flaps by carefully studying the global flow behaviour over the airfoil. First, the recirculation region draws the air from the suction surface around the trailing edge. The upper shear layer then interacts with the opposite-signed shear layer from the pressure surface, creating a stronger vortex shedding from the airfoil. Secondly, the recirculation region created by a DBD plasma actuator over the pressure surface displaces the positive shear layer away from the airfoil, thereby shifting the near-wake region downwards. The virtual Gurney flap also changes the dynamics of laminar separation bubbles and associated vortical structures by accelerating laminar-to-turbulent transition through the Kelvin–Helmholtz instability mechanism. In particular, the separation point and the start of transition are advanced. The reattachment point also moves upstream with plasma control, although it is slightly delayed at a large angle of attack.

scholarly journals Prediction of noise from serrated trailing edges

2016 ◽  
Vol 793 ◽  
pp. 556-588 ◽  
Author(s):  
B. Lyu ◽  
M. Azarpeyvand ◽  
S. Sinayoko
Keyword(s):  
Noise Reduction ◽  
Analytical Model ◽  
Leading Edge ◽  
Experimental Investigations ◽  
Destructive Interference ◽  
Trailing Edge ◽  
Trailing Edges ◽  
Mach Numbers ◽  
Trailing Edge Serrations ◽  
High Mach

A new analytical model is developed for the prediction of noise from serrated trailing edges. The model generalizes Amiet’s trailing-edge noise theory to sawtooth trailing edges, resulting in a complicated partial differential equation. The equation is then solved by means of a Fourier expansion technique combined with an iterative procedure. The solution is validated through comparison with the finite element method for a variety of serrations at different Mach numbers. The results obtained using the new model predict noise reduction of up to 10 dB at 90$^{\circ }$ above the trailing edge, which is more realistic than predictions based on Howe’s model and also more consistent with experimental observations. A thorough analytical and numerical analysis of the physical mechanism is carried out and suggests that the noise reduction due to serration originates primarily from interference effects near the trailing edge. A closer inspection of the proposed mathematical model has led to the development of two criteria for the effectiveness of the trailing-edge serrations, consistent but more general than those proposed by Howe. While experimental investigations often focus on noise reduction at 90$^{\circ }$ above the trailing edge, the new analytical model shows that the destructive interference scattering effects due to the serrations cause significant noise reduction at large polar angles, near the leading edge. It has also been observed that serrations can significantly change the directivity characteristics of the aerofoil at high frequencies and even lead to noise increase at high Mach numbers.

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