this is a polar cord file of three different airfoils , SD6062 7.5%,RG15 7% and FX791217operating with a 7.5 in cord at 300mphthe FX is the clear winner
i understand how ever ,that is the reality of the flight profile , that's the challenge eventually the Gs will get so high that you have to fly with an increasingly high Cl that you end up burning all your energy in the turns because of the increasing Cd and that's as fast as you can go . that's the barrier to going faster. btw love your formula . can you tell me what rho and nu are at sea lvl?
oops I used 387mph for the calcs, so reduce everything by a third.
Reynolds Number is the ratio of relative velocity to kinematic viscosity, or in equation form : (rho*V*L)/nu where rho is density, V is velocity, L is a characteristic length, and nu is dynamic viscosity (nu/rho is kinematic viscosity). It works using SI units or Psf (pounds seconds feet) but if you use the imperial measures you end up with density in slugs/f^3, and wtf is a slug?
Anyway, my main point is that a Cl of 0.8 is way outside the realm of efficient flight for any high-speed airfoil.
were did you get a Rn of 4x10^4 , or 40000 ? my formula is this for Rn (from Andy lennon's R/C model aircraft design)
Rn = mph x cord(inches) x K , K = 780 @ sea lvl
witch gives me a Rn of 1755000 @ mach 0.394
the supercritical airfoils will cause too much drag in the corners the name of the game is low drag while cornering im looking at the fx foil because it was designed for pylon racers . if you can show me an airfoil that gets a lower Cd at a Cl of 0.8 or higher let me know. thanks for the foil numbers i will take a look at them.im just in design phase so any input is appreciated
the reason for the high cl is this 2lbs glider is pulling 50g in the turns so you have to calculate your needed cl for 100lbs . my calcs show 0.84 needed cl to produce 100lbs of lift @ 300mph . and yes i am kicking around the idea of longer cord ect . i just picked up a sig ninja and should be hitting the slope next week .
PS, there are some specialist high-lift airfoil 66(421)-420 which can get L/Ds of up to 250, if you don't mind operating on the alpha wall (nearly stalled) all the time, but these wouldn't be suitable for your task, they're all very thick and you'd be seeing mach effects from M0.5ish.
There are two major sources of noise from a wing, the flow patterns (i.e. turbulence patterns) which are related to reynolds number and coefficient of lift and mach effects, which are related to mach number and coefficient of lift. The reynolds number is quiet low (4x10^4) and so is the mach number (0.687). However you're working at an exceedingly high coefficient of lift so it could be either.
Secondly is there any reason you're working at such a high Cl? is there a wing area limit? If not, you should increase wing area (probably cord, as more span at such high g is likely to be bad). You need to be operating at the best lift-to drag point for the airfoil. For example, a NACA 65-406 (high speed laminar flow supercritical type, 6% thick, med camber) has a best section L/D of apprx 133 at a Cl of .55, whereas the design point you've chosen has a section L/D of 120ish. So you may be better off picking a different airfoil, using a lower Cl to get a better L/D and increasing your wing size for lower total drag.
exactly if you watch the 357 mph video you will here one of the people say "its making a whole new noise" , this was happening around 320 from the sound of it , i think it was starting to delaminate the hope is we will be able to pull more G with out delam'.
btw they were useing the RG15 7% in that video
these are outputs from XFOIL ,a 2-D computational fluid dynamics program , these are theoretical performance numbers basically the name of the game in DS is to maintain a low Cd at high Cl , the large chart to the left illustrates this . so ... at 50g a 2 lb glider weighs 100 lbs in order to do this at 300mph the wing needs to produce a Cl of around 0.8 so if you look at the Cl to Cd graph the fx when producing a 0.8 Cl has the lowest Cd
FYI
Cl = coefficient of lift
Cd = coefficient of drag
Comments
Reynolds Number is the ratio of relative velocity to kinematic viscosity, or in equation form : (rho*V*L)/nu where rho is density, V is velocity, L is a characteristic length, and nu is dynamic viscosity (nu/rho is kinematic viscosity). It works using SI units or Psf (pounds seconds feet) but if you use the imperial measures you end up with density in slugs/f^3, and wtf is a slug?
Anyway, my main point is that a Cl of 0.8 is way outside the realm of efficient flight for any high-speed airfoil.
Rn = mph x cord(inches) x K , K = 780 @ sea lvl
witch gives me a Rn of 1755000 @ mach 0.394
the supercritical airfoils will cause too much drag in the corners the name of the game is low drag while cornering im looking at the fx foil because it was designed for pylon racers . if you can show me an airfoil that gets a lower Cd at a Cl of 0.8 or higher let me know. thanks for the foil numbers i will take a look at them.im just in design phase so any input is appreciated
the reason for the high cl is this 2lbs glider is pulling 50g in the turns so you have to calculate your needed cl for 100lbs . my calcs show 0.84 needed cl to produce 100lbs of lift @ 300mph . and yes i am kicking around the idea of longer cord ect . i just picked up a sig ninja and should be hitting the slope next week .
Secondly is there any reason you're working at such a high Cl? is there a wing area limit? If not, you should increase wing area (probably cord, as more span at such high g is likely to be bad). You need to be operating at the best lift-to drag point for the airfoil. For example, a NACA 65-406 (high speed laminar flow supercritical type, 6% thick, med camber) has a best section L/D of apprx 133 at a Cl of .55, whereas the design point you've chosen has a section L/D of 120ish. So you may be better off picking a different airfoil, using a lower Cl to get a better L/D and increasing your wing size for lower total drag.
btw they were useing the RG15 7% in that video
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930092747_19...
FYI
Cl = coefficient of lift
Cd = coefficient of drag
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