Nonlinear lift increase at high angles of attack for double swept waverider

2020 ◽  
Vol 34 (14n16) ◽  
pp. 2040124
Author(s):  
Chuan-Zhen Liu ◽  
Peng Bai
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Lower Surface ◽  
Aerodynamic Performance ◽  
Sweep Angle ◽  
Vital Interest ◽  
High Angles Of Attack ◽  
Nonlinear Increase

The nonlinear increase of the lift of the double swept waverider at high angles of attack is of vital interest. The aerodynamic performance of the double swept waverider is calculated and compared with that of single swept waveriders. Results suggest that the lift nonlinearity of the double swept waverider is stronger than that of equal-planform-area single swept one, and the nonlinearity increases as Mach number increases. Some scholars have proposed the “vortex lift” to explain the nonlinear lift increase, but it is questionable as the main lift of the waverider comes from the lower surface rather than the upper surface. This paper proposes another explanation that the nonlinear lift increase is related to the attachment of shock wave, influenced by the leading-edge sweep angle. The shock wave is more inclined to attach under the lower surface with smaller swept than that of larger swept as angle of attack increases. When the shock wave attaches, the pressure increase via angle of attack is nonlinear, leading to the nonlinearity of lift increase.

Optimization of Aerodynamic Parameters of Cropped Delta Wing with Fence at Sonic Mach Number

2019 ◽  
Vol 16 (2) ◽  
pp. 403-409
Author(s):  
M. P. Arun ◽  
M. Satheesh ◽  
Edwin Raja J. Dhas
Keyword(s):  
Shock Wave ◽  
Boundary Condition ◽  
Mach Number ◽  
Sound Speed ◽  
Delta Wing ◽  
Leading Edge ◽  
Supersonic Speed ◽  
Sweep Angle ◽  
Ansys Fluent ◽  
Aerodynamic Parameters

Manufacturing and maintaining different aircraft fleet leads to various purposes, which consumes more money as well as man power. Solution to this, nations that are leading in the field of aeronautics are performing much research and development works on new aircraft designs that could do the operations those were done by varied aircrafts. The foremost benefit of this delta wing is, along the huge rearward sweep angle, the wing’s leading edge would not contact the boundary of shock wave. Further, the boundary is produced at the fuselage nose due to the speed of aircraft approaches and also goes beyond the transonic to supersonic speed. Further, rearward sweep angle greatly worse the airspeed: wings under normal condition to leading edge, so permits the aircraft to fly at great transonic, subsonic, or supersonic speed, whereas the over wing speed is kept to minimal range than that of the sound speed. The cropped delta wing with fence has analysed in three cases: Fences at 3/4th distance from the centre, with fences at half distance from the centre and with fences at the centre. Further, the delta wing that cropped is exported to ANSYS FLUENT V14.0 software and analysed by making the boundary condition settings like sonic Mach number of flow over wing along with the angle of attack.

scholarly journals Leading edge serrations for the reduction of aerofoil self-noise at low angle of attack, pre-stall and post-stall conditions

2021 ◽  
Vol 20 (1-2) ◽  
pp. 130-156
Author(s):  
Giovanni Lacagnina ◽  
Paruchuri Chaitanya ◽  
Jung-Hoon Kim ◽  
Tim Berk ◽  
Phillip Joseph ◽  
...  
Keyword(s):  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Design Guidelines ◽  
Noise Performance ◽  
Text Design ◽  
Reduction Mechanisms ◽  
Wide Range ◽  
Number Formula ◽  
High Angles Of Attack

This paper addresses the usefulness of leading edge serrations for reducing aerofoil self-noise over a wide range of angles of attack. Different serration geometries are studied over a range of Reynolds number [Formula: see text]. Design guidelines are proposed that permit noise reductions over most angles of attack. It is shown that serration geometries reduces the noise but adversely effect the aerodynamic performance suggesting that a trade-off should be sought between these two considerations. The self-noise performance of leading edge serrations has been shown to fall into three angle of attack (AoA) regimes: low angles where the flow is mostly attached, moderate angles where the flow is partially to fully separated, and high angles of attack where the flow is fully separated. Leading edge serrations have been demonstrated to be effective in reducing noise at low and high angles of attack but ineffective at moderate angles. The noise reduction mechanisms are explored in each of three angle regimes.

scholarly journals Numerical simulation of flapping airfoil with alula

2020 ◽  
Vol 12 ◽  
pp. 175682932097798
Author(s):  
Han Bao ◽  
Wenqing Yang ◽  
Dongfu Ma ◽  
Wenping Song ◽  
Bifeng Song
Keyword(s):  
Aerodynamic Force ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Geometric Parameters ◽  
Flight Performance ◽  
Relative Angle ◽  
Vertical Distance ◽  
Unsteady Effect ◽  
Flapping Airfoil

Bionic micro aerial vehicles have become popular because of their high thrust efficiency and deceptive appearances. Leading edge or trailing edge devices (such as slots or flaps) are often used to improve the flight performance. Birds in nature also have leading-edge devices, known as the alula that can improve their flight performance at large angles of attack. In the present study, the aerodynamic performance of a flapping airfoil with alula is numerically simulated to illustrate the effects of different alula geometric parameters. Different alula relative angles of attack β (the angle between the chord line of the alula and that of the main airfoil) and vertical distances h between the alula and the main airfoil are simulated at pre-stall and post-stall conditions. Results show that at pre-stall condition, the lift increases with the relative angle of attack and the vertical distance, but the aerodynamic performance is degraded in the presence of alula compared with no alula, whereas at post-stall condition, the alula greatly enhances the lift. However, there seems to be an optimal relative angle of attack for the maximum lift enhancement at a fixed vertical distance considering the unsteady effect, which may indicate birds can adjust the alula twisting at different spanwise positions to achieve the best flight performance. Different alula geometric parameters may affect the aerodynamic force by modifying the pressure distribution along the airfoil. The results are instructive for design of flapping-wing bionic unmanned air vehicles.

Shockwave and Noise Abatement of Transonic Fans

2004 ◽  
Author(s):  
Liping Xu
Keyword(s):  
Shock Wave ◽  
Noise Reduction ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Test Results ◽  
Shock Strength ◽  
Noise Levels ◽  
Noise Abatement ◽  
Blade Passage

The aerodynamic sources of the forward tone noise of transonic civil fans are analysed. The leading edge shockwave near the rotor tip section is identified as the main source of tone noise. By comparing the tone noise levels of the same fan operating at two different working lines, numerical calculations show that on the lower working line, the main passage shockwave is swallowed and locked into the blade passage, and the fan blades act as a shield to prevent the strong passage shock wave from propagating upstream. The calculations show that, by running the fan at a lower working line, up to 6 db abatement in the blade passing frequency (BPF) tone can be achieved through shielding the shockwave. With three dimensional CFD it is possible to design swept rotors which have desired shockwave structures near the tip region. Fan rotors with different swept leading edges have been designed to study this effect and comparisons in aerodynamics performances as well as the tone noise levels are made. It is predicted that in a swept rotor the leading edge shock strength can be further weakened and up to 5db further reduction in tone noise is possible. With a more secure shockwave shielding, a forward swept rotor has the combination of better aerodynamic performance and better noise abatement feature. The design and test results of a three dimensional fan rotor LNR2, featuring localised forward swept rotor are presented. Rig test results show that although the noise reduction through shock shielding has been demonstrated, the aerodynamics and noise are complicated by the problems specific to such localised forward swept fan.

Effect of airfoil leading edge waviness on flow structures and noise

2016 ◽  
Vol 26 (6) ◽  
pp. 1821-1842 ◽  
Author(s):  
Man Zhang ◽  
Abdelkader Frendi
Keyword(s):  
Mach Number ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Space Time ◽  
Flow Structures ◽  
Mean Flow ◽  
Content Type ◽  
Acoustic Signature ◽  
Depth Analysis ◽  
Time Correlations

Purpose – The tubercles at the leading edge of Humpback Whale flippers have been shown to increase aerodynamic efficiency. The purpose of this paper is to compute the flow structures and noise signature of a NACA0012 airfoil with and without leading edge waviness, and located in the wake of a cylinder using the hybrid RANS-LES method. Design/methodology/approach – The mean flow Mach number is 0.2 and the angle of attack used is 2°. After benchmarking the method using existing experimental results, unsteady computations were then carried-out on both airfoil geometries and for a 2° angle of attack. Findings – Results from these computations confirmed the aerodynamic benefits of the leading edge waviness. Moreover, the wavy leading edge airfoil was found to be at least 4 dB quieter than its non-wavy counterpart. In-depth analysis of the computational results revealed that the wavy leading edge airfoil breaks up the large coherent structures which are then convected at higher speeds down the trough region of the waviness in agreement with previous experimental observations. This result is supported by both the two-point and space-time correlations of the wall pressure. Research limitations/implications – The limitations of the current findings reside in the fact that both the Reynolds number and the flow Mach number are low, therefore not applicable to aircrafts. In order to extend the study to practical aircrafts one needs huge grids and large computational resources. Practical implications – The results obtained here could have a huge implications on the design of future aircrafts and spacecrafts. More specifically, the biggest benefit from such redesign is the reduction of acoustic signature as well as increased efficiency in fuel consumption. Social implications – Reducing acoustic signature from aircrafts has been a major research thrust for NASA and Federal Aviation Administration. The social impact of such reduction would be improved quality of life in airport communities. For military aircrafts, this could results in reduced detectability and hence saving lives. Originality/value – Humpback Whales have been studied by various researchers to understand the effects of leading edge “tubercles” on flow structures. What is new in this study is the numerical confirmation of the effects of the tubercles on the flow structures and the resulting noise radiations. It is shown through the use of two-point correlations and space-time correlations that the flow structures in the trough area are indeed vortex tubes.

scholarly journals Aerodynamic and performance characteristics of a passive leading edge Kruger flap at low Reynolds numbers

2012 ◽  
Vol 116 (1181) ◽  
pp. 757-767 ◽  
Author(s):  
V. M. Moraris ◽  
N. J. Lawson ◽  
K. P. Garry
Keyword(s):  
Numerical Study ◽  
Leading Edge ◽  
Detailed Comparison ◽  
Reynolds Numbers ◽  
Aerodynamic Performance ◽  
Low Reynolds Numbers ◽  
Initial Stage ◽  
And Performance ◽  
High Angles Of Attack ◽  
Basic Configuration

Abstract An experimental and numerical study was performed on a Clark Y aerofoil with a 10% chord leading edge Kruger flap to examine its aerodynamic performance at Reynolds numbers of 0·6 × 106, 1 × 106, and 1·6 × 106, to help to identify the forces and moments acting on a basic configuration. A detailed comparison of the numerical and experimental data is presented in this paper. The leading edge flap was effective at high angles of attack with an increase in CL of up to 18% over a conventional no flap configuration and delayed separation by up to 3°. The moments around the Kruger flap rotation point were calculated from the numerical analysis as an initial stage in the design of a UAV passive flap system and they are also presented in the paper.

Impact of Nonuniform Leading Edge Coatings on the Aerodynamic Performance of Compressor Airfoils

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
Michael E. Elmstrom ◽  
Knox T. Millsaps ◽  
Garth V. Hobson ◽  
Jeffrey S. Patterson
Keyword(s):  
Mach Number ◽  
Computational Fluid Dynamic ◽  
Total Pressure ◽  
Fluid Dynamic ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Transition To Turbulence ◽  
Navier Stokes ◽  
Pressure Gradients ◽  
Small Separation

A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and nonuniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge nonuniformity were added. A nonuniform coating is obtained when a liquid coating is applied to a surface with high curvature, such as an airfoil leading edge. The CFD code used, RVCQ3D, is a Reynolds averaged Navier–Stokes solver, with a k-omega turbulence model. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of a coating nonuniformity parameter. Results are presented over a range of negative and positive incidences and inlet Mach numbers from 0.6 to 0.8. A map is provided that shows the allowable degree of coating nonuniformity as a function of incidence and inlet Mach number.

Analysing and Optimizing the Aerodynamic Performance of Wind Turbine Blades Using Injected-Air Jets at Variable Frequency and Amplitude for Flow Control

2005 ◽  
Vol 29 (4) ◽  
pp. 331-339 ◽  
Author(s):  
Liu Hong ◽  
Huo Fupeng ◽  
Chen Zuoyi
Keyword(s):  
Wind Turbine ◽  
Turbine Blade ◽  
Lift Coefficient ◽  
Turbine Blades ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Operating Conditions ◽  
Suction Surface ◽  
Air Jet

Optimum aerodynamic performance of a wind turbine blade demands that the angle of attack of the relative wind on the blade remains at its optimum value. For turbines operating at constant speed, a change in wind speed causes the angle of attack to change immediately and the aerodynamic performance to decrease. Even with variable speed rotors, intrinsic time delays and inertia have similar effects. Improving the efficiency of wind turbines under variable operating conditions is one of the most important areas of research in wind power technology. This paper presents findings of an experimental study in which an oscillating air jet located at the leading edge of the suction surface of an aerofoil was used to improve the aerodynamic performance. The mean air-mass flowing through the jet during each sinusoidal period of oscillation equalled zero; i.e. the jet both blew and sucked. Experiments investigated the effects of the frequency, momentum and location of the jet stream, and the profile of the turbine blade. The study shows significant increase in the lift coefficient, especially in the stall region, under certain conditions. These findings may have important implications for wind turbine technology.

Numerical Study on Rotor-Stator Interaction of a Supersonic Reaction Turbine for a Liquid Rocket Engine

2010 ◽  
Author(s):  
Kaname Kawatsu ◽  
Naoki Tani ◽  
Nobuhiro Yamanishi
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Flow Field ◽  
Nozzle Exit ◽  
Rocket Engine ◽  
Leading Edge ◽  
Liquid Rocket Engine ◽  
Turbine Efficiency ◽  
Rotor Stator Interaction ◽  
Liquid Rocket

For an open cycle liquid rocket engine, such as the expander bleed cycle, the mass flow rate of turbine driving gas should be small, especially to improve rocket engine performance. However, work output must be high as possible. As a result, pressure ratio of the turbine becomes high, and Mach number at both nozzle exit and rotor inlet becomes supersonic. As a result, strong shock wave interaction can be generated between nozzle exit and rotor inlet, and this interaction affects the turbine aerodynamic performance. However, this rotor-stator interaction of supersonic turbine has not yet been clarified. Therefore, as the first step, it is important to clarify the structure of the flow field and to evaluate the accuracy of CFD method as practical engineering tool for liquid rocket engine design. In the present study, quasi 3-D RANS simulations were applied to the NACA supersonic turbine and the numerical results were compared with the experimental ones to evaluate numerical methodology. Turbulence models and rotor/stator interface modeling method were compared, and their impacts to the turbine aerodynamic performance estimation were evaluated. In addition to these points, the flow field between nozzle and rotor region and the turbine efficiency were investigated. The present results clarify some features of rotor-stator interaction. The shock wave, which is generated near the nozzle exit caused by encounter of nozzle exit flow, reflects at the neighbor nozzle wall and affects the rotor region. At the same time, the shock wave from the rotor leading edge impinges the nozzle cascade, and these shocks interact with each other. The present results showed that Mach number at nozzle outlet becomes different due to each turbulence and rotor/stator interface models. This difference of Mach number influences the shape of detached shock wave at the leading edge of rotor blade, and changes the entire rotor region flow field such as static pressure profile of rotor region. Thus, turbine efficiency may be influenced by these different features of flow field.

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