section lift coefficient decreased
airfoil becoming supercritical. The maximum section lift coefficient decreased about 30 percent for an increase in Mach number from 0.10 to 0.36.
5. Section lift and pitching-moment coefficients at low Reynolds numbers for the smooth airfoil were in good agreement with results from another low-speed wind tunnel.
However, there were differences in the drag coefficients within the lift coefficient range where the laminar bucket would be expected.
6. Comparisons of experimental section lift coefficients, pitching-moment coeffi- cients, and chordwise pressure distributions with those calculated from a viscous-flow theoretical method were good as long as no boundary-layer flow separation was present; however, the theoretically calculated drag coefficients were less than the experimental drag coefficients.
Langley Research Center, National Aeronautics and Space Administration, Hampton, Va., December 7, 1974.
APPENDIX
APPENDIX CALIBRATION OF THE LANGLEY LOW-TURBULENCE PRESSURE TUNNEL A calibration of the Langley low-turbulence pressure tunnel has been performed using a long survey probe alined with the longitudinal center line of the test section. The nose of the probe contained a total-pressure tube and static-pressure orifices were installed flush with the probe surface at fixed interval distances over the probe length.
probe total-pressure and probe static-pressure distri- These were used to measure the of the airstream. A photograph of the probe is shown in figure 16 and a sketch of bution the calibration arrangement, relative to the wind tunnel at various longitudinal stations x cm (in.), is shown in figure 17. Also shown in figure 17 are the reference total- pressure tubes on the floor of the wind tunnel and the tunnel sidewall reference static- pressure orifices from which the tunnel test conditions are calculated.
Some typical Mach number distributions at various tunnel total pressures (that is, Mach number distributions show that Reynolds numbers) are shown in figure 18. The the flow is uniform with negligible deviations between x = -101.6 cm (-40 in.) and x = 101.6 cm (40 in.). Figure 19(a) shows the results of the calibration in terms of a calibration factor (C.F.) against average probe Mach number, and figure 19(b) shows the operational boundaries of the tunnel. The calibration factor is shown to vary by present less than 1 percent over the entire operational Mach number and Reynolds number range of the tunnel, and no specific trends of variation of the calibration factor with either The dashed line in figure 19(a) Mach number or Reynolds number could be identified.
indicates the root-mean-square value (rms value) of all the data (C.F. = 1.006), which therefore is defined as the calibration factor for the tunnel. This calibration factor is in good agreement with a previously used but unpublished value (1.005). The symbols used in the calibration of the tunnel (see figs. 16 to 19) are as follows: Pt,p- P C.F. calibration factor, , averaged from x = -30.48 cm (-12 in.)
Pt,ref ref to x = 60.96 cm (24 in.)
Mach number M free-stream Mav longitudinal average of probe Mach number between x = -30.48 cm (-12 in.)
and x = 60.96 cm (24 in.)
Mp probe Mach number (computed from pp and Pt,p)
APPENDIX - Concluded
APPENDIX - Concluded 2 2 kN/m (lb/in ) p probe static pressure, ) kN/m (lb/in sidewall static pressur'e, tunnel Pref 2 2 Ptp probe total pressure, kN/m (lb/in ) (lb/in pressure, kN/m Pt,ref tunnel total 2 2 ) q dynamic pressure, kN/m (lb/in R unit Reynolds number per m (per ft), based on stagnation temperature 300 K (5400 R) x longitudinal station, cm (in.); x = 0 coincides with axis through center of rotation of sidewall circular plates RE FERENCES of Air- 1. Abbott, Ira H.; Von Doenhoff, Albert E.; and Stivers, Louis S., Jr.: Summary foil Data. NACA Rep. 824, 1945. (Supersedes NACA WR L-560.)
Frank T., Jr.: The Langley Two-Dimensional 2. Von Doenhoff, Albert E.; and Abbott, Low-Turbulence Pressure Tunnel. NACA TN 1283, 1947.
3. Braslow, Albert L.; and Knox, Eugene C.: Simplified Method for Determination of Particles for Boundary-Layer Transition Critical Height of Distributed Roughness at Mach Numbers From 0 to 5. NACA TN 4363, 1958.
4. Baals, Donald D.; and Mourhess, Mary J.: Numerical Evaluation of the Wake-Survey the Effect of Energy Addition. NACA Equations for Subsonic Flow Including WR L-5, 1945. (Formerly NACA ARR L5H27.)
5. Allen, H. Julian; and Vincenti, Walter G.: Wall Interference in a Two-Dimensional- NACA Flow Wind Tunnel, With Consideration of the Effect of Compressibility.
NACA WR A-63.)
Rep. 782, 1944. (Supersedes of Compressibility on the Maximum Lift Coefficient of 6. Wootton, L. R.: The Effect Aerofoils at Subsonic Airspeeds. J. Roy.
Aeronaut. Soc., vol. 71, July 1967, pp. 476-486.
7. Englar, R. J.; and Ottensoser, J.: Calibration of Some Subsonic Wind Tunnel Inserts for Ship Res. & Develop.
Two Dimensional Airfoil Experiments. TN-AL-275, Naval Center, Sept. 1972. (Available from DDC as AD 913 412L.)
8. Stevens, W. A.; Goradia, S. H.; and Braden, J. A.: Mathematical Model for Two- Dimensional Multi-Component Airfoils in Viscous Flow. NASA CR-1843, 1971.
9. McGhee, Robert J.; and Beasley, William D.: Low-Speed Aerodynamic Characteristics of a 17-Percent-Thick Airfoil Section Designed for General Aviation Applications.
TN D-7428, 1973.
NASA 0.50, MEASURED AIRFOIL COORDINATES TABLE I.- NACA 651-213, a = c = 60.63 cm (23.87 in.)
Lower surface Upper surface x/c z/c x/c z/c 0.00000 0.00000 .00014 .00297 0.00021 -0.00096 -. 00215 .00042 .00431 .00054 .00080 .00551 .00098 -. 00327 .00127 .00669 .00152 -. 00431 .00132 .00775 .00214 -. 00533 .00285 -. 00626 .00244 .00880 .00362 -. 00717 .00313 .00981 .00444 -. 00797 .00386 .01074 .00530 -. 00871 .00624 .01311 .00619 -. 00943 .00884 -. 01114 .01111 .01656 .01403 -. 01363 .02348 .02298 .02680 -. 01788 .04840 .03281 .05214 -. 02404 .07345 .04052 .07737 -. 02868 .09858 .04703 .10252 -. 03256 .14893 .05750 .15271 -. 03858 .19939 .06550 .20279 -. 04308 .24993 .07153 .25280 -. 04640 .30052 .07604 .30275 -. 04872 .35115 .07897 .35267 -. 05016 .40182 .08031 .40254 -. 05065 .45254 .07990 .45238 -. 04999 .50339 .07749 .50207 -. 04811 .55416 .07291 .55185 -. 04504 .60462 .06665 .60193 -. 04106 .65493 .05918 .65216 -. 03625 .70513 .05066 .70251 -. 03086 .75522 .04151 .75297 -. 02504 .80526 .03201 .80347 -. 01899 .85525 .02250 .85403 -. 01298 .90527 .01329 .90456 -. 00729 .95532 .00520 .95506 -. 00250 .99999 .00015 1.0 -. 00013 II.- AIRFOIL ORIFICE LOCATIONS TABLE c = 60.63 cm (23.87 in.)] Lower surface Upper surface z/c x/c x/c z/c 0.0 0.0 0.00515 -0.00862 .00933 .00281 .00715 .01388 .01010 -. 01183 .01619 .01031 .01273 .01753 .01759 .02017 -. 01588 .02144 .02008 .02012 .02378 .02520 .03009 .02589 .05045 -. 02372 .03344 .05024 .07538 .04107 -. 03235 .04750 .10087 .10044 .15130 -. 03843 .05786 .15078 04296 .20129 -.
.20105 .06574 -. 04636 .07180 .25162 .25178 -. 04870 .07615 .30189 .30180 -. 05016 .35204 .35211 .07902 .37729 .07987 -. 05065 .40231 .08032 .40264 .42775 .08033 .45260 -. 04999 .07990 .45293 .07898 .47781 -. 04806 .07752 .50304 .50319 .52811 .07551 .07432 .54077 -. 04497 .07302 .55337 .55334 .07159 .56607 .07009 .57843 .06848 .59116 .60374 -. 04090 .60342 .06685 .06324 .62863 03609 .65377 -.
.65391 .05934 -. 03074 .05087 .70375 .70407 .75459 -. 02485 .75419 .04173 .80490 -. 01879 .80482 .03210 -. 01286 .02252 .85513 .85509 .90504 -. 00722 .90484 .01336 .00519 .95543 -. 00248 .95530 TABLE III.- NOMINAL ROUGHNESS PARTICLE HEIGHTS [Grit located at x/c = 0.05] R Grit number Nominal particle height, cm (in.)
3.0 x 10 0.021 (0.0083) 5.9 .012 ( .0049) 9.0 .0089 ( .0035) 11.8 .0074 ( .0029) 14.5 .0074 ( .0029) 17.4 .0074 ( .0029) 22.8 .0074 ( .0029)
Moment reference
z/c 0
- I
.9 1.0
.7 .8
.5 .6
.3 .4
.1 .2
S0
x/c
airfoil.
651-213, a = 0.50, for NACA I.- Section shape Figure CD.
Tunnel side walls Diam.= 1.68 c-- - -- .. . .
--- _ 1.51c Airflow A A Circular plate Metal seals - Airfoil positioning attachment Top view Model attachment plate Seal detail "z" -I _Zero incidence reference Tunnel center line c/4 c - End view ,section A-A Figure 2.- Airfoil mounted in wind tunnel.
All dimensions are in terms of airfoil chord c = 60.63 cm (23.87 in.).
o.126c .042c .021c Static-pressure probes .042c-- (typ.)
-1 1 L.17c Airflow .0052 c (typ.)
Total-pressure probes (tubes flattened) ,189c 0 yp.
Figure 3.- Drawing of wake rake. All dimensions are in terms of airfoil chord. c = 60.63 cm (23.87 in.).
a = 150; R=3.0 x 10 a =160; R =6.0x 10 C,max Stall cm (a) Tufts.
L-74-8543 Figure 4.- Flow-visualization photographs for NACA 651-213 airfoil.
M = 0.22; model smooth.
0 6 R= 18.0 x 10 a = 160; ; R= 12.0 x 10 a=17 Stall Stall O L-74-8544 Concluded.
(a) Tufts.
Figure 4.- Continued.
C- = 50; R= 3.0 x 106. L-74-8545 (b) Oil flow.
Figure 4.- Concluded.
2.0 - ___ Roughness o Off 1.
1.2 A- A :J- .8
- -
C 4 iiiii !i
-. 4 -2 ::-I::::I: _::25 -12 -8 -4 4 8 a,deg (a) R 3.0 x 10 M = 0.22.
Figure 5.- Effect of Reynolds number on section characteristics.
o ' Off
[ On
.020
.012
Cd
.008
.004
-I.2 -. 8 -. 4 0
.4 .8 1.2 1.6 2.0
(a) R = 3.0 x 10 . Concluded.
Figure 5.- Continued.
- - -i - -. - - 2.0
I
Roughness
Off
: i H\
1.2 -:_: rn 0- -. 4 -. :i~iri i.....iiii. ... .i .: ......
-1.2 ; -i -1 -. 2 16 20 4 8 12 -12 -8 -4 0 adeg (b) R = 5.9 x 10 Figure 5.- Continued.
V"4 4;t , Houghness
0 Off
o On
.020
.016
.012
c
d
.008
.004
-1.2 -.
8 -. 4
0 .4
.8
1.2
c l (b) R = 5.9 x 10 . Concluded.
Figure 5.- Continued.
Roughness o Off 1.6
i
-
0 On SL: .8 0- -. 8 i I i
-1 .2
-
-12 -8 -4 0 4 8 12 16 20 a,deg x 10 (c) R 9.0 Figure 5.- Continued.
[ On
.020
.0 16
.012
cd
.008
.004
01.2 -. 8
-. 4 O .4 .8
1.2 1.6 2.0
c
Cl (c) R = 9.0 x 10 . Concluded.
Figure 5.- Continued.
2.0 .
___ lllii .Roughness : __ - - - Off 1: 711 1 : :: ff o O :.6 : :!'i: ii ili i a On .2 . .... ...
. .. ... . . : ': ....... ..
c 4 F4 .2 l ; - I: , i i i:i :: . ::: 1 ij: jhi: -- -- l-- : I: : :I : i : .
- 1 i ! iii : iii; 2 2 Figure 5.- Continued.
E F': . . .. . - ' ' E i l '
0t I .I
I i!; i I'Eii: :i I i h F *i iEi:'uh iH : : ' iF ' -4 0 4 8 12 16 20 -12 -8
o Off
o On
.024
.020
.016
cd .012
.008
.004
1.2 -. 8 -,4 0 .4 .8 1.2 1.6 2,0
(d) R = 11.8 x 10 . Concluded.
5.- Continued.
Figure ii i Roughness o Off 1.6 o On 1.2
- .4 -,;:ii
,:: .'-! 7777 ,;i -. 4 ............
-1.2 1:: 1 .i!ii 'I h i H:!:1 l: lI* l: i iii1 iiii -12 -8 -4 0 4 8 12 16 20 a,deg 14.5 x 10 (e) Rr Figure 5.- Continued.
St
ougness
o
Off
.024
.020
.016
Cd .012
.008
,004
0,2 -. 8 -,4 O .4 .8
1.2 1.6 2.0
C1 (e) R = 14.5 x 10 . Concluded.
Continued.
Figure 5.- 2.0 Roughness
1.6 2
.2 ii iii
.8 .. . ...
jjj ilfi jt, .; ,I ii , li I lit -1.2 12 16 4 8 -4 0 - 12 -8 (f) R=17.4x10 (M) R : 17.4 x 106.
Figure 5.- Continued.
o Off
.024 [ On
.020
.0[6
Cd
.012
.008
.004
-1.2 -. 8 -. 4 0 .4 .8
1.2 1.6 2.0
CI
(f) R = 17.4 x 106. Concluded.
Figure 5.- Continued.
it 1 , 2.0 - siii!i l ii;iii!ii~ iiii!
T i,.! ii ' ' iii i ih !!' " iiiii-i , i " i R u h e - '2.0 t tii i l il Ro-ughn ess i !!!!!
-7 - L 0 Off ' ' iii ' t !liii!i; - v :iiii-, ii i i 1 .6 t' o Ofn S t it t i ! :i: '. ' t-e : : 5 55 ; ii!,: i : i'!! ! .. . . . . . . .
it i t ~ S tS '" .... ii- i i ..
HH...
ii iii r !iii ii 01
' c[i~i 4 idi F . i' .i q!! ::bl S it r~S jiil jj 1 ;it 1 1ii~ :;:: ,,[! , 111 1 1:ii ii, ill, 4 ~ ilj ijj ii!l~iiii' ": ir;i;~~tr::;
~iiii "H iiii! i I ii:!i il
1 F lit i uii nii Il [ ill If it H i i lit : •.iliiii I mii::;i::::r: I i lir: i! , ;tHI~r ~ it i ll, 'l: iiiiiii;i; O ~ ~
II tii ; litI~
Ktk if ij j~{~ {ti '. h i js s i N "js 14 ' . ir *; t It (jtRi n~j i
jjt t ;I; ijHil Irt iti ILI it ii lit It I!i ql'il~~'i i 11P I, ,fp 0i ii ', lkiiW.4.
iiL i Ipihi l V ii 1 714ii:il l iiilil f 1iiih tilh Vl i
, I f
' it c ii m (g) R =22.8tx10.
Figure 5.- Continued.
.- Cotnud Fiur s~ ii H37 a ,deg 3'7
.024
iWM M
On
.020
.016
Cd
.012
.008
.004
1.2 -. 8
-4 .4
.8 1.2
.6 2.0
cl (g) R = 22.8 x 106. Concluded.
Figure 5.- Concluded.
-20
-16
Cp,critical
min
Cp -
-8
-4
0 4
8 12 16 20 24xl06
- - - - -- - - - - of airfoil minimum upper surface pressure coefficient'with Reynolds Figure 6.- Variation numbers. M = 0.22; transition fixed at x/c = 0.05.
2.0 : F Roughness tLtF l i.i F i::ii :i! : $!: : ! : i i , . . .i .. L 1.6 Iu6 No.60 O No.120t A No.180 1.2 .8 . T 7 . .. !:... i i' : i:" i:j i:i ;:i!:i:: i~i'iii7iil .. . .... ~ili i ..... ii : : .. . ...... . ii i:ii ii iiiii: .... ....
- - - Fit .2" : ;, : '.: F4' E , 0 4 2 F 16.. 2 F (a) Lf adondt -12 -8=4 0; 52 4 8 16 2
----------
O ff
No.60 No. 120 No.180 .020
.020
.0 16
.012
cd
.008
.004
001.2 -. 8
,4
O .4
.8
1.2 1.6
2.0
Cl (b) Drag polars.
Concluded.
Figure 7.- 0.04 - - - 0 Wraparound roughness to at 0,04c ii Strip roughness 1.6 1.2 .8 C1 .4 : i 4 i i tur ii 11 i 1 1 :1~::-I: :
0 iiiii li iiii~ii
j:iirii '7 1ll :l:;::: il ;iiiillr ii;; iijj/ i;;iii~i~ii ll~t I: iiii~ilii Tliti :: ::i;i: !~LLL1 -. 4 8 12 16 2 -8 4 0 4 -12 aideg "':'i'' " (a) Lift and moment data.
chracerisics .i !:i- thestrp an wrparond oughesson sctin Figure B.i,: E -.2 M = .15;R = .9 106; no 60 rit 1:i11 ~ ~ ::I: i 42~iiiiitiiii''' 16 20 0 4 8 12 -12 -8 -4 adeg (a) Lift and moment data.
wraparound roughness on section characteristics.
Figure 8.- Effect of the strip and M = 0.15; R = 5.9 ×106; no.
60 grit.
.028
o Wraparound roughness to 0.04c
- Strip roughness at 0.04C
.024
.020
.016
cd
.012
.oos-
.008
.004
-,4 0 .4 .8 1.2 1.6 2.0
-1.2 -. 8
Cl
(b) Drag polars.
Figure 8.- Concluded.
i.8
Model smooth 1:1Up "I'll
o
"i ....
roughness
standard
[] NASA
,,~li u ,~~~jYW-;~~~t~
,. -.- iE, .. ...
I 6
7 1T 1 1 II .I i H-i I 1.4I -i i I ii .. . . .
, I 1 4 ... max .
' . . .
-•i - " "
1 .2 il j__ : '1 !- f 4
T I
IT P iati o it Fr
"R
1 LI Jill
30 0
,0
0,
3 4
t 2
R
number. M = 0.22.
with Reynolds maximum lift coefficient 9.- Variation of Figure
.014
-- -- os Model smooth
- ,,, 0 NASA standard roughness
+ F 7 , TjTi
.012 -i
o~: _L _ i , ii "j' -- " i ! In ....
I l
mini
cd
1-T
-- -- 4 IT
006 ' " '" '
-T TT 1 : I 7 Ii T1'
"001 2l ff 0 03X
-T j; Jill LIuL .006 ii, .008 r 10. Vaiaio 4- minmu rAgL- IT- fcin lith Renod'nmbr I =4 -1 7T,
.04 112
2 0xO
,, i ii'R
ariaionof dag oeficint wth eynldsnumbr. = .22 Figue 1.- inium ..
. 3 1 ' 4f0 . ..... .. ...
' : H it. ... : !
0ii '
• ,::i0
2i .2i M it i:i iL 1i If!
iil iii i!'ii. ... ::-
44 ';i i ii.36 .15-
1.2 .28 ) . 4 .8 ~ ~ 4 N I 4. , .... .. ... .
- I.24444 i , .......... it *t))*.+4
iiii 1 .... lit It iil~ijli !ii ililiii jIII ti! t;!! 1 Ifiili~iiii ""'"
4 4
o8
~ .4.4i~t,,i,1ii iir ii'li"l .: 11 E *4** 4 Ii~ ............
- i; iiti- iljjf tt ri :':i!Iiiiii liii jjii j ,: '{-! , Ii~ii ,F:3 : lit!
1V il . ii 1ii U -i H I f!'
lfil l 11 i 1-,I ll M 8 r;:;, 71 t I li ~44~ , 4. 44 11, 41144+1414 {( )I 4t4).)) H~ ~ ~~i , - !, ::,: ::: iiiiHi :: (:: i (H ,! 1.2 I'il: q'i il ii ti itii 'JiT 11Iff ff I'll i i iili ii i N it ki i i l t Iti id 1 111 "' 1 1 l 1 ]1 il li i l i i " !i!! ii!((t' ...... , ,, (J ((!( i' ...
.... ,iii, fl i iI ; i Oiltim it ijjjjiti 0I 11:ii ; i ! ; il 1 ::i::I H , Hit!; Oh :fi ... . ), . .. . .. f) ) If I~ 1 i 4 ! TI 1 11 il T I 11 i -i!1 iil; M i lit [ fiii '"~i l 11 it 11114 N lit 1[4 .2
-8 -4 4 8 12 16 iiii '!ii il:ii~i '
20ji
-. 1 + -, ,j ........ I'tiVtii t 1414f ji'[j 1t ~ir
.I i I i~ t I iiiir~i 'I;I, ii !;1a : de g ;i1i ii/i~ji (a) Lift and moment data.I~i::!ii umbr o setio chraceritic Figre 1.-Effct f Mch R= 59 X10 istibuion.
andchodwie pessre transtionfixe at / =0.05 -2 -8 4 0 4 8 12 16 20 cz,deg (a) Lift and moment data.
Figure 11.- Effect of Mach number on section characteristics and chordwise pressure distributions.
R = 5.9 X 106; transition fixed at x/c 0.05.
o 0.10
.024
. .15
<> .22
.020 A .28
. 36
-
.0 16
Cd
.012
.008
.004
- 1.2
-. 8 -. 4
.4 .8
1.2 1.6
2.0
cl
(b) Drag polars.
Figure 11.- Continued.
47,
-7.2 L
-10 -6.8 - M C. critical 0 0.10 - 66.9 - 13.4 O .22 -9 .28 - 8.1 1- -6.4 - 4.7 A .36(stalll - Upper surface - -6.0 - Lower surface - - - -7 - - -5.6 Cp -6 -5.2 - - - - - - --- - - -5 -4 .8 - i -4 -4.4 -4.0 - - - - - - - - -\ -3.
0 .04 .08 .12 x/c -3.2 Cp -2.8 - - - - - - - - - -2.4 - -2.0 -1.6 -1.2 .7 .8 .9 1.0 1 02 .2 . 4 .5 .6 x/c a = 120 Chordwise pressure distributions; (c) 11.- Concluded.
Figure
-20
Cp, mn
Cp
- 12
C, maxl.4
-8
1.2
1.0 -4
.3
.1 .2
.4 O
.2 .3
o .I
M
M
upper surface pressure and airfoil minimum of maximum lift coefficient Figure 12.- Variation transition fixed at x/c = 0.05.
Mach number. R = 5.9 x 106; coefficient with - 1.6 Airfoil a R 2 13 6 5 0.50 3.OxlO 0 NACA 1 1.2- - 1.2 / - NACA 651-212 .60 3.0 (ref. I) S- - - NACA 65,-213 .50 2.3 (ref.7) .8 .4 -. 4 .024 - - .02 -- - -8 -
0-- ----- -- O ---- t-- -3-
-12 _ - .016- ---------- --- -------- C d - \ cd (m 0 a-- .0082-- -a-.0 -- \ - .004- -- - - - - -. 1 O .2 .8 1.2 1.6 20 -1.2 -. 8 -. 4 0 .4 -8 -4 0 4 8 12 16 -12 a,deg x 10 2.3 x 10 and 3.0 (a) R= 6 5 2 1 2 1- (a 0.60) and 651-213 (a = 0.50) Figure 13.- Comparison of section characteristics for NACA < Models smooth; M 0.22.
airfoils.
Airfoil a NACA 651-213 0.50 1.6 NACA 651-212 .60(ref.1) 1.2 c .4
.8-/d -
-1.2 - .016 0 - --- .012 -,.0 -. 2 -00 -. 8 -,4 0 .4 .8 1.2 1.6 2.0 -12 -8 -4 0 4 8 12 16 20 -1.2 a,deg CI 6.0 x 10 (b) R Figure 13.- Concluded.
o Upper surface o Lower surface -2.4 Theory -2.0 - -I.6 Cp -.8 -. 4 .4 .8 1.2 .7 .8 .9 1.0 0 .1 .2 .3 .4 .5 .6 x/c =-4.1 .
(a) a pressure chordwise and theoretical of experimental Figure 14.- Comparison x 106; model smooth.
M = 0.22; R = 5.9 distributions.
-2.4 surface Upper o e Lower surface -2.0 - - Theory -1.6 -1,2 -.
cp
-. 4 I
-z
.4 .8 1.2 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/c (b) a = 0.00.
Figure 14.- Continued.
-2.4 surface o Upper Lower surface e -20
- -- Theory
-1.6 -1.2 -. 8 ,- -
Cp
-- S -.
.4 .8 1. 2 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 X/C (c) a = 0.600.
Figure 14.- Continued.
-2.4- 0 Upper surface e -2.0 - Lower surface - - - Theory
-1.6
-1.2 Cp -.
.4-- .8 - 1.2 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 X/C (d) a = 2.10.
Continued.
Figure 14.- -2.4 O Upper surface e Lower surface -2.0-- - - -Theory -1.6 -1,2 ) -. 8 _ .- -- Cp .4 ( 1.2 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 X/C a = 4.10 (e) Continued.
Figure 14.- -4. - 0 Upper surface e Lower surface 4.4 -- - - Theory -4.0 -3.6- -3.2 -2.4 .
-2.0 Cp -1.6 -,2- I - 0-.8
-. 4 --
.4 1.2 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 X/c (f) =8.3.
Figure 14.- Continued.
-4.8 o o Upper surface e Lower surface -4.4 - Theory -4.0 -0 .- 3.6 -3. - 9 -2.8 + -8 -2.4 Cp -7 \ k -6 -6 -2.0-- C- -. 6 -5--0 .04 .08 .12 X/C -1,2 1 2 ' I0 -. 4 'O(b .4 - -58 - - - - - - - - - - - .2 ----- 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/c Figure 14.- Continued. Figure 14.- Continued.
-4.8 0 I o Upper surface e Lower surface -4.4 - -12 -4.0 i -II -3.6 -10 -3.2 Cp _ 9 -2.8 -8 -2.4 -2.0 - 6 C Cp -1.6 - -5 0 .04 .08 .12 X/C -1.2 D0 / .4 .4 .8 - -- 1.2 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 X/C = i5.1 (h) a Figure 14.- Continued.
-1 -4.8 Upper surface O e Lower surface -13 -4.4 -3,6 ----- 10 i -3.2 0C -2.8 --- 9 -2.0 -7 Cp -1.6 -6 0 4 0 8 -1.2 0 5 x/c . .12 O o - -.
-. 4_-- .- ,- 0 - .4 .8 - 1. 2 - 1.0 .1 .2 .3 .4 .5 .6 .7 .8 .9 X/c (i) a =16.20.
Figure 14.- Concluded.
2.4 o Experiment Theory - 2.0 1.2 .8 cl .4 -1.2 -.
I -. 12 -8 -4 0 4 8 12 16 20 2 a,deg (a) R = 3.0 x 10 ; model smooth.
Figure 15.- Comparison of experimental and theoretical section characteristics. M = 0.22.
.024
Experiment o Theory .020
.016
cd
.008
.004 2.0 .8 .2 1.6 0 .4 -1.2 -. 8 -. 4 Cl x 10 . Concluded.
(a) R = 3.0 Figure 15.- Continued.
2.4 - Experiment o - - - Theory 2.0 1.6 1,2 ,4 -. 4 -. 8 1.2 Cm -22 -8 -4 0 4 8 12 16 20 a ,deg (b) R = 5.9 x 106; model smooth.
Figure 15.- Continued.
.028
o
Experiment
Theory
.024
.020
.016
c
d
.012
.008
.004
1.2 .6 2.0
2 -. 8 -,4 0 .4 .8
Cl
x 10 . Concluded.
(b) R = 5.9 Figure 15.- Continued.
MOM o Experiment Swep TM MUM MUM 8 Mt MMOMW 1 -4 SM Theory 1.11MMU MUN a8H i El iff 1 M i 4I i 1.2 111 ii:iii iiiiiliil i i ill tll U l 1, itit iiii i iiliiij :
:;j j iiliil!itlj: fllii:Ii: l lit!lll
HI I 1 81 i!&i it Haii~ l 000 HM~ H i Ii ii: ; ... Im m U !R 0 t:til I t .. ..
i Uilil ?llli!!lii !:: ' ;I:: ::
:: !iiliilif lit Ili Ii i
~i ::i ...liilit ... i 1; 1 llj/lij iiliiii~ m Wli 11:1 1;1: H1 O I il til i i l l ill 1 ; i ; 1- ~~iT~F -. 0 t lit, 1: i i:: l:: ii t l it it , E!E EU il000 11 ij;ER lili~l::lit iiii mi liq i l I: 0 il::1 lit Ill i.: it ;' l i .E a li W i u m i ;l m 'TElTHi ill i::! iU:I' l.it tillEEM .i it IE lit il 00 imE it 1H"E iH i ill l : 11 ll!jl m :11 . :: i inj iiii I;; ::: i !; ti):lll ii il ::: l:: lil it 11" Ii '- till I: it ill, p' i i1 til I 'I I ; 1! l T ;1it T WI h i !!:I i'i~f~i :':'"' !' ': lit, WIii:!
.4 10o moe smooth.
till R; =t 118 xl Cinued., i 15.- m~li IlHitm 1MRHE 1 M'M !i:: l ii till 1 ii iii li i t iii il :;!w il t lI l~ !! .ll i 1ii , 1 I d i i ii iiH I il !H H 11111 W ; it lll 0 i li l i 1 11 : lii11 I ll! ii:: litiiiii i: q ii llilitIMil: l !I i I Il iii !t/ .a.deg I I"~'i; ill c M !ill : I! iil ' i H it ti i !!J -12 Ht:0 2 S 2 H I ii~ i ... iii ill iii! Iiii iii ili il iiii til I ill Hill Ii ;iiiii 11 !I'll Ii t iiii jlji ll Jj i lt M l!)1 m i JH till T "" tt" acde i (c) R 11.8 x 10 6 ; model smooth.iii~i Figre 5. Cotined ::1 'iiiiii6 5
.028
0 Experiment
- Theory
.024
.020
.0 1 6
Cd
.012
.008
.004
2.0
-. 4 0 .4 .8 1.2 1.6
-1.2 -. 8
cl (c) R = 11.8 x 10 . Concluded.
Continued.
Figure 15.- 2.4 4i ,i o Experiment :ll1-lT1flI. . .. ' Theory 2.0t I IU H", illIiiii I lilt 1.6 lit- it I 'I lit 1UP , liii~i iii ifll 1 Il tIili iklit iii W ij li llfi il ilithii tii V! 1 111ql 1 1 ll !itfiiti 1 13 11 Ili illllt i~llttt iitil~ll - : lii 111 !1 1 , it:,. i li~i lt Illiff iliiiii ill N IA lilt lliiiiIIIIiII 1 lit lilt iq li i iiilitl ill liij li t 4l 1.2 ii'"; iiili'iMl HIWINNIIIH Ill'' liti~ it .8 i i ; .8 il tll i itili! i it T Al!il l il i N i iiflliflili 1; 1 WINN ONONlii lf it 1111 i~TjfTiiiiiI 11 .411t, 1:::: it i 111 11l II 111 ild it Ii j it Ill il 61 it I i l tl iiII 11 4til 1i i I w w l-l 1 0 T -.4 -. 8 -I.2 . I li C m -12 -8 I I0 4 8 12 16 20 a,deg (d) 1R = 22.8 × 10; model smooth.
Figure 15.- Continued.
.028
Experiment
o
Theory
.024
.020
.OI 6
Cd
.012
.004
2.0
.8 1.2 1.6
-. 4 0 .4
-1.2 -. 8
C1
22.8 x 106. Concluded.
(d) R = Continued.
Figure 15.- 2.4 o Experiment Theory 2.0 1.6 1.21 c I .4 -. 4 -. 8 -1.2 .1 -. 2 - -4 0 4 8 12 16 20 -. 412 Jil M adeg U110 mll l 106; transition fxed at x/c = 0.05.
(e) RF = 5.9 x Figure 15.- Continued.
.028 ii
Experiment
o
0 - -- TheoryI
.020
.01 6'
.0 2t " iiiiiiii:
.0 1 . .... 8 -. ... 6
(e) = 5.9 x 16 IConcluded
0 . 8 I 16 i20
-. 2 -. 8 - 4
Figure 15.- Concluded.
Figure 15.- Concluded.
L-72-3634 in wind tunnel.
probe mounted Figure 16.- Calibration
PA
ORIGINAL
QOBcuMT9~ts
iOoR
Center of rotation plates circular I-sidewall Pref Pt,p Typical pp Airflow Tunnel \ Probe Pt,ref -140 -120 -100 -80 -60 -40 -20 0 20 40 60 x,in.
I I i I I i I I I I I -350 -300 -250 -200 -150 -100 -50 0 50 100 150 x ,cm Figure 17.- General calibration arrangement.
.24
.22
Np
M I Pt,ref,obsolute Mo v .20 lb/in; kN/ma 4f W 0 14.7 101.3 0.222 t [ 60,0 413.7 .2 0 120.0 827.4 .22 Mp Center of rotation .24 .20 of circular plates .22 NiW .20 .18-72 -64 -56 -48 -40 -2 - 4 -16 -8 8 2 16 24 32 40 x in.
I I I I I I I I I I I I I I I -180 -160 -140 -120 -100 -80 -60 -40 -20 0 20 40 60 80 100 x ,cm center line of tunnel.
probe Mach number distribution along longitudinal Figure 18.- Typical -4 C, ,absolute Pt,ref Ib/ine kN/m 14.7 101.3 0 30.0 206.8 1.02 0 45.0 310.3 60.0 413.7 rms values 6 91.0 627.4 1.01 11 120.0 827.4 C.F.
M 1.00 .04 .08 .12 .16 .20 .24 .28 .32 .36 .40 .44 Mayo (a) Calibration factor (C.F.)
as a function of May.
riiT33(10) Speed control tank pressure, q, kN/m2 q,lb/ft 400 150 Ib/i 0 I200 1 M I I obs olute 14.7 Ib/in 101.3 kN/m' (b) Variation of q with M for constant values of Reynolds number.
Figure 19.- Calibration factor and operational characteristics of Langley low-turbulence pressure tunnel.
NASA-Langley, 1975 L-9773