The Analysis about External Compressible Flow around an Aircraft Wing

2014 ◽  
Vol 1044-1045 ◽  
pp. 654-658
Author(s):  
Wei Long ◽  
Zai Shuai Ling ◽  
Zhen Dang
Keyword(s):  
Flow Field ◽  
Velocity Vector ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Flow Simulation ◽  
Aircraft Wing ◽  
Wing Model ◽  
Fluent Simulation ◽  
Lift And Drag ◽  
Variation Tendency

The Steady flow simulation to selected the delta wing model for different angles of attack in the Maher number.The law of flow field changes with the angle of attack is gotten.Through the FLUENT simulation,The variation tendency of coefficient of lift and drag in the different angle of attack is gotten.Further reveals the change rule of Maher number, pressure, velocity and other parameters in the different angle of attack.With increasing angle of attack, Maher number distribution is sparse of the same position increases and the greater numerical.the distribution of velocity vector is sparse of the same position increases and the greater numerical.the pressure distribution is sparse of the same position increases and the greater numerical.

scholarly journals Effect of vortex generator on the flow field over a conventional delta wing in subsonic flow condition at higher angles of attack

2021 ◽  
Vol 49 (2) ◽  
pp. 395-400
Author(s):  
Manthan Patil ◽  
Rajesh Gawade ◽  
Shubham Potdar ◽  
Khushabu Nadaf ◽  
Sanoj Suresh ◽  
...  
Keyword(s):  
Flow Field ◽  
Flow Visualization ◽  
Drag Coefficient ◽  
Subsonic Flow ◽  
Delta Wing ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Vortex Generator ◽  
Oil Flow ◽  
Oil Flow Visualization

Flow over a conventional delta wing has been studied experimentally at a subsonic flow of 20 m/sec and the flow field developed at higher angle of attack varying from 10° to 20° has been captured. A vortex generator is mounted on the leeward surface of the delta wing and its effect on the flow field is studied. The set of wing tip vortices generated over the delta wing is captured by the oil flow visualization and the streamline over the delta wing surface captured with and without a vortex generator are compared. Based on the qualitative results, the effect of the vortex generator on the lift coefficient is anticipated. Further, force measurement is carried out to quantitatively analyze the effect of vortex generator on the lift and drag coefficient experienced by the delta wing and justify the anticipation made out of the qualitative oil flow visualization tests. In the present study, the effect of mounting of a vortex generator is found to be minimal on the lift coefficient experienced by the delta wing. However, a significant reduction in the drag coefficient with increase in angle of attack was observed by mounting a typical vortex generator.

scholarly journals Analysis on Flow Field Characteristics of Airfoil in Pitching Motion

2021 ◽  
Vol 2076 (1) ◽  
pp. 012069
Author(s):  
Rui Yin ◽  
Jing Huang ◽  
Zhi-Yuan He
Keyword(s):  
Flow Field ◽  
Drag Coefficient ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Decisive Role ◽  
Aerodynamic Characteristics ◽  
Pitching Motion ◽  
Maximum Value ◽  
Flow Field Characteristics ◽  
Lift And Drag

Abstract Based on CFD, the flow field characteristics of NACA4412 airfoil are analyzed under pitching motion, and its aerodynamic characteristics are interpreted. The results show that streamline changes on the upper surface of the airfoil play a decisive role in the aerodynamic characteristics. The interaction between the vortex leads to fluctuations in the lift and drag coefficients. Under a big angle of attack, the secondary trailing vortex on the upper surface of the airfoil adheres to the trailing edge of the airfoil, resulting in an increased drag coefficient. Under a small angle of attack, the secondary trailing vortex can break away from the airfoil. The lift coefficient reaches the maximum value of 2.961 before the airfoil is turned upside down, and the drag coefficient reaches the maximum value of 1.515 after the airfoil is turned upside down, but the corresponding angles of attack of the two are equal.

A Study of the Flow Field Associated with a Steadily-Rolling Slender Delta Wing

1964 ◽  
Vol 68 (638) ◽  
pp. 106-110 ◽  
Author(s):  
J. K. Harvey
Keyword(s):  
Flow Field ◽  
Static Pressure ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Rolling Moment ◽  
Air Stream ◽  
Total Head ◽  
Theoretical Predictions ◽  
Attached Flow

SummaryIn this paper an experiment is described in which a detailed study was made of the flow field associated with a slender sharp-edged delta wing which was rolling steadily at zero angle of attack to an air stream. The investigation was made by performing two pressure surveys: first , one of the static pressure acting on the wing’s surface and second by measuring the total-head distribution in the neighbourhood of the wing. From the former the local rolling-moment coefficients, Clp, are evaluated and these are compared with the predictions for attached flow, thus assessing the contributions to the forces acting on the wing which arise as a consequence of the leading-edge separations. The second set of surveys is used to construct a picture of the flow-field details and this is compared with that known to occur on a similar wing when it is set at an angle of attack to the airstream. One interesting finding is that the secondary separation which appears to cause the discrepancy between the theoretical predictions and the measurements made on slender wings at incidence, is absent in this configuration and thus it is concluded that these data could be used for a more meaningful test of the theory.

Surface Modifications for Improved Maneuverability and Performance of an Aircraft

2011 ◽  
Author(s):  
Deepanshu Srivastav ◽  
K. N. Ponnani
Keyword(s):  
Surface Modifications ◽  
Boundary Layer Separation ◽  
Aerodynamic Characteristics ◽  
Aircraft Wing ◽  
Pressure Drag ◽  
Wing Model ◽  
Study Results ◽  
Stall Angle ◽  
And Performance ◽  
Lift And Drag

The work describes a comparative study of aerodynamic characteristics of an aircraft wing model with and without surface modifications to it. The surface modifications that are considered here are outward dimples on the wing model. In the present study, results of computational fluid dynamics (CFD) analysis are presented showing variance in lift and drag of modified wing models at different angle of attacks. Dimples on the surface aircraft wing model doesn’t affect much the pressure drag since it’s already aerodynamic in shape but it can affect the angle of stall. This project verifies if the dimples that reduce a golf ball’s drag, can also increase an airplane’s critical angle of stall. Dimples delay the boundary layer separation by creating more turbulence over the surface. The airfoil profile considered here is NACA-0018 with uniform cross-section throughout the length of airfoil. Subsonic flows are considered for the study. The CAD model is prepared in CatiaV5 R18 and simulations are carried out in Comsol 3.4 and Comsol 4.0. The overall aim of the study is improved maneuverability and performance of an aircraft. The results justify the increase in the overall lift and reduction in drag of the aircraft, also change of stall angle with different surface modifications on the wing model is observed.

On the Vortex Breakdown Phenomenon in High Angle of Attack Flows Over Delta Wing Geometries

2014 ◽  
Author(s):  
Eric D. Robertson ◽  
Varun Chitta ◽  
D. Keith Walters ◽  
Shanti Bhushan
Keyword(s):  
Vortex Breakdown ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Navier Stokes ◽  
Detached Eddy Simulation ◽  
High Angle ◽  
High Angle Of Attack ◽  
Eddy Simulation ◽  
Wing Model

Using computational methods, an investigation was performed on the physical mechanisms leading to vortex breakdown in high angle of attack flows over delta wing geometries. For this purpose, the Second International Vortex Flow Experiment (VFE-2) 65° sweep delta wing model was studied at a root chord Reynolds number (Recr) of 6 × 106 at various angles of attack. The open-source computational fluid dynamics (CFD) solver OpenFOAM was used in parallel with the commercial CFD solver ANSYS® FLUENT. For breadth, a variety of classic closure models were applied, including unsteady Reynolds-averaged Navier-Stokes (URANS) and detached eddy simulation (DES). Results for all cases are analyzed and flow features are identified and discussed. The results show the inception of a pair of leading edge vortices originating at the apex for all models used, and a region of steady vortical structures downstream in the URANS results. However, DES results show regions of massively separated helical flow which manifests after vortex breakdown. Analysis of turbulence quantities in the breakdown region gives further insight into the mechanisms leading to such phenomena.

scholarly journals Effect of Angle of Attack on Pressure and Lift Coefficient of ONERA OA206 Wing Model Using Computational Fluid Dynamics Method

2019 ◽  
Vol 2 (2) ◽  
pp. 81
Author(s):  
Resti Anggraeni
Keyword(s):  
Lift Coefficient ◽  
Pressure Coefficient ◽  
Angle Of Attack ◽  
Aircraft Wing ◽  
Wing Model ◽  
Pressure Contour ◽  
The Difference ◽  
The Right ◽  
Expansion Point ◽  
Dynamics Method

In this study, we computed the lift force of the aircraft with ONERA OA206 airfoil type. It was positioned at 0%, 25%, 50%, 75%, and 100% of the wingspan for Angle of Attack (AoA) variations of 0o, 4o, 8o, 12o, and 16o. The research was to determine the effect of AoA on pressure, pressure coefficient (Cp), and lift coefficient (CL) on the ONERA OA206 aircraft wing. It shows that the greater AoA on the result of the pressure contour causes the increase in the difference of span at AoA 0o to 16o t these are 0.25%; 0.26%; 0.43%; 0.52%; and 0.53%. Through the graph of the pressure coefficient (Cp) against x/c, it can be seen that the greater AoA, the expansion point, and the stagnation point will shift to the right with the direction of x/c. In addition, the Cp at the lower is greater than the upper of the airfoil. Based on the research results, it was found that CL at the position of 0% to 50% increased when given AoA from 0o to 12o (CL max) and decreased at AoA = 16o (stall). Meanwhile, CL at 75% to 100% increased when given AoA from 0o to 8o (CL max) and decreased at AoA = 12o (stall). With these results, it can be concluded that the maximum AoA that can be applied to the wing of the ONERA OA206 aircraft is 8o. The closer to the end position of the airfoil, the higher the CL measured.

scholarly journals Theoretical and Experimental of Evaluation of Flow Field Over a Double Delta Wing at Supersonic Speeds

2019 ◽  
Vol 8 (2S11) ◽  
pp. 2939-2943
Keyword(s):  
Stainless Steel ◽  
Flow Field ◽  
Numerical Simulations ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Experimental Studies ◽  
The Body ◽  
Experimental Investigations ◽  
Supersonic Speed ◽  
Double Delta Wing

Numerical and experimental investigations of flay over double delta wing at supersonic speed are investigated . For numerical simulations , the body geometry is generated using SolidWorks software , mesh is created using Gambit and solution is obtained using Fluent sofivvare . For the experimental studies , stainless steel model is fabricated and tested in the supersonic Mind tunnel at Mach 2.5 at 0 and 5 angle of attack . The results are presented in the present paper

scholarly journals Pulsed Blowing Interacting with a Leading-Edge Vortex

Aerospace ◽  
2020 ◽  
Vol 7 (1) ◽  
pp. 4
Author(s):  
Andrei Buzica ◽  
Christian Breitsamter
Keyword(s):  
Flow Field ◽  
Delta Wing ◽  
Active Flow Control ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Mean Flow ◽  
Sweep Angle ◽  
Leading Edge Vortex ◽  
Wing Model ◽  
Edge Vortex

Manipulation of vortex instabilities for aerodynamic performance increase is of great interest in numerous aeronautical applications. With increasing angle of attack, the leading-edge vortex of a semi-slender delta wing becomes unsteady and eventually collapses, endangering the flight stability. Hence, active flow control by pulsed blowing stabilizes the vortex system, enlarging the flight envelope for such wing configurations. The most beneficial outcome is the reattachment of the separated shear layer during post-stall, contributing to a lift increase of more than 50%. In contrast to high power consuming brute-force actuation, manipulating the flow instabilities offers a more efficient alternative for mean flow field control, which has direct repercussions on the aerodynamic characteristics. However, the flow mechanisms involving jet–vortex and vortex–vortex interactions and the disturbance convection through the flow field are little understood. This paper reports on the unsteady flow field above a generic half delta wing model with a 65 ° sweep angle and its response to periodic blowing. Numerical and experimental results are presented and discussed in a synergistic manner.

scholarly journals Lift and Drag of Non-conventional Wings at Subsonic Speeds and Zero Angle of Attack - An Experimental Investigation

2018 ◽  
Vol 152 ◽  
pp. 02017
Author(s):  
Abdulkareem Shafiq Mahdi Al-Obaidi ◽  
Ting Chern Wei
Keyword(s):  
Case Studies ◽  
Lift Coefficient ◽  
Angle Of Attack ◽  
Aerodynamic Performance ◽  
Notch Depth ◽  
Flight Condition ◽  
Wing Model ◽  
Subsonic Speed ◽  
Continue Research ◽  
Lift And Drag

Various non-conventional wing development shows potential in increasing the aerodynamic performance of airplanes. If the non-conventional wing only improves the aerodynamic performance by a small margin, conventional wing is still a better option for airline operators. This provides opportunity to continue research on non-conventional configurations that can greatly saves the fuel consumption. This research was conducted to examine the lift and drag of non-conventional wings at low subsonic speed and low angle of attack. Analytical method based on DATCOM was used to calculate the lift and drag coefficients of non-conventional cranked wing for comparison with experimental results obtained experimentally using Taylor’s wind tunnel (TWT). Experimental lift coefficient shows similar values with the analytical results but experimental drag coefficient had an average difference of 44%. The experimental setup and calibration of TWT were verified and further case studies on nonconventional wing model featuring trailing edge notches were carried out. Analysis of the results from case studies shows that generally the effect of varying the number of notches only had significant effect on drag reduction if the notch depth was higher. For flight condition that does not exceed 4° angle of attack, lower number of notches at higher notch depth had the best aerodynamic performance. On the other hand, for flight condition that requires cruise angle of attack that exceeds 4°, higher number of notches at higher notch depth had the best aerodynamic performance.

EFFECTS OF STEADY ANGLE OF ATTACK ON NONLINEAR GUST RESPONSE OF A DELTA WING MODEL

2002 ◽  
Vol 16 (8) ◽  
pp. 1093-1110 ◽  
Author(s):  
D.M. TANG ◽  
J.K. HENRY ◽  
E.H. DOWELL
Keyword(s):  
Delta Wing ◽  
Angle Of Attack ◽  
Wing Model ◽  
Gust Response
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