scholarly journals IMPLEMENTATION OF TRANSONIC AREA RULE AND SWEPT BACK DELTA WING DESIGN ON AN AIRCRAFT.

2022 ◽  
pp. 20-25
Author(s):  
S. Rajat Singh ◽  
Y.D. Dwivedi
Keyword(s):  
Delta Wing ◽  
Wave Drag ◽  
Wing Design ◽  
Cross Sectional ◽  
Ansys Fluent ◽  
Area Distribution ◽  
Area Rule ◽  
Drag Ratio ◽  
Commercial Aircrafts ◽  
Lift To Drag Ratio

The transonic area rule was first implemented in the 1950s. It is an important concept related to the drag on an aircraft or other body in transonic and supersonic flight which states that two airplanes with the same longitudinal cross-sectional area distribution have the same wave drag, independent of how the area is distributed laterally. A swept back delta wing increases the critical Mach number of the wing and performs well at low speeds, as a result of unique swirling vortices that form on the upper surface of the wing. BOOM Supersonic plans to bring back Supersonic Commercial aircrafts by implementing these modifications in the famous Concorde. In this paper two aircraft designs inspired by Concorde and BOOM Overture are compared using ANSYS Fluent. These were designed in CATIA with changes in fuselage dimensions, wing configuration and engine configuration. The lift to drag ratio of both the designs are calculated and compared. Pressure contours, velocity vectors, vector pathlines, turbulence pathlines and pressure pathlines are also compared. The results show that the design with the implementation of transonic area rule and swept back delta wing has a better Lift to Drag ratio when compared to the design with a wide fuselage and a delta wing design.

Aerodynamic study of single corrugated variable-camber morphing aerofoil concept

2021 ◽  
pp. 1-29
Author(s):  
K. Dhileep ◽  
D. Kumar ◽  
P.N. Gautham Vigneswar ◽  
P. Soni ◽  
S. Ghosh ◽  
...  
Keyword(s):  
Finite Volume ◽  
Interpolation Method ◽  
Beam Theory ◽  
Small Deformation ◽  
Aerodynamic Characteristics ◽  
Ansys Fluent ◽  
Aerodynamic Efficiency ◽  
Finite Element Software ◽  
Drag Ratio ◽  
Lift To Drag Ratio

Abstract Camber morphing is an effective way to control the lift generated by any aerofoil and potentially improve the range (as measured by the lift-to-drag ratio) and endurance (as measured by $C_l^{3/2}/C_d$ ). This can be especially useful for fixed-wing Unmanned Aerial Vehicles (UAVs) undergoing different flying manoeuvres and flight phases. This work investigates the aerodynamic characteristics of the NACA0012 aerofoil morphed using a Single Corrugated Variable-Camber (SCVC) morphing approach. Structural analysis and morphed shapes are obtained based on small-deformation beam theory using chain calculations and validated using finite-element software. The aerofoil is then reconstructed from the camber line using a Radial Basis Function (RBF)-based interpolation method (J.H.S. Fincham and M.I. Friswell, “Aerodynamic optimisation of a camber morphing aerofoil,” Aerosp. Sci. Technol., 2015). The aerodynamic analysis is done by employing two different finite-volume solvers (OpenFOAM and ANSYS-Fluent) and a panel method code (XFoil). Results reveal that the aerodynamic coefficients predicted by the two finite-volume solvers using a fully turbulent flow assumption are similar but differ from those predicted by XFoil. The aerodynamic efficiency and endurance factor of morphed aerofoils indicate that morphing is beneficial at moderate to high lift requirements. Further, the optimal morphing angle increases with an increase in the required lift. Finally, it is observed for a fixed angle-of-attack that an optimum morphing angle exists for which the aerodynamic efficiency becomes maximum.

Effects of leading-edge bevel angle on the aerodynamic forces of a non-slender 50° delta wing

2005 ◽  
Vol 109 (1098) ◽  
pp. 403-407 ◽  
Author(s):  
J. J. Wang ◽  
S. F. Lu
Keyword(s):  
Delta Wing ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Relative Thickness ◽  
Aerodynamic Performances ◽  
Stall Angle ◽  
Drag Ratio ◽  
Low Speed Wind Tunnel ◽  
Lift To Drag Ratio ◽  
Blunt Leading Edge

Abstract The aerodynamic performances of a non-slender 50° delta wing with various leading-edge bevels were measured in a low speed wind tunnel. It is found that the delta wing with leading-edge bevelled leeward can improve the maximum lift coefficient and maximum lift to drag ratio, and the stall angle of the wing is also delayed. In comparison with the blunt leading-edge wing, the increment of maximum lift to drag ratio is 200%, 98% and 100% for the wings with relative thickness t/c = 2%, t/c = 6.7% and t/c = 10%, respectively.

Note on the Evaluation of the Integral

1954 ◽  
Vol 58 (527) ◽  
pp. 787-788 ◽  
Author(s):  
N. A. Routledge ◽  
W. T. Lord ◽  
E. Eminton
Keyword(s):  
Logarithmic Singularity ◽  
Wave Drag ◽  
Cross Sectional Area ◽  
Sectional Area ◽  
Direct Evaluation ◽  
Cross Sectional ◽  
Numerical Work ◽  
Singular Form ◽  
Area Distribution ◽  
Kinetic Pressure

The Integral I occurs frequently in aerodynamics. To cite a particular case, Ward has shown that ifS(x)is the cross-sectional area distribution of a slender body its wave dragDis given to a first approxition by D = qI/(2π), whereqis the kinetic pressure. In this case,S(x)is often defined numerically, and the direct evaluation of I is then complicated by the presence of the logarithmic singularity. Several methods may be employed to avoid this complication. Legendre has recently suggested rewriting I in a non-singular form, which unfortunately is not well suited to numerical work. Here, a method for the evaluation of the unmodified integral is presented.

Control of shock-induced vortex breakdown on a delta-wing-body configuration in the transonic regime

2021 ◽  
pp. 1-25
Author(s):  
Rajan B. Kurade ◽  
L. Venkatakrishnan ◽  
G. Jagadeesh
Keyword(s):  
Vortex Breakdown ◽  
Delta Wing ◽  
Pressure Fluctuations ◽  
Angle Of Attack ◽  
Aerodynamic Coefficients ◽  
Delta Wings ◽  
Drag Ratio ◽  
Lift To Drag Ratio ◽  
Sudden Jump ◽  
Modest Level

Abstract Shock-induced vortex breakdown, which occurs on the delta wings at transonic speed, causes a sudden and significant change in the aerodynamic coefficients at a moderate angle-of-attack. Wind-tunnel tests show a sudden jump in the aerodynamic coefficients such as lift force, pitching moment and centre of pressure which affect the longitudinal stability and controllability of the vehicle. A pneumatic jet operated at sonic condition blown spanwise and along the vortex core over a 60° swept delta-wing-body configuration is found to be effective in postponing this phenomenon by energising the vortical structure, pushing the vortex breakdown location downstream. The study reports that a modest level of spanwise blowing enhances the lift by about 6 to 9% and lift-to-drag ratio by about 4 to 9%, depending on the free-stream transonic Mach number, and extends the usable angle-of-attack range by 2°. The blowing is found to reduce the magnitude of unsteady pressure fluctuations by 8% to 20% in the aft portion of the wing, depending upon the method of blowing. Detailed investigations carried out on the location of blowing reveal that the blowing close to the apex of the wing maximises the benefits.

Making Air Travelling More Economical, An Innovative Drag Reduction Approach For A Supercritical Wing Section Using Shock Control Bumps

2014 ◽  
Vol 1 (1) ◽  
pp. 215-220
Author(s):  
A Saeed ◽  
Malik. S. Raza ◽  
Ahmed Mohsin Khalil
Keyword(s):  
Drag Reduction ◽  
Three Dimensional ◽  
Wave Drag ◽  
Two Dimensional ◽  
Supercritical Airfoil ◽  
Shock Control ◽  
Wing Section ◽  
First Choice ◽  
Drag Ratio ◽  
Lift To Drag Ratio

AbstractAir travelling is the second largest travelling medium used by people. In future it is expected to be the first choice for the travellers. As increase in the price of oil cost of air travelling is getting higher. Engineers are forced to find the cheaper means of travelling by innovating new techniques. This paper presents the new idea to reduce air travelling cost by reducing drag, which is major driving factor of high fuel consumption. Two-dimensional and three-dimensional shock control contour bumps have been designed and analysed for a supercritical wing section with the aim of transonic wave drag reduction. A supercritical airfoil (NACA SC (02)-0714) has been selected for this study considering the fact that most modern jet transport aircraft that operate in the transonic flow regime (cruise at transonic speeds) employ supercritical airfoil sections. It is to be noted that a decrease in the transonic wave drag without loss in lift would result in an increased lift to drag ratio, which being a key range parameter could potentially increase both the range and endurance of the aircraft. The major geometric bump parameters such as length, height, crest and span have been altered for both the two-dimensional and three-dimensional bumps in order to obtain the optimum location and shape of the bump. Once an optimum standalone three-dimensional bump has been acquired an array of bumps has been manually placed spanwise of an unswept supercritical wing and analysed under fully turbulent flow conditions. Different configurations have been tested with varying three-dimensional bump spacing in order to determine the contribution of bump spacing on overall performance. The results show a 14 percent drag reduction and a consequent 16 percent lift to drag ratio rise at the design Mach number for the optimum arrangement of bumps along the wing span. This innovative technique proves to be a bridge between economical problems and engineering solutions and a milestone for aviation engineering.

Effect of anhedral on a reverse delta wing

2017 ◽  
Vol 232 (12) ◽  
pp. 2317-2325 ◽  
Author(s):  
T Lee ◽  
LS Ko ◽  
V Tremblay-Dionne
Keyword(s):  
Lift Force ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Tangential Velocity ◽  
Bottom Surface ◽  
Drag Ratio ◽  
Lift To Drag Ratio

The effect of anhedral on the vorticity flowfield and aerodynamic loadings of a slender reverse delta wing was studied experimentally. The addition of anhedral always led to a reduced lift and lift-to-drag ratio in comparison with their clean-wing counterparts. The drag was increased with increasing anhedral compared to the clean wing at the same lift condition. The reverse delta wing vortex retained its regularity to a higher angle of attack as the anhedral was increased. The persistent outboard location of the reverse delta wing vortex suggests that the lift force was mainly produced by the pressure exerted on the bottom surface of the wing. The anhedral also led to an increased vorticity level and tangential velocity.

scholarly journals Computational Investigation of Aerodynamic Characteristics for Elliptic UAVs’ Wing

2018 ◽  
Vol 6 (3) ◽  
Author(s):  
Amaal Attiah ◽  
Ibrahim Elbadawy ◽  
Osama E. Mahmoud
Keyword(s):  
Reynolds Number ◽  
Computational Study ◽  
Angle Of Attack ◽  
Aerodynamic Characteristics ◽  
Thickness Ratio ◽  
Ansys Fluent ◽  
Rotor Wing ◽  
Drag Ratio ◽  
Lift To Drag Ratio ◽  
Elliptic Airfoil

Unmanned Aerial Vehicles, UAVs, gained an important role in modern military and civilian applications. Developments in UAVs technology improve its performance and maneuverability with acceptable cost. Elliptic airfoil had been widely used in the development of Rotor/Wing subsonic aircraft. The present work aims to investigate the effect of various elliptic airfoil parameters, such as Reynolds number, angle of attack and airfoil thickness, on aerodynamic behavior using two-dimensional computational study. The computational results were validated by experimental results. Angles of attack was evaluated from 0° to 18° in order to analyze aerodynamic characteristics up to stall condition, while Reynolds number was evaluated at values of 1×10⁵, 3×105, 2×106, and 8×106, to cover the range of rotary and fixed wing flight conditions. Thickness ratio was ranged from 5% to 25% to include the UAVs airfoil thicknesses so that choice best thickness gets max lift to drag ratio. In addition, the thicknesses location was evaluated for a range of 30% to 70% to get suitable location gets max left to drag ratio. The ANSYS-Fluent software was used with Spalart-Allmaras turbulence model, and found that the maximum lift to drag ratio which improve the UAV capability in this study is at Re=2×106, angle of attack at 8°, max thickness ratio of (0.1chord) located at (0.3chord).

scholarly journals Shape Optimization of an Airfoil in Ground Effect for Application to WIG Craft

2014 ◽  
Vol 2014 ◽  
pp. 1-11 ◽  
Author(s):  
Yilei He ◽  
Qiulin Qu ◽  
Ramesh K. Agarwal
Keyword(s):  
Genetic Algorithm ◽  
Flow Field ◽  
Lift Coefficient ◽  
Ground Effect ◽  
Navier Stokes ◽  
Ansys Fluent ◽  
Multiobjective Genetic Algorithm ◽  
Lift And Drag ◽  
Drag Ratio ◽  
Lift To Drag Ratio

This paper employs a multiobjective genetic algorithm (MOGA) to optimize the shape of a widely used wing in ground (WIG) aircraft airfoil NACA 4412 to improve its lift and drag characteristics, in particular to achieve two objectives, that is, to increase its lift and its lift to drag ratio. The commercial software ANSYS FLUENT is employed to calculate the flow field on an adaptive structured mesh generated by ANSYS ICEM software using the Reynolds-Averaged Navier-Stokes (RANS) equations in conjunction with a one equation Spalart-Allmaras (SA) turbulence model. The results show significant improvement in both the lift coefficient and lift to drag ratio of the optimized airfoil compared to the original NACA 4412 airfoil. It is demonstrated that the performance of a wing in ground (WIG) aircraft can be improved by using the optimized airfoil.

Some Effects of Thickness on the Longitudinal Characteristics of Sharp-Edged Delta Wings at Low Speeds

1968 ◽  
Vol 72 (686) ◽  
pp. 151-155 ◽  
Author(s):  
L. C. Squire
Keyword(s):  
Aspect Ratio ◽  
Delta Wing ◽  
Leading Edge ◽  
Air Transport ◽  
Wing Area ◽  
Delta Wings ◽  
Amount Of Information ◽  
Drag Ratio ◽  
Lift To Drag Ratio ◽  
Passenger Cabin

Recently there has been renewed interest in the concept of an all-wing aircraft as a means of producing cheap air transport over relatively short distances. It is natural that with the large amount of information on slender wings now available an all-wing aircraft based on a sharp-edged slender planform should be considered for this role. One of the difficulties immediately faced in developing this concept is that since the aircraft must carry a large number of passengers it is necessary that as much of the wing area as possible should be deep enough to provide for a large passenger cabin. Thus the wing will be very thick over a large part of its area. If this condition is not met, then the aircraft has too much wing area and hence too high a structure weight. Typically one may think of an aircraft with a delta wing of aspect ratio 2 and with a wing thickness of from 15% to 20% of the root chord over as much of the wing area as possible. At first sight thickness of this order eliminates the main advantage of slender wings since the effect of thickness is usually to reduce the strength of the leading-edge vortices and hence the non-linear lift. Thus the incidence for a given lift is increased above that for a thin wing. This in turn means that the lift-to-drag ratio may be smaller.

Note on the Numerical Evaluation of the Wave Drag of Smooth Slender Bodies Using Optimum Area Distributions for Minimum Wave Drag

1956 ◽  
Vol 60 (541) ◽  
pp. 61-63 ◽  
Author(s):  
E. Eminton ◽  
W. T. Lord
Keyword(s):  
Numerical Evaluation ◽  
Practical Method ◽  
Wave Drag ◽  
The Body ◽  
Cross Sectional ◽  
Area Distribution ◽  
Slender Bodies ◽  
Kinetic Pressure ◽  
Image Position ◽  
Sectional Shape

The linearised theory value of the wave drag D at zero lift of a smooth slender body of arbitrary cross-sectional shape was shown by Ward to be given bywhere S (x), 0 ⩽ x ⩽ 1, is the cross-sectional area distribution of the body and q is the kinetic pressure. The development of this result has aroused interest in two problems: the derivation of the optimum area distribution for minimum wave drag under certain specified conditions and the numerical evaluation of the wave drag of a specified area distribution. These apparently distinct problems have hitherto been treated separately, but it is shown here how an attempt to solve the first problem has led to a practical method of solving the second.

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