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aircraft:tmd:airfoil [2018/09/20 12:03] – [Airfoil] jhaircraft:tmd:airfoil [2020/09/28 22:44] (current) jh
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 === Other Aerodynamic classes === === Other Aerodynamic classes ===
-[[aerowing]], [[aerofuselage]], [[airbrake]], [[propeller]], [[aerodrag]], [[bladeforce]]+[[aerowing]], [[aerofuselage]], [[airbrake]], [[propeller]], [[aerodrag]]
  
 ===== Angle of attack ===== ===== Angle of attack =====
  
-The angle of attack (also known as AOA or α ("alpha")) is defined as the angle between the undisturbed incoming air-flow and the chord line ("the zero chamber line") of an airfoil. The lift, drag and moment of an [[aircraft:tmd:airfoil|airfoil]] are all functions of the angle of attack.+The angle of attack (also known as AOA or α ("alpha")) is defined as the angle between the undisturbed incoming air-flow and the chord line ("the zero camber line") of an airfoil. The lift, drag and moment of an [[aircraft:tmd:airfoil|airfoil]] are all functions of the angle of attack.
  
 {{ Alpha.png?nolink }} {{ Alpha.png?nolink }}
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 The lift coefficient Cl and momentum coefficient Cm increase linear with alpha in the Range of ''-AttachedRange'' to ''+AttachedRange'' The lift coefficient Cl and momentum coefficient Cm increase linear with alpha in the Range of ''-AttachedRange'' to ''+AttachedRange''
 The gradient of the lift coefficient in respect to alpha is ''ClAlpha'' which equals 2 * Math.Pi for most real airfoils.   The gradient of the lift coefficient in respect to alpha is ''ClAlpha'' which equals 2 * Math.Pi for most real airfoils.  
-> Leave it at ''6.28'' (2 * pi)+> Leave it at ''6.28'' (2 * pi) unless you are dealing with a delta wing
  
 {{ Airfoil_AttachedRange_ClAlpha.png?nolink }} {{ Airfoil_AttachedRange_ClAlpha.png?nolink }}
  
 ==== Cl0 ==== ==== Cl0 ====
-The ''Cl0'' parameter increases the lift coefficient when the angle of attack is zero. This is achieved in real life by giving the airfoil a chamber. For symmetrical airfoils ''Cl0'' always equals ''0''.+The ''Cl0'' parameter increases the lift coefficient when the angle of attack is zero. This is achieved in real life by giving the airfoil a camber. For symmetrical airfoils ''Cl0'' always equals ''0''.
  
 > This attribute directly reflects the attitude in which an aircraft flies at high speed (low AOA). It also requires a larger down force from the horizontal stabilizer which increases the drag. It greatly affects the glide ratio. > This attribute directly reflects the attitude in which an aircraft flies at high speed (low AOA). It also requires a larger down force from the horizontal stabilizer which increases the drag. It greatly affects the glide ratio.
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 ==== CmAlpha ==== ==== CmAlpha ====
-The gradient of the momentum coefficient Cm is zero for symmetrical airfoils, positive for most typical airfoils that show movement of center of pressure (unstable) and negative for most airfoils with an s-bend (stable), used for flying wings. The ''CmAlpha'' and ''Cm0'' (see below) are related to the airfoils chamber+The gradient of the momentum coefficient Cm is zero for symmetrical airfoils, positive for most typical airfoils that show movement of center of pressure (unstable) and negative for most airfoils with an s-bend (stable), used for flying wings. The ''CmAlpha'' and ''Cm0'' (see below) are related to the airfoil's camber
 > The ''CmAlpha'' attribute influences the amount of elevator required for a level turn and therefor also the rate of turn. For gliders this value needs to be high to be be as unstable as possible and use as little elevator deflection as possible. > The ''CmAlpha'' attribute influences the amount of elevator required for a level turn and therefor also the rate of turn. For gliders this value needs to be high to be be as unstable as possible and use as little elevator deflection as possible.
 > Typical ''CmAlpha'' values are ''0.09'' for normal and ''0.0'' for symmetric airfoils. > Typical ''CmAlpha'' values are ''0.09'' for normal and ''0.0'' for symmetric airfoils.
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 ==== Cm0 ==== ==== Cm0 ====
-''Cm0'' is the equivalent to ''Cl0'' for the momentum coefficient. It is also zero for symmetrical airfoils. For chambered airfoils this value is usually smaller than zero which creates a pitch up tendency when flying fast (at low angle of attack). +''Cm0'' is the equivalent to ''Cl0'' for the momentum coefficient. It is also zero for symmetrical airfoils. For cambered airfoils this value is usually smaller than zero which creates a pitch up tendency when flying fast (at low angle of attack). 
 > With given geometry, incidence and fixed elevator position the ''Cm0'' value the ''Cm0'' attribute directly effects the pitch trim of the aircraft. Increase the value to get a pitch up tendency and decrease it to get a pitch down trim.  > With given geometry, incidence and fixed elevator position the ''Cm0'' value the ''Cm0'' attribute directly effects the pitch trim of the aircraft. Increase the value to get a pitch up tendency and decrease it to get a pitch down trim. 
  
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 ''ClAlphaBase'' is usually the maximum momentum coefficient an airfoil will ever create. ''ClAlphaBase'' is usually the maximum momentum coefficient an airfoil will ever create.
  
-> The lower the value of ''ClAlphaBase'' the more the nose will drop when approaching a stall. Values near ''-0.2'' tend to soften the stall behavior. +> The lower the value of ''CmAlphaBase'' the more the nose will drop when approaching a stall. Values near ''-0.2'' tend to soften the stall behavior. 
  
 {{ Airfoil_Stall_ClAlphaBase.png?nolink }} {{ Airfoil_Stall_ClAlphaBase.png?nolink }}
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 ==== Minimal ==== ==== Minimal ====
 This will fall back to default values for all attributes. This will fall back to default values for all attributes.
-<code>        <[string8][object][airfoil+<code>            <[airfoil][AirfoilTip][] 
-            <[string8][Name][AirfoilTip]> +            ></code>
-        ></code>+
                  
 ==== All ===== ==== All =====
-<code>        <[string8][object][airfoil] +<code>            <[airfoil][NACA0009][] 
-            <[string8][Name][AirfoilTip]> +                <[float64][Cl0][0.0]> 
-            <[float64][Cl0][0.0]> +                <[float64][Cd0][0.007]> 
-            <[float64][Cd0][0.007]> +                <[float64][Cm0][0.0]> 
-            <[float64][Cm0][0.0]> +                <[float64][ClAlpha][6.28]> 
-            <[float64][ClAlpha][6.28]> +                <[float64][CdAlpha][0.4]> 
-            <[float64][CdAlpha][0.4]> +                <[float64][CmAlpha][0.0]> 
-            <[float64][CmAlpha][0.0]> +                <[float64][ClFlap][2.1]> 
-            <[float64][ClFlap][2.1]> +                <[float64][CdFlap][0.05]> 
-            <[float64][CdFlap][0.0]> +                <[float64][CmFlap][-0.3]> 
-            <[float64][CmFlap][-0.2]> +                <[float64][Cd0Base][0.005]> 
-            <[float64][Cd0Base][0.005]> +                <[float64][ClAlphaBase][2.4]> 
-            <[float64][ClAlphaBase][2.4]> +                <[float64][CdAlphaBase][1.8]> 
-            <[float64][CdAlphaBase][1.8]> +                <[float64][CmAlphaBase][-0.45]> 
-            <[float64][CmAlphaBase][-0.2]> +                <[float64][AttachedCenter][0.0]> 
-            <[float64][ClFlapBase][0.5]> +                <[float64][AttachedRange][0.17]> 
-            <[float64][AttachedCenter][0.0]> +                <[float64][StallRange][0.1]> 
-            <[float64][AttachedRange][0.25]> +                <[float64][ClFlapBase][0.5]> 
-            <[float64][StallRange][0.25]> +                <[float64][AttachedCenterFlap][-0.12]> 
-            <[float64][AttachedCenterFlap][-0.12]> +            ></code>
-        ></code>+
  
aircraft/tmd/airfoil.1537437818.txt.gz · Last modified: 2018/09/20 12:03 by jh