A • TRANSITION FROM LAMINAR TO TURBULENT FLOW 



was determined by the surface-Pitot technique, but boundary layer data 

 were not taken and a small separation bubble might possibly be over- 

 looked. During the same period the laminar flow airfoils were being de- 

 veloped, permitting much more extensive runs of laminar flow because 

 of the extent of the favorable pressure gradient. Both wind tunnel and 

 flight data give estimated equivalent flat plate Reynolds numbers cover- 

 ing a range from about 600,000 to about 14,000,000, the high values being 

 obtained much more recently on laminar flow sections in wind tunnels of 

 turbulence less than 0.1 per cent and in flight on models in which extreme 

 care had been taken to remove surface roughness and waviness. In any 

 particular measurement it is almost impossible to separate the effects of 

 the many variables. 



Fig. A, 17a illustrates wind tunnel data on the chord wise position of 

 transition on airfoils at approximately zero angle of attack [43,48,57,58, 



0.8 r? 



Xt/c 0.4 



Re X 1 0"^ 



Fig. A, 17a. Transition position on airfoils as a function of airfoil Reynolds number. 



69,60,61]. In every case transition moves forward with increasing Reyn- 

 olds number but the rate and the value of the Reynolds number at which 

 the most rapid change occurs are dependent on the type of airfoil section, 

 the smoothness and fairness of the surface of the airfoil, and on the turbu- 

 lence of the wind tunnel in which the measurements are made. 



The results for the 65215-114 airfoil [48], showing transition as far back 

 as 25 to 30 per cent of the chord at Reynolds number of 40 to 55 million, 

 were obtained in the low turbulence wind tunnel of the Langley Aero- 

 nautical Laboratory in which the turbulence intensity is a few hundredths 

 of 1 per cent. The results for the Tani-Mituisi airfoil at the extreme left 

 of Fig. A, 17a were obtained in the FFA wind tunnel at Stockholm [43] 

 behind a turbulence grid giving a turbulence level of approximately 1 per 

 cent. Although there is some influence of airfoil shape and surface wavi- 

 ness in this comparison, the principal differences between these two curves 

 are believed to be due to effects of wind tunnel turbulence. 



(42 > 



