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Wing planform complete but aerofoil adding in and solidification problem

14 REPLIES 14
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Message 1 of 15
inaamh
345 Views, 14 Replies

Wing planform complete but aerofoil adding in and solidification problem

The wing planform is perfectly done, so please don't change the geometry on this. Now how do I add the aerofoil cross-sections to the root and tip chords? I've spent hours getting nowhere on this!
 

 

This is the aerofoil I am adding to the root and tip chords. But how do I add the airfoils in to the root and tip chords? And how do I close the gap between the last two points of the trailing edge in this aerofoil?

 

And if you're done, please send me an updated and corrected fusion 360 file of the solution of the solid wing body with the same cross-sectional aerofoils at the root and tip chords.

These are my xy aerofoil coordinates for the supercritical aerofoil for the cross-sections at the tip and root chords:

1-0.0104
0.99-0.0071
0.98-0.0039
0.97-0.0009
0.950.0049
0.920.0131
0.90.0181
0.870.0251
0.850.0294
0.820.0353
0.80.0389
0.770.0439
0.750.0469
0.720.0509
0.70.0533
0.680.0555
0.650.0585
0.620.061
0.60.0625
0.570.0645
0.550.0656
0.530.0666
0.50.0678
0.480.0684
0.450.0692
0.430.0695
0.40.0697
0.380.0698
0.350.0696
0.330.0692
0.30.0685
0.270.0673
0.250.0664
0.220.0646
0.20.0632
0.170.0606
0.150.0585
0.120.0548
0.10.0518
0.070.0462
0.050.0411
0.040.0381
0.030.0343
0.020.0293
0.010.0219
0.0050.0158
0.0020.0095
00
0.002-0.0093
0.005-0.016
0.01-0.0221
0.02-0.0295
0.03-0.0344
0.04-0.0381
0.05-0.0412
0.07-0.0462
0.1-0.0517
0.12-0.0547
0.15-0.0585
0.17-0.0606
0.2-0.0633
0.22-0.0647
0.25-0.0666
0.28-0.068
0.3-0.0687
0.32-0.0692
0.35-0.0696
0.37-0.0696
0.4-0.0692
0.42-0.0688
0.45-0.0676
0.48-0.0657
0.5-0.0644
0.53-0.0614
0.55-0.0588
0.58-0.0543
0.6-0.0509
0.63-0.0451
0.65-0.041
0.68-0.0346
0.7-0.0302
0.73-0.0235
0.75-0.0192
0.77-0.015
0.8-0.0093
0.83-0.0048
0.85-0.0024
0.87-0.0013
0.89-0.0008
0.92-0.0016
0.94-0.0035
0.95-0.0049
0.96-0.0066
0.97-0.0085
0.98-0.0109
0.99-0.0137
1-0.0163
14 REPLIES 14
Message 2 of 15
davebYYPCU
in reply to: inaamh

The Airfoil File needs the Z coordinate for each point. (way too many Points)

Saved as CSV, and the Spline from CSV add-in or script, will draw the airfoil on vertical plane from the Chord line.

 

For model aircraft there is no such thing as critical airfoil data. 

For best Fusion results you only need 4 point spline for top and bottom curves.

 

However, I still had your original step file, so I modified it -

I pasted your sketch into it.  (you should have it fully defined for accuracy.)

Moved the airfoil plates to the sketch trailing edge/s (presumed no wash out) and Lofted the half wing.

 

 

Might help...

 

 

Message 3 of 15
inaamh
in reply to: davebYYPCU

Send me the fusion 360 file too. Please. Otherwise I wouldn’t be able to learn from this all. I’ll be able to see the dimensions, and steps taken to the solution.
Message 4 of 15
inaamh
in reply to: inaamh

I didn’t want you to change my planform area, I just wanted you to add the aerofoils to the root and tip chords of the wing.
Message 5 of 15
davebYYPCU
in reply to: inaamh

I have not changed anything in the sketch.  (Cut and Pasted from your new file to the old one)

Something wrong with that sketch?

It was / is not fully defined.

I have only used one point in that sketch to position the tip rib.

My file was supplied.  Been fixed.

 

Might help....

Message 6 of 15
inaamh
in reply to: davebYYPCU

These are my xy aerofoil coordinates for the supercritical aerofoil for the cross-sections at the tip and root chords:

1-0.0104
0.99-0.0071
0.98-0.0039
0.97-0.0009
0.950.0049
0.920.0131
0.90.0181
0.870.0251
0.850.0294
0.820.0353
0.80.0389
0.770.0439
0.750.0469
0.720.0509
0.70.0533
0.680.0555
0.650.0585
0.620.061
0.60.0625
0.570.0645
0.550.0656
0.530.0666
0.50.0678
0.480.0684
0.450.0692
0.430.0695
0.40.0697
0.380.0698
0.350.0696
0.330.0692
0.30.0685
0.270.0673
0.250.0664
0.220.0646
0.20.0632
0.170.0606
0.150.0585
0.120.0548
0.10.0518
0.070.0462
0.050.0411
0.040.0381
0.030.0343
0.020.0293
0.010.0219
0.0050.0158
0.0020.0095
00
0.002-0.0093
0.005-0.016
0.01-0.0221
0.02-0.0295
0.03-0.0344
0.04-0.0381
0.05-0.0412
0.07-0.0462
0.1-0.0517
0.12-0.0547
0.15-0.0585
0.17-0.0606
0.2-0.0633
0.22-0.0647
0.25-0.0666
0.28-0.068
0.3-0.0687
0.32-0.0692
0.35-0.0696
0.37-0.0696
0.4-0.0692
0.42-0.0688
0.45-0.0676
0.48-0.0657
0.5-0.0644
0.53-0.0614
0.55-0.0588
0.58-0.0543
0.6-0.0509
0.63-0.0451
0.65-0.041
0.68-0.0346
0.7-0.0302
0.73-0.0235
0.75-0.0192
0.77-0.015
0.8-0.0093
0.83-0.0048
0.85-0.0024
0.87-0.0013
0.89-0.0008
0.92-0.0016
0.94-0.0035
0.95-0.0049
0.96-0.0066
0.97-0.0085
0.98-0.0109
0.99-0.0137
1-0.0163

 

This is the precise xy aerofoil coordinates I want in my wing cross section as this one is supercritical with a flat upper surface at the trailing edge and the curve cambered curve at the lower surface of the trailing edge, making it suitable for transonic speeds.

 

 

Not the previous one, as that one the NACA 65A004 is symmetrical, and not suited for transonic speeds (see image below).

 

inaamh_0-1707033535645.jpeg

 

Message 7 of 15
inaamh
in reply to: inaamh

Please @davebYYPCU , I’m so close. The wing planform I recently sent you in the fusion 360 file dimensions are correct. And the only thing left is the supercritical aerofoil cross section at both the root and tip chords.

 

These are my xy aerofoil coordinates for the supercritical aerofoil for the cross-sections at the tip and root chords:

1-0.0104
0.99-0.0071
0.98-0.0039
0.97-0.0009
0.950.0049
0.920.0131
0.90.0181
0.870.0251
0.850.0294
0.820.0353
0.80.0389
0.770.0439
0.750.0469
0.720.0509
0.70.0533
0.680.0555
0.650.0585
0.620.061
0.60.0625
0.570.0645
0.550.0656
0.530.0666
0.50.0678
0.480.0684
0.450.0692
0.430.0695
0.40.0697
0.380.0698
0.350.0696
0.330.0692
0.30.0685
0.270.0673
0.250.0664
0.220.0646
0.20.0632
0.170.0606
0.150.0585
0.120.0548
0.10.0518
0.070.0462
0.050.0411
0.040.0381
0.030.0343
0.020.0293
0.010.0219
0.0050.0158
0.0020.0095
00
0.002-0.0093
0.005-0.016
0.01-0.0221
0.02-0.0295
0.03-0.0344
0.04-0.0381
0.05-0.0412
0.07-0.0462
0.1-0.0517
0.12-0.0547
0.15-0.0585
0.17-0.0606
0.2-0.0633
0.22-0.0647
0.25-0.0666
0.28-0.068
0.3-0.0687
0.32-0.0692
0.35-0.0696
0.37-0.0696
0.4-0.0692
0.42-0.0688
0.45-0.0676
0.48-0.0657
0.5-0.0644
0.53-0.0614
0.55-0.0588
0.58-0.0543
0.6-0.0509
0.63-0.0451
0.65-0.041
0.68-0.0346
0.7-0.0302
0.73-0.0235
0.75-0.0192
0.77-0.015
0.8-0.0093
0.83-0.0048
0.85-0.0024
0.87-0.0013
0.89-0.0008
0.92-0.0016
0.94-0.0035
0.95-0.0049
0.96-0.0066
0.97-0.0085
0.98-0.0109
0.99-0.0137
1-0.0163

 

This is the precise xy aerofoil coordinates I want in my wing cross section as this one is supercritical with a flat upper surface at the trailing edge and the curve cambered curve at the lower surface of the trailing edge, making it suitable for transonic speeds.

 

 

Not the previous one, as that one the NACA 65A004 is symmetrical, and not suited for transonic speeds (see image below).

 

inaamh_0-1707037287154.jpeg

 

 

 

Message 8 of 15
inaamh
in reply to: inaamh

Just how do I change the aerofoil cross sections at both root and tip chords to these coordinates @davebYYPCU?

 

These are my xy aerofoil coordinates for the supercritical aerofoil for the cross-sections at the tip and root chords:

1-0.0104
0.99-0.0071
0.98-0.0039
0.97-0.0009
0.950.0049
0.920.0131
0.90.0181
0.870.0251
0.850.0294
0.820.0353
0.80.0389
0.770.0439
0.750.0469
0.720.0509
0.70.0533
0.680.0555
0.650.0585
0.620.061
0.60.0625
0.570.0645
0.550.0656
0.530.0666
0.50.0678
0.480.0684
0.450.0692
0.430.0695
0.40.0697
0.380.0698
0.350.0696
0.330.0692
0.30.0685
0.270.0673
0.250.0664
0.220.0646
0.20.0632
0.170.0606
0.150.0585
0.120.0548
0.10.0518
0.070.0462
0.050.0411
0.040.0381
0.030.0343
0.020.0293
0.010.0219
0.0050.0158
0.0020.0095
00
0.002-0.0093
0.005-0.016
0.01-0.0221
0.02-0.0295
0.03-0.0344
0.04-0.0381
0.05-0.0412
0.07-0.0462
0.1-0.0517
0.12-0.0547
0.15-0.0585
0.17-0.0606
0.2-0.0633
0.22-0.0647
0.25-0.0666
0.28-0.068
0.3-0.0687
0.32-0.0692
0.35-0.0696
0.37-0.0696
0.4-0.0692
0.42-0.0688
0.45-0.0676
0.48-0.0657
0.5-0.0644
0.53-0.0614
0.55-0.0588
0.58-0.0543
0.6-0.0509
0.63-0.0451
0.65-0.041
0.68-0.0346
0.7-0.0302
0.73-0.0235
0.75-0.0192
0.77-0.015
0.8-0.0093
0.83-0.0048
0.85-0.0024
0.87-0.0013
0.89-0.0008
0.92-0.0016
0.94-0.0035
0.95-0.0049
0.96-0.0066
0.97-0.0085
0.98-0.0109
0.99-0.0137
1-0.0163

 

This is the precise xy aerofoil coordinates I want in my wing cross section as this one is supercritical with a flat upper surface at the trailing edge and the curve cambered curve at the lower surface of the trailing edge, making it suitable for transonic speeds.

 

 

 

 

 

Message 9 of 15
davebYYPCU
in reply to: inaamh

I will need the airfoil data file to add the 3rd row of numbers.  Not sure but might have to be saved as txt file.

So add both to a message.

 

Transonic, on a model wing?

 

Might help....

Message 10 of 15
inaamh
in reply to: davebYYPCU

The third row, z axis is 0 for each points.

Message 11 of 15
davebYYPCU
in reply to: inaamh

I know.

 

Fusion needs the file with 3 numbers for each point saved in a CSV file, in a spreadsheet format.

I can add the zero but I can't use a picture of the numbers.

 

 

Might help...

Message 12 of 15
inaamh
in reply to: davebYYPCU

Will this suffice (though the trailing edge isn't closed)? @davebYYPCU 

Message 13 of 15
davebYYPCU
in reply to: inaamh

Looks like you have the 3 set numbers, will check it out later in my day.

 

Might help….

Message 14 of 15
davebYYPCU
in reply to: inaamh

Your CSV is ok.

 

Fusion can use it, and makes such a mess, that I would not be wasting time on it.

I don't have any transonic experience, you sure this airfoil is suitable?

 

ycsv1.PNG

 

The workflow to make your wing with this airfoil is.

Utilities > Scripts and Addins > ImportSplineCSV < Run.

Edit the sketch, join the open gap at trailing edge with a line.

 

New Sketch - Project the Trailing Edge Line, the Datum point and the other points as needed.

 

ycsv2.PNG

 

Fit Point Spline command, snap the spline to the top purple points, 

Constrain and adjust the handles to have the curve match sketch 2.

Fit Point Spline command, snap the spline to the bottom purple points, 

Constrain and adjust the handles to have the curve match sketch 2.

 

ycsv3.PNG

 

Surface > Patch . select the sketch 3 profile.

Modify > Scale > Uniform, scale the patch body to be 180mm long. (scale factor 18)

Move - Patch body - the datum point to the Leading edge point in sketch 1.

Sweep with Path and Guide Rail - Profile - select the Patch body, Path is Leading Edge line, and Rail is Trailing edge Line.  Set to full Extent, and Scale.

 

ycsv4.PNG

 

Might help....

 

Message 15 of 15
inaamh
in reply to: davebYYPCU

This solution is admirable and useful. Thanks a bunch Dave. Can you send me a youtube video of how you did it? It will help me learn how to make the wing design along with the aerofoil csv coordinates at the root and tip chords of the wing. It will help me learn as well.

 

All the best Dave,

Inaam Hussain

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