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Mohammad H. Sayrarfie - One of the best experts on this subject based on the ideXlab platform.
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Vibrational Response vs. Change of Trailing Sweep Angle, Tip Angle and Wing’s Thickness of a Small Wing Under Aerodynamic and Aeroelastic Forces in Super Sonic Range
Computational Mechanics ’95, 1995Co-Authors: Mohammad H. Kargarnovin, Mohammad H. SayrarfieAbstract:The equation of motion for vibration of a small wing in a flying object is derived. Then, this equation is solved numerically by using the finite element method in which two dimensional isoparametric nine nodded elements are utilized. In this study the effect of frame’s vibration is assumed to be small compared to wing’s vibrations. The loadings comprise of two types namely; the aerodynamic forces and aeroelastic forces. The aerodynamical force is defined from the kinematics of motion of the flying object and the aeroelastic force is determined by using the Second Piston Theory. Based on obtained formulations, a computer program is written in which by changing of trailing Sweep Angle or tip Angle or wing’s thickness one at the time, the wing’s natural frequencies and its time response are derived. Furthermore, for each point on the wing the stress distribution at each time step is calculated.
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vibrational response vs change of trailing Sweep Angle tip Angle and wing s thickness of a small wing under aerodynamic and aeroelastic forces in super sonic range
1995Co-Authors: Mohammad H. Kargarnovin, Mohammad H. SayrarfieAbstract:The equation of motion for vibration of a small wing in a flying object is derived. Then, this equation is solved numerically by using the finite element method in which two dimensional isoparametric nine nodded elements are utilized. In this study the effect of frame’s vibration is assumed to be small compared to wing’s vibrations. The loadings comprise of two types namely; the aerodynamic forces and aeroelastic forces. The aerodynamical force is defined from the kinematics of motion of the flying object and the aeroelastic force is determined by using the Second Piston Theory. Based on obtained formulations, a computer program is written in which by changing of trailing Sweep Angle or tip Angle or wing’s thickness one at the time, the wing’s natural frequencies and its time response are derived. Furthermore, for each point on the wing the stress distribution at each time step is calculated.
Shinji Matsumura - One of the best experts on this subject based on the ideXlab platform.
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Rotation Speed Control of Horizontal-Axis Wind Turbine by Tip Vane
JSME International Journal Series B, 1994Co-Authors: Yukimaru Shimizu, Shinji MatsumuraAbstract:This paper describes the rotation speed control of a horizontal-axis wind turbine. A tip vane has the excellent capability of improving the performance of the horizontal-axis wind turbine. Also, we found that the rotation speed of the wind turbine is controlled by changing the Sweep Angle of the tip vane. The relationships among the Sweep Angle of the tip vane, the change of wind speed and the rotor speed are investigated. As a result, a method to maintain constant rotor speed for changing wind speed, by means of an electromechanical apparatus, is developed.
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Rotation Speed Control of Horizontal Axis Wind Turbine by Tip Vane.
Transactions of the Japan Society of Mechanical Engineers Series B, 1992Co-Authors: Yukimaru Shimizu, Shinji MatsumuraAbstract:The paper describes about the rotation speed control of horizontal axis wind turbine by tip vane. Tip vane has the excellent ability to improve the performance of horizontal axis wind turbine. Also, we found that the rotation speed of wind turbine is controlled by the change of Sweep Angle of tip vane. The relationships among the Sweep Angle of tip vane, the change of wind speed and the rotor speed are investigated. As the results, the method to keep the constant rotor speed by means of electro-mechanical apparatus is developed, when wind speed changed.
Sher Afghan Khan - One of the best experts on this subject based on the ideXlab platform.
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Effect of Sweep Angle and a half sine wave on roll damping derivative of a Delta Wing
2019Co-Authors: Renita Sharon Monis, Asha Crasta, Aysha Shabana, Sher Afghan KhanAbstract:This paper presents the effect of Sweep Angle on a roll of damping derivative of a delta wing with half-sine wave for an attached shock case in supersonic/hypersonic flow has been studied analytically. The Ghosh Strip theory is replicated. By combining this with the similitude at high-speed flows lead to giving a piston theory. The initial conditions for the applicability of the theory are that the attached wave must be attached with the leading edge of the wing. The results of the present study reveal that with the increments in the Sweep Angles; it results in continuous decrease in the roll damping derivative, it is also seen that the magnitude of the decrement for lower Sweep Angle is considerable as compared to the higher values of the Sweep Angles due to the drastic change in the surface area of the wing. Roll damping derivative progressively increases with the Angle of attack; however, with the increase in the inertia level of the flow, it results in the decrement in the damping derivative and later conforms to the Mach number independence principle. Effect of the leading edge bluntness and viscous effects are neglected. Results have been obtained for the supersonic/hypersonic flow of perfect gases over a wide range of Angle of attack, planform area for different Mach numbers. In the present study, attention is on the effect of Sweep Angle of the wing on roll damping derivative at a different Angle of attack and inertia level has been studied. In contemporary theory, Leeward surface is taken along with shock waves attached with the leading edge.
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An Effect of Sweep Angle on Roll Damping Derivative for a Delta Wing with Curved Leading Edges in Unsteady Flow
International Journal of Mechanical and Production Engineering Research and Development, 2019Co-Authors: Renita Sharon Monis, Asha Crasta, Sher Afghan KhanAbstract:This paper presents the results of an analytical study to account the effect of the Sweep Angle of a delta wing whose leading edges are curved on roll damping derivative at various Angles of attack and the amplitude of the full sine waves. In the present theory, the effect of Leeward surface has been taken into consideration with the attached shock case at the leading edge. For a detached shock case, this theory will not be valid. Results have been obtained for the hypersonic flow of perfect gases over a wide range of Angle of attack and the Mach number. The results indicate that the roll damping derivative decreases with a Sweep Angle, but increase with the increase in the flow deflection Angle δ as well as with Mach M.
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effect of Sweep Angle on rolling moment derivative of an oscillating supersonic hypersonic delta wing
2014Co-Authors: Ijmer Journal, Asha Crasta, Sher Afghan Khan, Antony A. JAbstract:In the Present paper effect of Sweep Angle on roll of damping derivative of a delta wing with straight leading edges for an attached shock case in supersonic/hypersonic flow has been studied analytically. A Strip theory is used in which strips at different span wise location are independent. This combines with similitude which leads to give a piston theory. The Present theory is valid for attached shock case only. The results of the present study reveals that with the increase in the Sweep Angle; it results in continuous decrease in the roll damping derivative, it is also seen that the magnitude of the decrement for lower Sweep Angle is very large as compared to the higher values of the Sweep Angles due to the drastic change in the plan form area. Roll damping derivative progressively increases with the Angle of attack, however, with the increase in the Mach number it results in the decrement in the damping derivative and later conforms to the Mach number principle. Effects of wave reflection, leading edge bluntness, and viscosity have not been taken into account. Results have been obtained for supersonic/hypersonic flow of perfect gases over a wide range of Angle of attack, plan form area, and the Mach number.
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Effect of Sweep Angle on Rolling Moment Derivative of an Oscillating Supersonic/Hypersonic Delta Wing
2014Co-Authors: Ijmer Journal, Asha Crasta, Sher Afghan Khan, Antony A. JAbstract:In the Present paper effect of Sweep Angle on roll of damping derivative of a delta wing with straight leading edges for an attached shock case in supersonic/hypersonic flow has been studied analytically. A Strip theory is used in which strips at different span wise location are independent. This combines with similitude which leads to give a piston theory. The Present theory is valid for attached shock case only. The results of the present study reveals that with the increase in the Sweep Angle; it results in continuous decrease in the roll damping derivative, it is also seen that the magnitude of the decrement for lower Sweep Angle is very large as compared to the higher values of the Sweep Angles due to the drastic change in the plan form area. Roll damping derivative progressively increases with the Angle of attack, however, with the increase in the Mach number it results in the decrement in the damping derivative and later conforms to the Mach number principle. Effects of wave reflection, leading edge bluntness, and viscosity have not been taken into account. Results have been obtained for supersonic/hypersonic flow of perfect gases over a wide range of Angle of attack, plan form area, and the Mach number.
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Stability Derivatives of a Delta Wing with Straight Leading Edge in the Newtonian Limit
2013Co-Authors: Asha Crasta, Sher Afghan KhanAbstract:This paper presents an analytical method to predict the aerodynamic stability derivatives of oscillating delta wings with straight leading edge. It uses the Ghosh similitude and the strip theory to obtain the expressions for stability derivatives in pitch and roll in the Newtonian limit. The present theory gives a quick and approximate method to estimate the stability derivatives which is very essential at the design stage. They are applicable for wings of arbitrary plan form shape at high Angles of attack provided the shock wave is attached to the leading edge of the wing. The expressions derived for stability derivatives become exact in the Newtonian limit. The stiffness derivative and damping derivative in pitch and roll are dependent on the geometric parameter of the wing. It is found that stiffness derivative linearly varies with the pivot position. In the case of damping derivative since expressions for these derivatives are non-linear and the same is reflected in the results. Roll damping derivative also varies linearly with respect to the Angle of attack. When the variation of roll damping derivative was considered, it is found it also, varies linearly with Angle of attack for given Sweep Angle, but with increase in Sweep Angle there is continuous decrease in the magnitude of the roll damping derivative however, the values differ for different values in Sweep Angle and the same is reflected in the result when it was studied with respect to Sweep Angle. From the results it is found that one can arrive at the optimum value of the Angle of attack Sweep Angle which will give the best performance.
Sandeep Saha - One of the best experts on this subject based on the ideXlab platform.
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effect of Sweep Angle on wing strake vortex interaction and breakdown over double delta wings
2017 First International Conference on Recent Advances in Aerospace Engineering (ICRAAE), 2017Co-Authors: Abhishek Sinha, Arun Kumar Suthar, Sourabh Sahoo, Amit A. Pawar, Kumar Sanat Ranjan, Sandeep SahaAbstract:The lifting surface of present-day aircrafts cruising at high speeds is often delta wings or a variant in order to enhance lift and maneuverability. These aircrafts are most vulnerable to accidents during take-off and landing because of the danger of abrupt loss in lift due to vortex-breakdown at high Angles of attack on the top-surface. In particular, a double delta wing experiences the vortex-breakdown phenomenon that is further complicated by the wing-strake vortex-interaction. We attempt to gain insight into the wing-strake vortex-interaction and vortexbreakdown phenomenon using a two-pronged approach: Windtunnel experiments and numerical simulations of two models of the double delta wing having same apex Angle but different wing Sweep Angle. The pressure distribution on the upper surface of the wing at varying Angles of attack is reported and compared to numerical simulations to glean substantive data on the physical mechanism. We also identify the location of vortex breakdown from pressure distribution for both the models.
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Effect of Sweep Angle on wing-strake vortex — Interaction and breakdown over double delta wings
2017 First International Conference on Recent Advances in Aerospace Engineering (ICRAAE), 2017Co-Authors: Abhishek Sinha, Arun Kumar Suthar, Sourabh Sahoo, Amit A. Pawar, Kumar Sanat Ranjan, Sandeep SahaAbstract:The lifting surface of present-day aircrafts cruising at high speeds is often delta wings or a variant in order to enhance lift and maneuverability. These aircrafts are most vulnerable to accidents during take-off and landing because of the danger of abrupt loss in lift due to vortex-breakdown at high Angles of attack on the top-surface. In particular, a double delta wing experiences the vortex-breakdown phenomenon that is further complicated by the wing-strake vortex-interaction. We attempt to gain insight into the wing-strake vortex-interaction and vortexbreakdown phenomenon using a two-pronged approach: Windtunnel experiments and numerical simulations of two models of the double delta wing having same apex Angle but different wing Sweep Angle. The pressure distribution on the upper surface of the wing at varying Angles of attack is reported and compared to numerical simulations to glean substantive data on the physical mechanism. We also identify the location of vortex breakdown from pressure distribution for both the models.
Mohammad H. Kargarnovin - One of the best experts on this subject based on the ideXlab platform.
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Vibrational Response vs. Change of Trailing Sweep Angle, Tip Angle and Wing’s Thickness of a Small Wing Under Aerodynamic and Aeroelastic Forces in Super Sonic Range
Computational Mechanics ’95, 1995Co-Authors: Mohammad H. Kargarnovin, Mohammad H. SayrarfieAbstract:The equation of motion for vibration of a small wing in a flying object is derived. Then, this equation is solved numerically by using the finite element method in which two dimensional isoparametric nine nodded elements are utilized. In this study the effect of frame’s vibration is assumed to be small compared to wing’s vibrations. The loadings comprise of two types namely; the aerodynamic forces and aeroelastic forces. The aerodynamical force is defined from the kinematics of motion of the flying object and the aeroelastic force is determined by using the Second Piston Theory. Based on obtained formulations, a computer program is written in which by changing of trailing Sweep Angle or tip Angle or wing’s thickness one at the time, the wing’s natural frequencies and its time response are derived. Furthermore, for each point on the wing the stress distribution at each time step is calculated.
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vibrational response vs change of trailing Sweep Angle tip Angle and wing s thickness of a small wing under aerodynamic and aeroelastic forces in super sonic range
1995Co-Authors: Mohammad H. Kargarnovin, Mohammad H. SayrarfieAbstract:The equation of motion for vibration of a small wing in a flying object is derived. Then, this equation is solved numerically by using the finite element method in which two dimensional isoparametric nine nodded elements are utilized. In this study the effect of frame’s vibration is assumed to be small compared to wing’s vibrations. The loadings comprise of two types namely; the aerodynamic forces and aeroelastic forces. The aerodynamical force is defined from the kinematics of motion of the flying object and the aeroelastic force is determined by using the Second Piston Theory. Based on obtained formulations, a computer program is written in which by changing of trailing Sweep Angle or tip Angle or wing’s thickness one at the time, the wing’s natural frequencies and its time response are derived. Furthermore, for each point on the wing the stress distribution at each time step is calculated.
