pawsephus
Agricultural
- Aug 28, 2005
- 1
I am trying to refine a wing design for a modern
version of an old WWI triplane type, a pet project similar to but not a replica of the Sopwith/Fokker types.
I've gone through all the old NACA reports and some old engineering texts of the time and later that referenced multi-wing design in an effort to learn as much as possible about wing theory and old research
in regards to triplanes.
Question 1.....
I've looked at Munk's report (NACA No. 256) where a particular wing cellule of a gap to chord ratio, G/C, of 0.9, and a forward stagger of 30°, each individual wing having an aspect ratio of 6.0 with a fairly thin RAF 15 airfoil, yielded a lift curve slope of 0.075 for the top wing, 0.050 for the middle wing and 0.045 for the bottom
wing, all originating at nearly the same zero lift.
Munk's data did not go much beyond the stall point of the top wing so I couldn't tell much about the maximum lift of the middle and bottom wing except that their lift was still reasonably linear at the last reported AOA.
There was very little difference in the lift and drag curves between the top wing and a monoplane wing of the same dimensions. In fact, Munks tests showed a very slight increase in lift of the triplane top wing and the mono-wing.
Running the RAF 15 profile as a triplane setup equivalent to Munk's through Hepperle's Java Foil as a multifoil yields very similar ratios of the lift curve slopes to Munk's. Doing the same with the NACA 6515 airfoil also yields similar ratios to Munk's.
Looking at the theory in the texts along with the above, I'm led to believe that if I use the same wing aspect ratio, the same G/C, and the same stagger, and a relatively thin profile such as the NACA 23012, that a reasonable approach would be to use the same lift curve
slopes as Munk. Is this a reasonable approach? Suggestions?
Questions 2,3,and 4....
I've about decided to go with a NACA 23012 airfoil with a 20%chord NACA 23012 external flap/aileron. This will allow some camber change flexibility when theories falter and some NACA data in Report 573 is available for projecting maximum lift and drag and moment at AR=6 and AR=infinite. I've put the lift curves for the various flap
settings in a spreadsheet, adjusted to the lift slopes above. I've taken these past the stall as shown in the report in order to look mainly at the top wing at and beyond stall from a drag standpoint.
This means the profile drag can't be related to the lift by a constant formula as the lift drops off as the profile drag increases beyond the stall. It seems to me that the profile drag should be related to the AoA rather than the lift. Is this right? Suggestions..?
Running the 23012 through Hepperle's Java Foil at surface condition "smooth" and again with surface condition "Paint and Fabric"allows me to figure a correction factor from smooth to real world. I
came up with nearly the same ratios using AeroFoil. Is this a reasonable way to adjust the NACA profile drag data? Shouldn't this varying correction factor be based on the Cl rather than AoA?
Best I can tell, Cmac and aerodynamic center adjustments for a triplane over a monoplane are small and fuzzy and should be masked by all the other bigger fuzz. Sound reasonable?
Thanks for any and all ideas, suggestions...
Dave
version of an old WWI triplane type, a pet project similar to but not a replica of the Sopwith/Fokker types.
I've gone through all the old NACA reports and some old engineering texts of the time and later that referenced multi-wing design in an effort to learn as much as possible about wing theory and old research
in regards to triplanes.
Question 1.....
I've looked at Munk's report (NACA No. 256) where a particular wing cellule of a gap to chord ratio, G/C, of 0.9, and a forward stagger of 30°, each individual wing having an aspect ratio of 6.0 with a fairly thin RAF 15 airfoil, yielded a lift curve slope of 0.075 for the top wing, 0.050 for the middle wing and 0.045 for the bottom
wing, all originating at nearly the same zero lift.
Munk's data did not go much beyond the stall point of the top wing so I couldn't tell much about the maximum lift of the middle and bottom wing except that their lift was still reasonably linear at the last reported AOA.
There was very little difference in the lift and drag curves between the top wing and a monoplane wing of the same dimensions. In fact, Munks tests showed a very slight increase in lift of the triplane top wing and the mono-wing.
Running the RAF 15 profile as a triplane setup equivalent to Munk's through Hepperle's Java Foil as a multifoil yields very similar ratios of the lift curve slopes to Munk's. Doing the same with the NACA 6515 airfoil also yields similar ratios to Munk's.
Looking at the theory in the texts along with the above, I'm led to believe that if I use the same wing aspect ratio, the same G/C, and the same stagger, and a relatively thin profile such as the NACA 23012, that a reasonable approach would be to use the same lift curve
slopes as Munk. Is this a reasonable approach? Suggestions?
Questions 2,3,and 4....
I've about decided to go with a NACA 23012 airfoil with a 20%chord NACA 23012 external flap/aileron. This will allow some camber change flexibility when theories falter and some NACA data in Report 573 is available for projecting maximum lift and drag and moment at AR=6 and AR=infinite. I've put the lift curves for the various flap
settings in a spreadsheet, adjusted to the lift slopes above. I've taken these past the stall as shown in the report in order to look mainly at the top wing at and beyond stall from a drag standpoint.
This means the profile drag can't be related to the lift by a constant formula as the lift drops off as the profile drag increases beyond the stall. It seems to me that the profile drag should be related to the AoA rather than the lift. Is this right? Suggestions..?
Running the 23012 through Hepperle's Java Foil at surface condition "smooth" and again with surface condition "Paint and Fabric"allows me to figure a correction factor from smooth to real world. I
came up with nearly the same ratios using AeroFoil. Is this a reasonable way to adjust the NACA profile drag data? Shouldn't this varying correction factor be based on the Cl rather than AoA?
Best I can tell, Cmac and aerodynamic center adjustments for a triplane over a monoplane are small and fuzzy and should be masked by all the other bigger fuzz. Sound reasonable?
Thanks for any and all ideas, suggestions...
Dave
