Is drag coefficient lowest at zero angle of attack? Announcing the arrival of Valued Associate #679: Cesar Manara Planned maintenance scheduled April 23, 2019 at 23:30 UTC (7:30pm US/Eastern)How do insects decrease aircraft performance?How to draw NACA 6-Series Airfoils?How can the zero-lift drag coefficient (parasitic drag) be calculated?What is the relation between the Lift Coefficient and the Angle of Attack?Is it possible to fly horizontally with zero angle of attack?How to find trim condition of a sectional airfoil without knowing the angle of attack?What is the effect of flow separation on lift, pressure distribution and drag?How can the zero-lift drag coefficient (parasitic drag) be calculated?Do negative angles-of-attack create lift?How do you calculate the lift coefficient of an airfoil at zero angle of attack?Calculating induced drag approximation using XFoil generated parasitic dragDoes speed or angle of attack generally have the greatest impact on total induced drag?What's the theoretical background of the critical angle of attack?
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Is drag coefficient lowest at zero angle of attack?
Announcing the arrival of Valued Associate #679: Cesar Manara
Planned maintenance scheduled April 23, 2019 at 23:30 UTC (7:30pm US/Eastern)How do insects decrease aircraft performance?How to draw NACA 6-Series Airfoils?How can the zero-lift drag coefficient (parasitic drag) be calculated?What is the relation between the Lift Coefficient and the Angle of Attack?Is it possible to fly horizontally with zero angle of attack?How to find trim condition of a sectional airfoil without knowing the angle of attack?What is the effect of flow separation on lift, pressure distribution and drag?How can the zero-lift drag coefficient (parasitic drag) be calculated?Do negative angles-of-attack create lift?How do you calculate the lift coefficient of an airfoil at zero angle of attack?Calculating induced drag approximation using XFoil generated parasitic dragDoes speed or angle of attack generally have the greatest impact on total induced drag?What's the theoretical background of the critical angle of attack?
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The drag coefficient of a symmetric airfoil is lowest when its angle of attack is zero. I'm not sure if this is true in general.
aerodynamics airfoil drag angle-of-attack
asked Apr 14 at 22:17
simple jacksimple jack
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The drag coefficient of a symmetric airfoil is lowest when its angle of attack is zero. I'm not sure if this is true in general.
aerodynamics airfoil drag angle-of-attack
asked Apr 14 at 22:17
simple jacksimple jack
283
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simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
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– Ralph J
Apr 14 at 22:34
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The drag coefficient of a symmetric airfoil is lowest when its angle of attack is zero. I'm not sure if this is true in general.
aerodynamics airfoil drag angle-of-attack
asked Apr 14 at 22:17
simple jacksimple jack
283
New contributor
simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
Check out our Code of Conduct.
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The drag coefficient of a symmetric airfoil is lowest when its angle of attack is zero. I'm not sure if this is true in general.
aerodynamics airfoil drag angle-of-attack
aerodynamics airfoil drag angle-of-attack
asked Apr 14 at 22:17
simple jacksimple jack
283
New contributor
simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
Check out our Code of Conduct.
asked Apr 14 at 22:17
simple jacksimple jack
283
New contributor
simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
Check out our Code of Conduct.
asked Apr 14 at 22:17
simple jacksimple jack
283
New contributor
simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
Check out our Code of Conduct.
asked Apr 14 at 22:17
simple jacksimple jack
283
asked Apr 14 at 22:17
simple jacksimple jack
283
283
New contributor
simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
Check out our Code of Conduct.
New contributor
simple jack is a new contributor to this site. Take care in asking for clarification, commenting, and answering.
Check out our Code of Conduct.
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Welcome to Av.SE!
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– Ralph J
Apr 14 at 22:34
add a comment |
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Welcome to Av.SE!
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– Ralph J
Apr 14 at 22:34
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Welcome to Av.SE!
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– Ralph J
Apr 14 at 22:34
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Welcome to Av.SE!
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– Ralph J
Apr 14 at 22:34
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Obviously, drag should be smallest for symmetrical airfoils at zero angle of attack.
However, most airfoils have camber, and then the lowest drag is at positive lift coefficients in case of positive camber. Where that point is exactly depends on many parameters; in case of laminar airfoils even local imperfections can have a noticeable effect. Generally, the lowest drag can be found at the angle of attack where the stagnation point is exactly at the center of the leading edge, where the local curvature is highest. A deviation from this point will force the flow on one side to negotiate this point of highest curvature all by itself, resulting in a suction peak which will increase the losses in the boundary layer.
This is a theoretical drag polar (calculated with XFOIL) of an airfoil with a 20% camber flap at different settings and a Reynolds number of 1.5 million. The laminar bucket is clearly visible and produces a range of lift coefficients with nearly identical drag. The small waviness at the lower end of the laminar bucket is an artificial result of smoothing the plot.
What is obvious is how camber shifts the area of minimum drag up and down. If you use the right side of the plot to find the angle of attack of minimum drag, you will find that this is not constant but goes down as flap angles go up. For the 0° flap polar it is at about -2° AoA. This is caused by the induced angle of attack which increases with the lift coefficient.
The 6-series NACA airfoils were the first to be systematically designed with the pressure distribution in mind, and the design lift coefficient is where the condition of the ideal stagnation point location is met. This is indicated by the digit right after the hyphen in the airfoil name: Divide this digit by 10 and you have the lift coefficient of minimum drag.
Example: The $63_1-412$ airfoil has its lowest drag at a lift coefficient of 0.4.
If you want to know the angle of attack with the lowest drag of a whole airplane, this is a very different matter and needs to include the drag due to lift, which is of course smallest at the zero lift polar point.
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
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$begingroup$
Obviously, drag should be smallest for symmetrical airfoils at zero angle of attack.
However, most airfoils have camber, and then the lowest drag is at positive lift coefficients in case of positive camber. Where that point is exactly depends on many parameters; in case of laminar airfoils even local imperfections can have a noticeable effect. Generally, the lowest drag can be found at the angle of attack where the stagnation point is exactly at the center of the leading edge, where the local curvature is highest. A deviation from this point will force the flow on one side to negotiate this point of highest curvature all by itself, resulting in a suction peak which will increase the losses in the boundary layer.
This is a theoretical drag polar (calculated with XFOIL) of an airfoil with a 20% camber flap at different settings and a Reynolds number of 1.5 million. The laminar bucket is clearly visible and produces a range of lift coefficients with nearly identical drag. The small waviness at the lower end of the laminar bucket is an artificial result of smoothing the plot.
What is obvious is how camber shifts the area of minimum drag up and down. If you use the right side of the plot to find the angle of attack of minimum drag, you will find that this is not constant but goes down as flap angles go up. For the 0° flap polar it is at about -2° AoA. This is caused by the induced angle of attack which increases with the lift coefficient.
The 6-series NACA airfoils were the first to be systematically designed with the pressure distribution in mind, and the design lift coefficient is where the condition of the ideal stagnation point location is met. This is indicated by the digit right after the hyphen in the airfoil name: Divide this digit by 10 and you have the lift coefficient of minimum drag.
Example: The $63_1-412$ airfoil has its lowest drag at a lift coefficient of 0.4.
If you want to know the angle of attack with the lowest drag of a whole airplane, this is a very different matter and needs to include the drag due to lift, which is of course smallest at the zero lift polar point.
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
$endgroup$
add a comment |
$begingroup$
Obviously, drag should be smallest for symmetrical airfoils at zero angle of attack.
However, most airfoils have camber, and then the lowest drag is at positive lift coefficients in case of positive camber. Where that point is exactly depends on many parameters; in case of laminar airfoils even local imperfections can have a noticeable effect. Generally, the lowest drag can be found at the angle of attack where the stagnation point is exactly at the center of the leading edge, where the local curvature is highest. A deviation from this point will force the flow on one side to negotiate this point of highest curvature all by itself, resulting in a suction peak which will increase the losses in the boundary layer.
This is a theoretical drag polar (calculated with XFOIL) of an airfoil with a 20% camber flap at different settings and a Reynolds number of 1.5 million. The laminar bucket is clearly visible and produces a range of lift coefficients with nearly identical drag. The small waviness at the lower end of the laminar bucket is an artificial result of smoothing the plot.
What is obvious is how camber shifts the area of minimum drag up and down. If you use the right side of the plot to find the angle of attack of minimum drag, you will find that this is not constant but goes down as flap angles go up. For the 0° flap polar it is at about -2° AoA. This is caused by the induced angle of attack which increases with the lift coefficient.
The 6-series NACA airfoils were the first to be systematically designed with the pressure distribution in mind, and the design lift coefficient is where the condition of the ideal stagnation point location is met. This is indicated by the digit right after the hyphen in the airfoil name: Divide this digit by 10 and you have the lift coefficient of minimum drag.
Example: The $63_1-412$ airfoil has its lowest drag at a lift coefficient of 0.4.
If you want to know the angle of attack with the lowest drag of a whole airplane, this is a very different matter and needs to include the drag due to lift, which is of course smallest at the zero lift polar point.
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
$endgroup$
add a comment |
$begingroup$
Obviously, drag should be smallest for symmetrical airfoils at zero angle of attack.
However, most airfoils have camber, and then the lowest drag is at positive lift coefficients in case of positive camber. Where that point is exactly depends on many parameters; in case of laminar airfoils even local imperfections can have a noticeable effect. Generally, the lowest drag can be found at the angle of attack where the stagnation point is exactly at the center of the leading edge, where the local curvature is highest. A deviation from this point will force the flow on one side to negotiate this point of highest curvature all by itself, resulting in a suction peak which will increase the losses in the boundary layer.
This is a theoretical drag polar (calculated with XFOIL) of an airfoil with a 20% camber flap at different settings and a Reynolds number of 1.5 million. The laminar bucket is clearly visible and produces a range of lift coefficients with nearly identical drag. The small waviness at the lower end of the laminar bucket is an artificial result of smoothing the plot.
What is obvious is how camber shifts the area of minimum drag up and down. If you use the right side of the plot to find the angle of attack of minimum drag, you will find that this is not constant but goes down as flap angles go up. For the 0° flap polar it is at about -2° AoA. This is caused by the induced angle of attack which increases with the lift coefficient.
The 6-series NACA airfoils were the first to be systematically designed with the pressure distribution in mind, and the design lift coefficient is where the condition of the ideal stagnation point location is met. This is indicated by the digit right after the hyphen in the airfoil name: Divide this digit by 10 and you have the lift coefficient of minimum drag.
Example: The $63_1-412$ airfoil has its lowest drag at a lift coefficient of 0.4.
If you want to know the angle of attack with the lowest drag of a whole airplane, this is a very different matter and needs to include the drag due to lift, which is of course smallest at the zero lift polar point.
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
$endgroup$
Obviously, drag should be smallest for symmetrical airfoils at zero angle of attack.
However, most airfoils have camber, and then the lowest drag is at positive lift coefficients in case of positive camber. Where that point is exactly depends on many parameters; in case of laminar airfoils even local imperfections can have a noticeable effect. Generally, the lowest drag can be found at the angle of attack where the stagnation point is exactly at the center of the leading edge, where the local curvature is highest. A deviation from this point will force the flow on one side to negotiate this point of highest curvature all by itself, resulting in a suction peak which will increase the losses in the boundary layer.
This is a theoretical drag polar (calculated with XFOIL) of an airfoil with a 20% camber flap at different settings and a Reynolds number of 1.5 million. The laminar bucket is clearly visible and produces a range of lift coefficients with nearly identical drag. The small waviness at the lower end of the laminar bucket is an artificial result of smoothing the plot.
What is obvious is how camber shifts the area of minimum drag up and down. If you use the right side of the plot to find the angle of attack of minimum drag, you will find that this is not constant but goes down as flap angles go up. For the 0° flap polar it is at about -2° AoA. This is caused by the induced angle of attack which increases with the lift coefficient.
The 6-series NACA airfoils were the first to be systematically designed with the pressure distribution in mind, and the design lift coefficient is where the condition of the ideal stagnation point location is met. This is indicated by the digit right after the hyphen in the airfoil name: Divide this digit by 10 and you have the lift coefficient of minimum drag.
Example: The $63_1-412$ airfoil has its lowest drag at a lift coefficient of 0.4.
If you want to know the angle of attack with the lowest drag of a whole airplane, this is a very different matter and needs to include the drag due to lift, which is of course smallest at the zero lift polar point.
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
answered Apr 15 at 2:49
Peter KämpfPeter Kämpf
162k12413659
162k12413659
add a comment |
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simple jack is a new contributor. Be nice, and check out our Code of Conduct.
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– Ralph J
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