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aircraft:tmd:airfoil [2017/03/02 11:09] 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]], [[bodyaerodynamics]]+[[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 attached range is the range of the angle of attack at which the airflow is simulated as attached to the airfoil. It reaches from ''-AttachedRange'' to ''+AttachedRange'' The attached range is the range of the angle of attack at which the airflow is simulated as attached to the airfoil. It reaches from ''-AttachedRange'' to ''+AttachedRange''
  
-=== AttachedCenter ===+==== AttachedCenter ====
  
 The attached center creates an asymmetric attached center. Using an attribute grater than zero will offset the stall angle towards a greater angle of attack. The attached center creates an asymmetric attached center. Using an attribute grater than zero will offset the stall angle towards a greater angle of attack.
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 {{ Airfoil_Ranges2.png?nolink }} {{ Airfoil_Ranges2.png?nolink }}
  
-=== ClAlpha ===+==== ClAlpha ====
 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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 {{ Airfoil_AttachedRange_Cl0.png?nolink }} {{ Airfoil_AttachedRange_Cl0.png?nolink }}
  
-=== 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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 {{ Airfoil_AttachedRange_CmAlpha.png?nolink }} {{ Airfoil_AttachedRange_CmAlpha.png?nolink }}
  
-=== 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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 {{ Airfoil_AttachedRange_Cm0.png?nolink }} {{ Airfoil_AttachedRange_Cm0.png?nolink }}
  
-=== Cd0 ===+==== Cd0 ====
 ''Cd0'' is the drag coefficient at alpha = 0. For a symmetrical airfoil and when ''Cd0Base = 0'' and ''AttachedCenter = 0'' then ''Cd0'' is the lowest drag an airfoil ever produces. To either side of that angle, while in the attached range, the drag grows with a parabolic function.  ''Cd0'' is the drag coefficient at alpha = 0. For a symmetrical airfoil and when ''Cd0Base = 0'' and ''AttachedCenter = 0'' then ''Cd0'' is the lowest drag an airfoil ever produces. To either side of that angle, while in the attached range, the drag grows with a parabolic function. 
 > This attribute determines the drag at high speed (low AOA). It effects the glide ratio.  > This attribute determines the drag at high speed (low AOA). It effects the glide ratio. 
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 > Values can get as low as ''0.004'' for laminar glider airfoils. It is mostly a function of the airfoils thickness and aerodynamically quality. Airfoils for model aircraft have values as high as ''0.012''. > Values can get as low as ''0.004'' for laminar glider airfoils. It is mostly a function of the airfoils thickness and aerodynamically quality. Airfoils for model aircraft have values as high as ''0.012''.
  
-=== CdAlpha ===+==== CdAlpha ====
 The drag an airfoil produces due to the effective front area and the beginning of the airflow detachment is modeled with the ''CdAlpha'' attribute. This value does not take the induced drag into account because that is modeled entirely with inside the AeroWing class. The drag an airfoil produces due to the effective front area and the beginning of the airflow detachment is modeled with the ''CdAlpha'' attribute. This value does not take the induced drag into account because that is modeled entirely with inside the AeroWing class.
 > Leave the ''CdAlpha'' at ''0.4'' at the beginning > Leave the ''CdAlpha'' at ''0.4'' at the beginning
  
-=== Cd0Base ===+==== Cd0Base ====
 The ''Cd0Base'' increases the separation drag of the airflow. With higher values the stall produces a lot more drag. The ''Cd0Base'' increases the separation drag of the airflow. With higher values the stall produces a lot more drag.
  
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 {{ Airfoil_AttachedRange_CdAlpha.png?nolink }} {{ Airfoil_AttachedRange_CdAlpha.png?nolink }}
  
-==== Flaps ==== +==== AttachedCenterFlap ====
-=== AttachedCenterFlap ===+
 The ''AttachedCenterFlap'' attribute defines how much the local angle of attack is moved when the flap is deflected. With a negative value the airfoil stalls at a lower angle of attack with a deflected flap than in a clean configuration.  The ''AttachedCenterFlap'' attribute defines how much the local angle of attack is moved when the flap is deflected. With a negative value the airfoil stalls at a lower angle of attack with a deflected flap than in a clean configuration. 
 > Leave it at ''-0.12'' first.  > Leave it at ''-0.12'' first. 
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 > Set a value that is lower to increase the effect of reversed ailerons near stall. > Set a value that is lower to increase the effect of reversed ailerons near stall.
  
-=== ClFlap ===+==== ClFlap ====
 ''ClFlap'' represents the increase of the lift coefficient per ''1 rad ≈ 57.296°'' flap deflection. While the default value is ''1.5'' it is usually a different value in reality. A comparison [[#TableFlap|table]] for different kinds of flap types can be found later in this section.  ''ClFlap'' represents the increase of the lift coefficient per ''1 rad ≈ 57.296°'' flap deflection. While the default value is ''1.5'' it is usually a different value in reality. A comparison [[#TableFlap|table]] for different kinds of flap types can be found later in this section. 
 > ''2.1'' is an effective flap and a good value to start with. Use higher values for Fowler flaps and lower values for split flaps.  > ''2.1'' is an effective flap and a good value to start with. Use higher values for Fowler flaps and lower values for split flaps. 
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 {{ Airfoil_AttachedRange_ClFlap.png?nolink }} {{ Airfoil_AttachedRange_ClFlap.png?nolink }}
  
-=== CmFlap ===+==== CmFlap ====
 ''CmFlap'' simulates the amount of pitch up tendency of a flap. A negative value represents a pitch down effect. The value is usually negative. Also refer to the [[#TableFlap|table]] below for different ''CmFlap'' values for different flap types.  ''CmFlap'' simulates the amount of pitch up tendency of a flap. A negative value represents a pitch down effect. The value is usually negative. Also refer to the [[#TableFlap|table]] below for different ''CmFlap'' values for different flap types. 
 > Some aircraft need tremendous amount of negative ''CmFlap'' (''-0.7'' or even lower) to maintain level flight when extending the flaps.  > Some aircraft need tremendous amount of negative ''CmFlap'' (''-0.7'' or even lower) to maintain level flight when extending the flaps. 
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 {{ Airfoil_AttachedRange_CmFlap.png?nolink }} {{ Airfoil_AttachedRange_CmFlap.png?nolink }}
  
-=== CdFlap ===+==== CdFlap ====
 ''CdFlap'' is the increase of the drag coefficient per ''1 rad ≈ 57.296°'' flap deflection. ''CdFlap'' must be positive. The default value is ''0'' which means the flap does not create additional drag other than the induced drag caused by its lift coefficient ''ClFlap'' and its deflection angle (AdverseYaw for ailerons). Refer to the [[#TableFlap|table]] below for some examples. ''CdFlap'' is the increase of the drag coefficient per ''1 rad ≈ 57.296°'' flap deflection. ''CdFlap'' must be positive. The default value is ''0'' which means the flap does not create additional drag other than the induced drag caused by its lift coefficient ''ClFlap'' and its deflection angle (AdverseYaw for ailerons). Refer to the [[#TableFlap|table]] below for some examples.
 > Real life flaps always create additional drag when deflected. ''0.005'' or a little bit higher even is a good value to start with. Some large flaps like double Fowler flaps create more drag and this drag is simulated with a greater value.  > Real life flaps always create additional drag when deflected. ''0.005'' or a little bit higher even is a good value to start with. Some large flaps like double Fowler flaps create more drag and this drag is simulated with a greater value. 
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 >Increasing the ''CdFlap'' can reduce the AdverseYaw. >Increasing the ''CdFlap'' can reduce the AdverseYaw.
  
-<div id = "TableFlap"> 
 ==== Table of flap coefficients ==== ==== Table of flap coefficients ====
  
-> These are just examples+Recommended values for the Aerofly are:
  
-^ Type of flap ^ ClFlap       ^ CdFlap          ^ CmFlap           +^ Type of flap ^ ClFlap      ^ CdFlap   ^ CmFlap         
-| Plain flap   0.7 ... 1.5  | 0.005 ... 0.06  | -0.15 ... -0.7   +| Plain flap   2.1         | 0.0      | -0.... -0.5  
-| Split flap   0.7          | 0.18            | -0.19            +| Split flap   2.4         | 0.2      | -0.2 ... -0.9  
-| Slotted flap | 0.8 ... 2.0  | 0.06            | -0.14            +| Slotted flap | 2.7         | 0.06     | -0.2           
-| Fowler flap  | 1.... 3.0  | 0.14 ... 0.3    | -0.73            | +| Fowler flap  | 2.... 3.4 | 0.2      -0.... -1.1  |
-</div>+
  
-==== StallRange and beyond ==== +==== ClAlphaBase ====
- +
-=== ClAlphaBase ===+
 The lift coefficient can be approximated with a sinus function with the wave length of ''180° = pi''. Only in the AttachedRange and at -180° the airflow significantly deviates.  The lift coefficient can be approximated with a sinus function with the wave length of ''180° = pi''. Only in the AttachedRange and at -180° the airflow significantly deviates. 
 ''ClAlphaBase'' is the amplitude of the sin function and therefor the maximum lift that the airfoil can produce in the stall.  ''ClAlphaBase'' is the amplitude of the sin function and therefor the maximum lift that the airfoil can produce in the stall. 
 > Decrease this value to get a greater wing drop in a stall and during spins. Lower values increase the dynamics of accelerated stalls and maneuvers. Also decrease CdAlphaBase. > Decrease this value to get a greater wing drop in a stall and during spins. Lower values increase the dynamics of accelerated stalls and maneuvers. Also decrease CdAlphaBase.
  
-=== CmAlphaBase ===+==== CmAlphaBase ====
 Follows the same principle as the ClAlphaBase attribute. Usually this value is negative and the airfoil creates a pitch down force when the angle of attack increases beyond the stall angle (stable). The aircraft will pitch down on its own and accelerate again. A positive value is destabilizing. Follows the same principle as the ClAlphaBase attribute. Usually this value is negative and the airfoil creates a pitch down force when the angle of attack increases beyond the stall angle (stable). The aircraft will pitch down on its own and accelerate again. A positive value is destabilizing.
 ''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 ''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 }}
  
-=== ClFlapBase ===+==== ClFlapBase ====
 The ''ClFlapBase'' is the increment in lift coefficient in full stall. It adds to the ''ClAlphaBase'' when the flap is deflected.  The ''ClFlapBase'' is the increment in lift coefficient in full stall. It adds to the ''ClAlphaBase'' when the flap is deflected. 
 > Use the value to regulate the control surface authority in stall or at a high deflection angle. This increases the roll rate for aerobatic airplanes. (See ''ClFlap'' above as well) > Use the value to regulate the control surface authority in stall or at a high deflection angle. This increases the roll rate for aerobatic airplanes. (See ''ClFlap'' above as well)
  
-=== CdAlphaBase ===+==== CdAlphaBase ====
 The drag coefficient can be approximated with a sin^2 (sinus squared) function that repeats every ''180° = pi''. As for the ClAlphaBase the drag of an airfoil only deviates from this approximation when the airflow is attached, which is of course the case inside the AttachedRange but also at -180°. ''CdAlphaBase'' is the amplitude of that sin function and therefor the maximum lift that an airfoil can ever create.  The drag coefficient can be approximated with a sin^2 (sinus squared) function that repeats every ''180° = pi''. As for the ClAlphaBase the drag of an airfoil only deviates from this approximation when the airflow is attached, which is of course the case inside the AttachedRange but also at -180°. ''CdAlphaBase'' is the amplitude of that sin function and therefor the maximum lift that an airfoil can ever create. 
 > Decrease the ''CdAlphaBase'' to lose less energy in accelerated maneuvers and keep the momentum going.  > Decrease the ''CdAlphaBase'' to lose less energy in accelerated maneuvers and keep the momentum going. 
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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.1488449368.txt.gz · Last modified: 2017/03/02 11:09 by jh