looking for A330-300 oswald efficiency factor
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looking for A330-300 oswald efficiency factor
Hello.
im new at this forum so hi everyone!
well im having problems searching for the oswald efficiency factor of a A330-300, so if someone does know it post it please! it would be an awesome help :P
thanks!.
im new at this forum so hi everyone!
well im having problems searching for the oswald efficiency factor of a A330-300, so if someone does know it post it please! it would be an awesome help :P
thanks!.
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I believe at Mach .82 Oswald efficiency number is likely to be 0.2
Take a look at Raymer Daniels, Aircraft Design: A Conceptual Approach
Take a look at Raymer Daniels, Aircraft Design: A Conceptual Approach
Last edited by jimmyg; 25th Nov 2013 at 10:41.
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Yes you are right had to go back and check, Oswald efficiency does decrease with speed
Clean sub-sonic would be somewhere between 0.7 and 0.8
This function lets you use different methods to estimate Oswald efficiency factor, e, for purposes of approximating a simple parabolic drag polar:
C_D = C_D0 + C_L^2/(pi*e*AR), where C_L is lift coefficient.
Based on some or all of the following:
- Wing aspect ratio
- Wing sweep
- Parasite drag coefficient (drag independent of lift)
- Ratio of fuselage diameter (or width) to wing span
- Planform efficiency - usually .98 < u < 1
You can also choose between two common methods as proposed by Raymer ("Aicraft Design: A Conceptual Approach") or Shevell ("Fundamentals of Flight"). OSWALDFACTOR can also calculate using both methods and return the most pessimistic or optimistic result or a weighted average of the two results.
http://www.mathworks.com/matlabcentr...6/untitled.png
Clean sub-sonic would be somewhere between 0.7 and 0.8
This function lets you use different methods to estimate Oswald efficiency factor, e, for purposes of approximating a simple parabolic drag polar:
C_D = C_D0 + C_L^2/(pi*e*AR), where C_L is lift coefficient.
Based on some or all of the following:
- Wing aspect ratio
- Wing sweep
- Parasite drag coefficient (drag independent of lift)
- Ratio of fuselage diameter (or width) to wing span
- Planform efficiency - usually .98 < u < 1
You can also choose between two common methods as proposed by Raymer ("Aicraft Design: A Conceptual Approach") or Shevell ("Fundamentals of Flight"). OSWALDFACTOR can also calculate using both methods and return the most pessimistic or optimistic result or a weighted average of the two results.
http://www.mathworks.com/matlabcentr...6/untitled.png
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looking for B767-400ER Oswald efficiency coef
hello again!.
I posted yesterday asking for the oswald efficiency coef of a A330-300, and your replies were very helpful!.
now im looking for the same coef but of a B767-300ER, so again if someone does know it, please reply it would be very helpful!
thanks.
I posted yesterday asking for the oswald efficiency coef of a A330-300, and your replies were very helpful!.
now im looking for the same coef but of a B767-300ER, so again if someone does know it, please reply it would be very helpful!
thanks.
Last edited by pipox; 25th Nov 2013 at 23:49.
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This stuff goes way beyond my pay grade and as mentioned by Pugilistic Animus, not a simple mathematical formulation that goes into some weighty theory reserved for much more gifted minds than mine.
Different types of models have been applied with different random variables, theoretical, nonlinear regression and empirical.
In most cases it is a conservative estimate in aircraft design based on corrective factors in relation to optimal taper ratio for a given sweep angle. Which can give a theoretical Oswald factor. Then correction factors are applied for span loading, fuselage,compressibility, mach number and lift to drag characteristics. Basically for everything for non planar configuration.
In other words, based on whose model you apply and without specific wing/fuselage design information. I am able to only give a conservative estimate based on a speed just short of super sonic of somewhere between 0.6 and .8 You would need more data and a specific computer design program to generate better conclusions.
Try and contact Boeing or Airbus design engineering departments for more information.
Different types of models have been applied with different random variables, theoretical, nonlinear regression and empirical.
In most cases it is a conservative estimate in aircraft design based on corrective factors in relation to optimal taper ratio for a given sweep angle. Which can give a theoretical Oswald factor. Then correction factors are applied for span loading, fuselage,compressibility, mach number and lift to drag characteristics. Basically for everything for non planar configuration.
In other words, based on whose model you apply and without specific wing/fuselage design information. I am able to only give a conservative estimate based on a speed just short of super sonic of somewhere between 0.6 and .8 You would need more data and a specific computer design program to generate better conclusions.
Try and contact Boeing or Airbus design engineering departments for more information.
Last edited by jimmyg; 26th Nov 2013 at 12:21.
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You know when I think about 
I am not sure if our poster was setting bait on such a complex subject matter.
In all of mankind's prose endeavors, nothing proves nature's laws of aerodynamics more succinctly than does physics and mathematics.
I am not sure if our poster was setting bait on such a complex subject matter.In all of mankind's prose endeavors, nothing proves nature's laws of aerodynamics more succinctly than does physics and mathematics.
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Interesting paper! thanks for posting that...
Ventus, to break it down, the Oswald factor is used , along with a myriad of other calcs, to estimate the wing span, and wing span efficiency in aircraft design.
I did a broad brush look at it, especially with some of the assumptions, coupled with the data provided and the corresponding source.
I did find some interesting issues:
1. the inviscid part. The assumption that the induced velocities from the wake or vortex drag are inviscid is not currently accepted by everyone any longer... I note how much of the foundation of the calculations is based on the inviscid parameter.
2. Corrections for span loading due to the fuselage ignore centerwing parameters, it simply uses fuselage diameter. I feel this is a significant overgeneralization of the component couple.
3. Winglets. It appears that they use the assumption that winglet height has the same effect as span increase, with a variable. This is oversimplified to the point of being useless. There are far too many variables associated with winglets and when they are efficient.
4. They reference specifically Boeing data, so it comes as no surprise that the paper shows the 737-800 having more drag reduction at cruise that the A320 Neo....
Ventus, to break it down, the Oswald factor is used , along with a myriad of other calcs, to estimate the wing span, and wing span efficiency in aircraft design.
I did a broad brush look at it, especially with some of the assumptions, coupled with the data provided and the corresponding source.
I did find some interesting issues:
1. the inviscid part. The assumption that the induced velocities from the wake or vortex drag are inviscid is not currently accepted by everyone any longer... I note how much of the foundation of the calculations is based on the inviscid parameter.
2. Corrections for span loading due to the fuselage ignore centerwing parameters, it simply uses fuselage diameter. I feel this is a significant overgeneralization of the component couple.
3. Winglets. It appears that they use the assumption that winglet height has the same effect as span increase, with a variable. This is oversimplified to the point of being useless. There are far too many variables associated with winglets and when they are efficient.
4. They reference specifically Boeing data, so it comes as no surprise that the paper shows the 737-800 having more drag reduction at cruise that the A320 Neo....
For some of their (the authors of the above paper) compressibility issues, they would probably have gotten a better data fit if they used the Karman-Tsien rather than the Prandtl-Glauert, just my opinion though...
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When ventus 45 wrote of
I wasn’t sure if that was an invitation or a challenge, especially since I share jimmyg’s feelings a little.
But since it was an interesting paper and generated some ideas that others here might also find interesting I thought I would follow it through although some illusions might be shattered. My apologies to those who know all this already!
The paper contains a lot of complicated and off-putting equations, but when one gets down to it one finds that for most of the time the authors are simply repeating very similar thoughts in sixteen different ways. Worse, they spend this time on what is probably the least important bit of the drag due to lift and they apply an empirical correction for Mach number that simply doesn’t work above about 0.8M
I find it easier to explain and understand the problem using the Limey version of the expression for drag due to lift : Cdi = k*CL^2/(pi*A)
where the “induced drag factor” (k) is the reciprocal of the Oswald efficiency (e).
That way an increase in the coefficient corresponds to an increase in drag.
Most researchers assume the drag due to lift to be composed of two parts – the inviscid (theoretical) vortex drag and the viscous (practical) drag. Using this concept the induced drag factor is:
k = (1+delta)+ (pi*A)*dCd/dCL^2
1 is the much discussed but never achieved ideal vortex drag that goes with an elliptical lift distribution
delta is the increment to be applied to that ideal to allow for taper ratio, fuselage effects etc.
dCd/dCL^2 is to account for all the bits of viscous dependent drag that vary with lift coefficient.
If you apply the equations given in the paper to the airplanes listed in the Appendix (leaving out the propeller driven machines and the two labelled unreliable) the average (1+delta) is 1.07 and the average “k” deduced from the tabulated values of “e” is 1.35.
In other words the additional lift dependent drag coming from viscous effects is four times greater than that from less than ideal lift distribution and makes the concentration on inviscid drag terms and the finer points of taper and sweep variations rather academic.
So when pugilistic animus writes
one can only agree!
BTW underfire when you write
It is difficult to see how you arrive at this conclusion, since theA320 span efficiency is given as 0.783 and the B737-800 as 0.66
Consideration of the viscous effects is where the black art comes in, and where some of the classic assumptions start to go awry.
The paper mentions that “e” falls as wing t/c is increased, which is right because the flow over the rear part of the upper surface starts to separate earlier (starting from the TE) as CL increases. At the other end of the airfoil there is an area of very low pressure where the flow moves around from the stagnation point on the lower surface to go over the top (leading edge suction in the vernacular). If you can arrange to have a significant amount of forward facing surface in this zone you will get a “thrust” component of the aerodynamic force that lowers the effective drag coefficient. Airfoils with a large nose radius can exploit this, this wings with sharp leading edges cannot. If you examine the profile of modern supercritical wing sections you will find that they have large nose radii despite the fact that they are intended to fly at highish Mach numbers. Balancing the effects at either end of the airfoil and their influence on the structure is where the art comes in.
However, the fact that they are “supercritical” sections means that they have significant areas of supersonic flow over the upper surface, so the drag is not simply vortex plus some viscous effects – there is also wave drag to consider.
A modern airliner wing is designed using CFD methods to arrange camber, twist and thickness distributions to minimise drag at some defined design condition of Mach number and lift coefficient. That means balancing vortex drag, viscous drag and wave drag to get the best result. One consequence of this is that the sectional drag is not a minimum at zero lift coefficient. The old classic: Cd = Cdo + k*CL^2/pi.A is not accurate; instead one has to use Cd= Cdo + k*(CL-CLo)^2/pi.A. This in turn means that one cannot predict lift dependent drag using a simple value of “e” (or “k”); one must know the associated value of CLo as well.
Worth mentioning perhaps is that the intention is to design the best wing, not the best aerodynamic design. This means that it may be (usually is) better to back off a little from the ideal lift distribution to bring the centre of pressure further inboard which reduces the wing bending moment and weight.
The paper goes on to attempt a correction for Mach number effects, but if you calculate this using their expressions you will find that the span efficiency falls to zero at 0.825M for any airplane – not very useful if your design is meant to cruise at 0.84M!
BTW pugilistic animus, following up on your thoughts:
I tried Karman-Tsien and although Prandtll-Glauert was hopeless as a basis for Mach number correction, K-T was worse!
All in all, you can have your choice of methods to estimate “e” for Mach numbers up to about 0.7 (the Douglas Aeroplane Company method described in Dick Shevell’s book is good and easy to use) but above that, and for modern designs, you need a more sophisticated technique.
All of which is just a complicated way of agreeing with jimmyg:
And the best of luck
http://www.google.com.au/url?sa=t&rc...57155469,d.aGc
Now, most of that makes my head hurt, big time.
If you could study it, digest it, and then explain it to the rest of us, in simple terms, we would all be well pleased.
Now, most of that makes my head hurt, big time.
If you could study it, digest it, and then explain it to the rest of us, in simple terms, we would all be well pleased.
You know when I think about I am not sure if our poster was setting bait on such a complex subject matter.
The paper contains a lot of complicated and off-putting equations, but when one gets down to it one finds that for most of the time the authors are simply repeating very similar thoughts in sixteen different ways. Worse, they spend this time on what is probably the least important bit of the drag due to lift and they apply an empirical correction for Mach number that simply doesn’t work above about 0.8M
I find it easier to explain and understand the problem using the Limey version of the expression for drag due to lift : Cdi = k*CL^2/(pi*A)
where the “induced drag factor” (k) is the reciprocal of the Oswald efficiency (e).
That way an increase in the coefficient corresponds to an increase in drag.
Most researchers assume the drag due to lift to be composed of two parts – the inviscid (theoretical) vortex drag and the viscous (practical) drag. Using this concept the induced drag factor is:
k = (1+delta)+ (pi*A)*dCd/dCL^2
1 is the much discussed but never achieved ideal vortex drag that goes with an elliptical lift distribution
delta is the increment to be applied to that ideal to allow for taper ratio, fuselage effects etc.
dCd/dCL^2 is to account for all the bits of viscous dependent drag that vary with lift coefficient.
If you apply the equations given in the paper to the airplanes listed in the Appendix (leaving out the propeller driven machines and the two labelled unreliable) the average (1+delta) is 1.07 and the average “k” deduced from the tabulated values of “e” is 1.35.
In other words the additional lift dependent drag coming from viscous effects is four times greater than that from less than ideal lift distribution and makes the concentration on inviscid drag terms and the finer points of taper and sweep variations rather academic.
So when pugilistic animus writes
All in all aerodynamics is truly an art...because nature laughs at complex mathematics
one can only agree!
BTW underfire when you write
They reference specifically Boeing data, so it comes as no surprise that the paper shows the 737-800 having more drag reduction at cruise that the A320 Neo....
It is difficult to see how you arrive at this conclusion, since theA320 span efficiency is given as 0.783 and the B737-800 as 0.66
Consideration of the viscous effects is where the black art comes in, and where some of the classic assumptions start to go awry.
The paper mentions that “e” falls as wing t/c is increased, which is right because the flow over the rear part of the upper surface starts to separate earlier (starting from the TE) as CL increases. At the other end of the airfoil there is an area of very low pressure where the flow moves around from the stagnation point on the lower surface to go over the top (leading edge suction in the vernacular). If you can arrange to have a significant amount of forward facing surface in this zone you will get a “thrust” component of the aerodynamic force that lowers the effective drag coefficient. Airfoils with a large nose radius can exploit this, this wings with sharp leading edges cannot. If you examine the profile of modern supercritical wing sections you will find that they have large nose radii despite the fact that they are intended to fly at highish Mach numbers. Balancing the effects at either end of the airfoil and their influence on the structure is where the art comes in.
However, the fact that they are “supercritical” sections means that they have significant areas of supersonic flow over the upper surface, so the drag is not simply vortex plus some viscous effects – there is also wave drag to consider.
A modern airliner wing is designed using CFD methods to arrange camber, twist and thickness distributions to minimise drag at some defined design condition of Mach number and lift coefficient. That means balancing vortex drag, viscous drag and wave drag to get the best result. One consequence of this is that the sectional drag is not a minimum at zero lift coefficient. The old classic: Cd = Cdo + k*CL^2/pi.A is not accurate; instead one has to use Cd= Cdo + k*(CL-CLo)^2/pi.A. This in turn means that one cannot predict lift dependent drag using a simple value of “e” (or “k”); one must know the associated value of CLo as well.
Worth mentioning perhaps is that the intention is to design the best wing, not the best aerodynamic design. This means that it may be (usually is) better to back off a little from the ideal lift distribution to bring the centre of pressure further inboard which reduces the wing bending moment and weight.
The paper goes on to attempt a correction for Mach number effects, but if you calculate this using their expressions you will find that the span efficiency falls to zero at 0.825M for any airplane – not very useful if your design is meant to cruise at 0.84M!
BTW pugilistic animus, following up on your thoughts:
For some of their (the authors of the above paper) compressibility issues, they would probably have gotten a better data fit if they used the Karman-Tsien rather than the Prandtl-Glauert, just my opinion though..
I tried Karman-Tsien and although Prandtll-Glauert was hopeless as a basis for Mach number correction, K-T was worse!
All in all, you can have your choice of methods to estimate “e” for Mach numbers up to about 0.7 (the Douglas Aeroplane Company method described in Dick Shevell’s book is good and easy to use) but above that, and for modern designs, you need a more sophisticated technique.
All of which is just a complicated way of agreeing with jimmyg:
You would need more data and a specific computer design program to generate better conclusions.
Try and contact Boeing or Airbus design engineering departments for more information.
Try and contact Boeing or Airbus design engineering departments for more information.
Last edited by Owain Glyndwr; 30th Nov 2013 at 10:25. Reason: arithmetic!
