mkL-n ^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED April I9U5 as Advance Restricted Eeport L5C13a INTOSTIGATIOH OF EFBTCT OF SUESLIP ON LATERAL STABILITr CHARACTERISTICS in - RECTAHGULAR LOU WHKJ ON CIRCDLAR FUSELAOE WITH VARIATIONS IN VERTICAL-TAIL AREA AND FDh'KT.AflT: LENGTH WITH AND WITHOUT HORIZONTAL TAIL SURFACE Bj Thomas A. HoUingvortli Langley Memorial Aeronautical Laboratory Langley Field, Va. ►rxex Sr^ WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution. L-17 DOCUMENTS DEPARTMENT Digitized by tine Internet Arclnive in 2011 witli funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/investigationofe inf^nn 3 and 5^ over an angle-of-attack range from -10''-' to 20° and at an angle of attac;c of 10.2° over an angle-of-yaw range from -50° to 12°. All tests were run at a dynamic pressure of 65 pounds per square foot, which corresoonds to a test RejTiolds number of approximately 388,000 based on an 3-inch v;ing chord. The data may have been affected by compressibility at large angles of attack. PRFJSENTATION OP DATA The results of the tests are presented in standard NACA coefficients of forces and moments. Rolling-moment and yawing -m.oment coefficients are given about the center- of -gravity location shown in figure 1. The data are referred to a system of axes in vi'hich the Z-axis is in the plane of S3rmmetry and perpendicular to the relative wind, the X-axis is in the plane of symmetry and perpen- dicular to the Z-axis. and the Y-axis is perpendicular to the plane of symmetry. The coefficients and symbols used are defined as follows ; Cl lift coefficient (l/iSv/) Cd drag coefficient (p/qS^^^ Cy lateral-force coefficient ^^Y/qSw) Cy, slope of curve of lateral-force coefficient against angle of yaw (^'Cy/0\[/) Cj rolling -moment coefficient (L'/qbSw) NACA ARR No. L5C15a Cj, slope of curve of rolling-moment coefficient '■^ against angle of yaw fdCi/d-if^ Cn yawing -moment coefficient (K/qbSvA )pe of curve of yawing -moment ci against angle of yaw fdCn/^^') Cn, slope of curve of yawing -moment coefficient A]_ increment of Cn,i; or Cy-.w caused by wing- fuselage interference A2 increment of Cn-i; -^^ Cy Cj., and Cv was determined experimentally to be about ±0.0005, ±0.0003, and ±0.001, respectively., The average experimental accuracy of Cnj;^ Gj^, and Cy^ is about ±0.00010, +O.OOOI6, and ±0.0002, respectively. The accuracies of the angle-of -attack and angle-of-yaw measurements are about 0.1^ and O.O5", respectively. Angle of attack and drag coefficient were corrected for tunnel-v;all effect by the following formulas; Aa =^ 57.5aw~-CL = O.609CL (deg) aCd = ^v^'cl^ = 0.0 106c r3 v/here 5w jet-boundary correction fc^ctor at wing (O.I525) C cross-sectional area of tunnel (Jo sq ft) NACA ARR No. L^ClJa 7 Both corrections are additive. No jet-boundary correc- tions v.'ere apolicd to Cj. Cn, and Cy- The correction to Cy is within the experimental error, v/hereas the corrections to Cn and Cj would be subtractive and equal to about 1 percent. The Cl and Cd daba were corrected for the support- strut effect; no corrections were aoplied to Cy, Cj, or Cn since previous results indicated the magnitude of these corrections to be sirall for this model and support system. The values of o.i and ^o ^'^^ Cn-u for the model v^ithout wing fillets were obtained by the follovi/ing formulas : ^l'~'n,:, ~ ^n,i; " ( Cn\i/ "•" Cn■^|; ^ ^ "^wing- fuselage combination . wing fuselage^ ^ ^complete model I ^v/mg '^fuselage with hor. and vert, tails on -^ ^. The values of Ai and A^ for Cy,u may be obtained in the same manner. The method used, in this investigation to obtain Ai and A2 is the same as that of reference 5 The follov/ing formula (by which the value of Cn^j for the com.-Dlete model is obtained.) is an example of the application of the increments A]_ and A2: Cn^i, = Cn^ + Cn^ wing fuselage with hor. and vert, tails on The interference betv/een the fuselage and vertical tail and the interference between the fuselage and horizontal tail were not determined. Lift-coefficient and drag-coefficient data for representative model configurations are shovi/n in figure 3« The lateral-stability slooes Cn 1 and Cy. for the wing 8 KACA ARR No. L|3C15a a are presented in figure 14.. The data presented in the figures are summarized in table IV. DISCUSSION For the complete model, the static-lateral-stability slopes Cn,!, and Cy^ usually decreased with a positive increase in angle of attack (figs. 12 and 15) • The results of the present Investigation indicate that this decrease was caused by interference (figs. 5 ^-^d 6). Vlfith the vertical tall off, the variation of these values with angle of attack was irregular apparently also because of interference (figs. 9 ^^^^ 11). Such variations with a of the lateral-stability slopes as were obtained in the present investigation for the low-wing model both with and without a vertical tail -were not shown in the midwlng investigation (reference 5). The slopes Cy^ and C^. were practically always greater for the low-wing than for the mldwing configuration, apparently because of a change in interference with wing location. At negative and sometimes at small nositive angles of attack, Cj, decreased as the angle of attack became less negative (figs. 12 and 15). In the positive angle- of -attack range, Cj^ generally increased with angle of attack. The slope Cj^ was increased because of the side force on the vertical tail at negative and small positive angles of attack but the opposite was true at large positive angles of attack. This effect may be attributed to the system of axes used. For this system of axes, the center of pressure of the vertical tail is above the X-axis at negative and small positive angles of attack; consequently the side force on the vertical tail caused a positive increment of Cj^. The opposite was true at large oositive angles of attack because the center of pressure was below the X-axis. The slope Cj^ was appreciably greater for the midwing configuration than for the low-wing configuration, frequently by as much as 3° oi' effective dihedral (reference $)• This change in slope is evidently caused by a change in the nature of the flow around the wing near the wing-fuselage juncture . NACA ARR No. L5C15a Interference Effects The ircrenents caused "by wing-fuselage inter- ference Ai and by v/ing-fuselage interference on the vertical tail L2. v;ere computed by the equations previously given. The fuselage data (vv'ith and without tail surfaces) used in these coniputations were taken from reference 2. The other data were ob-cained frora the present investigation. The quantities AlCn,. a-^d AiCyi '.vere generally appreciable and had a stabilizing effect on the model (fig. 5)- The variation of these values with angle of attack '-vas irregular, but -^I'-^Yj, generally tended to decrease with a positive increase in angle of attack. The irregularity of the curves may be caused by a burble at the juncture of the wing and fuselage, additional evidence of which may be seen in the curves of lift and drag coefficients in figure J. An appreciabls part of the value of Cy,!, ^or the wing-fuselage combination can be attributed to interference. The changes in AiCn,i; and AiCv. with fuselage length were within the experi- mental accuracy for the fuselage lengths tested. At negative and small positive angles of attack, the quantities A2Cn.| ^i^d A2Cyi vvere generally appreciable and had a stabilizing effect on the laodel (fig. 6). V.'ith an additional positive increase in angle of attack, the values changed in such a manner as to become destabilizing. The effect of replacing vertical tail 2 by vertical tail ]\ (a kS-percent increase in area) on these quantities was generally small in the unstalled range. The variations of A2Cn 1 a.nd AZ'-Yi with fuselage length ivere somev^rhat irregular. Because the model tested in this investigation had no wiiig fillet, caution should be used in applying the results to design since the presence of a fillet may appreciably change the lateral stability characteristics . In view of this fact, an investigation of the lateral stability characteristics of a model with wing fillets might be desirable. 10 NaCa ARR No. L5C13a Effect of Horizontal Tail Theory indicates that the presence of the horizontal tail v.'ould increase the effective asnect ratio of the vertical tail and thus increase Cny, and Cy^i,- A small increase in these quantities was ODtained by the addition of the horizontal tail (figs. 7 ^^d 8). This increment varied somevv'hat irregularly with angle of attack, a coiTiparison with the results of reference 5 showed that the end-plate effect -.vas greater for the midv/ing conlig;u- raticn. Data fro.-n reference 5 indicate that this difference is due to a change in the v.'ing-fuselage inter- ference on the vertical tail with wing location. In reference 3 an incremental increase of 0.0010 in Gy, was Vf computed for the end-plate effect of the horizontal tail on vertical tail L. An average increase of 0.0005 was obtained from the oresent experimental investigation. The end-plate effect jf the horizontal tail on Gj,, airiounted to less than 1'^ of effective dihedral. The results of the present investigation (fig. S) indicate that, although separation begins to occur on the vertical tail at about the same time with the horizontal tail on and off, it progresses more rapidly with tlie horizontal tail on. '/kith the vertical tail off, the magnitude of the static-lateral-stability slopes was not appreciably affected by the addition of the horizontal tail (figs. 9 and 11) . Effect of Changes in Fuselage Length Within the scope of the present investigation, a slight increase in Cn |, "•as generally obtained with a longer fuselage for the model having no vertical tall (figs. 9 t'^ 11). The effect was more pronounced at the larger negative angles of attack. For the complete model equipped with vertical tail J4., the increase in Cn^ v/ith fuselage length was approxi- mately linear (figs, 12 and IJ ) - This increment of Cn,|, v/ith vertical- tail area was greatest in a small region betv/een angles of attack of -ii'- and 0° and decreased as the ang].e of attack varied from this range. This change in vertical-tail effectiveness with angle of attack might be attributed to interference. For the mldvi'lng configuration, the Increases in these slopes with vertical-tail area were also approximately linear and fairly constant over the unst^lied angle-of- attack range (reference y) . As would be expected, at negative and small positive angles of attack, Ci^ increased "with vertical-tail area whereas, at large positive angles of attack, the opposite was true. A similar result Vifas obtained for the midwing conf igui'ation (reference 5). Effect of Changes with Constant Tail Volume In figures l3 and 19 the result of changing the fuselage length and vertical-tail area in such a manner as to hold the tail volume constant is shov/n. The configurations in v;hich the tall volume remained constant are shown in table V. Data from figures l3 and 19 are cross-plotted in figure ll+. The vertical tails tested all had an asoect ratio of 2.I5. The slone Cn,j, should remain approximately the same with constant tail volume. The experimental variation is small over the norm.al flight range and may be partly caused by interference or might be explained by the arbitrary manner in which the tail-voliame coef- ficient was defined. 12 NACA ARR No. L5C13a The values of Ci, and Cv, are dependent mainly on tail area and are practically independent of tail length (fig. llj.) . For the range of variations giving constant tail volume, the changes in both Cy, and Cj, were appreciable . Effects of Changes in Dihedral For the model having no vertical tail, the change in Cn-ij with dihedral angle was small (figs. 9 to 11). Yi/ith the vertical tail on, Cnvi, was slightly larger for r - 0° than for p = 5° (figs. 12 to 17). Similar trends were also obtained for the midwing configuration (reference 3)- Figure ll| shows that increased, dihedral angle slightly decreased the rate of change of Cn^i/ with vertical-tail area but had a negligible effect on the rate of change of Cn,i, with fuselage length. The slope Cy| was generally slightly greater for P = 0*^ than for T = 5^ except at large positive angles of attack . The changes with dihedral angle of wing-fuselage interference and wing-fuselage interference on the vertical tail were small. Comparison of Data from Langley J- by 10-Foot and Langley Stability Tunnels The model tested in the Langley stability tunnel is 0.8 as large and geometrically similar to the one tested in the Langley '/- ^7 10-foot tunnel for the investigation of reference I4.. The test Reynolds niomber, based on the wing chord, was about 619,000 for the Lane,ley 7- ^y 10-foot tunnel compared with about 338,000 for the Langley stability tunnel. The effective Reynolds number, however, was about the sa>ne since the turbulence factor for the Langley 7- by 10-foot tunnel is 1.6, compared v;ith less than 1.1 for the Langley stability tunnel. Data taken from reference Ij. were converted to the stability axes and the angle of attack was corrected for tunnel- v;all effect in order to make the data comparable with data from the Langley stability tunnel. Figure 20 shows NACA ARR No. L5C13a I5 that satisfactory agreement was obtained, in general, for all ttiree static-lateral-rstability slopes. In both tunnels the stall occurred at about the sar-e angle of attack and the model, when yawed, tended to roll violently at the stall. CONCLUSIONS The results of tests of a model consisting of a rectangular low Vv'ing on a circular fuselage v/ith variations in vertical-tail area and fuselage length with and without a horizontal tt.ll indicated, for the range of configurations tested, the following conclusions: 1. The influence of wing-fuselage interf ei-ence on the slope of the curve of yawing-mojaent coefficient against angle of yaw Cn^i, s^nd the slope of the curve of lateral-i crce coefficient against angle of yaw Cy 1, was usually stabilizing, appreciable, and varied with angle of attack. The effect of wing-fuselage interference on the values of Cn 1 and Cy , contributed bv the vertical tail was also generally stabilizing and appreciable at negative and small nositive angles of at back but varied with angle of attack. 2. The end-plate effect of the horizontal tail slightly increased the efficiency of the vertical tall. The experimental Increment obtciined was only one-half the computed value. 3. Increasing the fuselage length with no vertical tall resulted in a slight increase in Cn^ for the model, both with and v;ithout a horizontal tail. At the larger negative angles of attack, the effect was more pronounced. For the complete m.odel, the incre^^se in 0^,1; was approximately linear with fuselage length. The magnitude of the increase appreciably diminished with a positive increase in angle of attack. The changes in the slope of the curve of rolling -moment coefficient against yaw Cj^ and in Cy^ with fuselage length v/ere small. 1J4. NACA ARR No. L5C15a 1|. The slopes Cn,]; smd CYiI; increased approxi- mately linearl;/ v/ith vertical- tail area. For the system of axes used, C j i increased with vertical-tail area at V negative and small positive angles of attack but the opposite v/as true at large positive angles of attack. Increased dihedral angle slightly decreased the rate of change of Cn,i, 'vvith vertical-tail area hut had a negligible effect on the rate of change of Cn^ with fuselage length. Except at large positive angles of attack, Cy-(!/ w£^s greater vulth the smaller dihedral angle Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va . KACA ARR No. L5G15a I5 REFERENCES 1. Ziiraneriran, Charles H.: An analysis of Lateral otatil.lty in Power-Off Plight with Charts for Use in Design. NACA Rep. No. 569, 1957. 2. Pehlner, Leo F., and LacLachla:i, Robert: Investi- gation cf '-Effect of Sl'-Ieslip on Lateral Stability Characteristics. I - Circular Fuselage with Variabions in Vertical-Tail Area and Tail Length with and without Horizontal Tail Surface. HACA ARR ^io. LI4L25, l?kh' 5. Eollingworth , 'iliomas A.j Inveatigaticn cf Effect of oidesilp on Lateral Stability Characteristics. II - Rectangular I.Iidwing on Circular Fuselage with Variations in Vertical-Tail Area and Fi'.sclage Length with and without liorizontal Tail Surface, NACA ARR No. L5CI5, 19hJ}. '4. Barnber, I.;. J., and Rouse, R. 0.: Vvlnd-Tunnel Investi- gation of Effect of Yaw on Lateral-'-^tability Characteristics. II - Rectangular II.A.C .A ." 23012 Wing with a Circular Fuselare and s Fin. NACA T'cl iJo. 730, 1935. 5. V/allace, Ai-thar R,, and Turn^i", 'Tlioinas il.i w'ind- Tunnsl Investigation of Effect oi^ Yaw on Lateral- Stability Characteristics. V - SyrfiTije trie ally Tapered •i'lng with a Circular Puselar--e Having a Horizontal and a Vertical Tail. IIACA ARR 'io . 3F23, I9I3. NACA ARR No. L5C13a 16 TABLE I FUSELAGE DIMENSIONS Fuselage Fuselage length (in.) Tall-cone length (in.) Tail length, I (in.) , , Tail length l Wing span ' b Short 52.25 9.35 20.07 0.i|l8 Medii,im 37.05 111 -65 2I+.87 .513 Long kl.35 19.^5 29.67 .618 TABLE II TAIL-SURFACE DIMENSIONS Vertical- Tail surface Designation tail area (sq in. ) Vertical-tail area vVing area Aspect ratio (1) Vertical 1 10.83 0.0300 2.15 Do--- 2 25.78 .0659 2.15 Do 3 28.37 .0786 2.15 Do--- h 55-16 .097i| 2.15 Do--- 5 ^6.20 .1280 2.15 Horlzontal 6Ji.21 .178 3.99 lA rea measured from root chord at center line of fuselage NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NAGA ARR No. L5C13a 17 TABLE Til MODEL COMBINATIONS TESTED Dihedral Horizontal Vertical x^selage angle Variable | tail tail (aeg) ! Off 1 2 Short , medium. a 3 and long li 5 On 2 and 5 Long \!/ k 5 tl e c 1 um k h Short Off a and \[/ Off Off Long tt Off 5 ^1' ^ Short and 5 a and \J; NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ARR No. L5C13?. TABLE IV PRESENTATION OP RESULTS Figure 3 k 5 6 7 8 9 10 11 13 Description of figure Data Dresented Lift and drag curves for repre- sentative model conf i^ijUrations Slope of yawing-moment and lateral-force coefficients for NACA 23012 rectangular v;ing Effect of wing-fuselage interference Effect of wing-fuselage interference on vertical tail End-plate effect of horizontal tail End-plate effect of horizontal tail Effect of changing fuselage length (no tail surfaces) Effect of changing fuselage length (no tail surfaces) Cl and I Cd a-s f ( a) ic n\i/ and CYii; ^s f(a: UlCn.i^ and AiCy^ as f (a) ^^2Cn,i; and /^ZCy^, as f(a) On^, C^, and Cy,|/ as f(u) Cn, Cj, and Cy as f(\l/) Cn^) '^l^\f> and Cy,I; as f(a) On, Ci, and Cy as fi'it) Effect of changing fuselage length ! Cj-^,,^, Ci^, and (horizontal tail on; vertical j "Cyi, as f(a) tail off) ^ Effect of changing fuselage length (horizontal tail and vertical tail I4. on) m-j/ > ^l^> and CY^ as f(a) Effect of changing fuselage length; Cn, C 1 , and (horizontal tail and vertical | Cy as f ('Jf) tail [|. on) I NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ARR No. L5C13a 19 TABLE IV - Concluded PRESENTATION OF RESULTS - Concluded Figure 15 16 17 13 19 20 Description of figure II4. Effect of changing fuselage length Effect of changing vertical- tail area Increment of slope of yawing- moinent coefficient against angle of yaw caused by vertical-tail area Effect of changing vertical-tail area Effect of changes .vith tall volume constant Effect of changes with tail volume constant Comparison of data from Langley stability and Langley 7- by 10-foot tunnels Data presented Cy,.. as 1 and V'^W Cy^ as f(a) ''-^ '' 'fe Cn, Cj, and Cy as f (\^) ^n^ll> C^. and Cy^i; as f(a) Cn, Cj, and Cy as f (\!/) ^n-^> ^l\if> and Cy^ as f(a) NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ARR No. L5C13a 20 > c/D .? 1-0 X> -p 03 a H ^ •H cc: < tM C^ t^ t>- s CsH 1-/3 C:i (!) -^ -^ J M « > W w • * " --i^-^ § J (-:) 3 -a; fo rH IS > > w M W i-q EH EH M r-i =t: Eh < «^ a M E-J S! i ^ Eh > r- CO rjO CO a •> cd a -^ ^ a-. > r-^ bO =1; j; •H fl Eh w C!3 Eh •H CO '^ b M r^lP EH <; •, -P r c bO cO GO ■00 cc- M c! fX iH rH rH [■--, CD m -:i- U^ MD S rH • C bO •rH ^:i •■•ti W [H bO jj ^ rH fn •■H fcO 03 T3 ti CO ,a > 1 NACA ARR NO. L5C13a Fig. 1 ■1.15 ff 1.47 f? — —2.03 Figure l.~ Reciangular NACA 230/2 win<^ in combination with circular fuselage , yeriical and horizontal fails, and loil cones ■ /\ll dimensions given in inches • NACA ARR No. L5C13a Fig, T3 G 03 +J >> U *J o -H x: i-H CO -H J3 x: Cd <-> «-> ■H ro s >i T3 OJ (U ■-{ P. bo P. C ■H 03 3 J cr 0) C — ) .-( 0) ra T3 +J O (Q E 0) +3 tlD c t-1 •H o 1 v-l s t3 o QJ r-\ +J 1 C u 3 cd O M 6 3 bo lO C n3 -H +j • H o 03 (D ♦J cc: rH 1 «3 • CJ CM --1 ■H> 0) u U 0) 3 > bo NACA ARR NO. L5C13a Fig, /.£ /.O .8 ^■^ o 4: -.4 -.6 -.8 Fuselage. Horiz.o/7to/ Verf/caJ- to// to// o Shorf Off t S^orf On X Shorf On D l\Aec/iu/T> On o Long' On Off Off -/z Angle of affaoAj cC j deg Figure 3.-\/ariation of lift and drag coefficients wi/A> ang/Ze of attack for represeniafive model corifigurations. IZ -^"j y, 0°- c^,6Slb/sqff. Fig. 4 NACA ARR No. L5C13a «.S ^ ^ ^ 8 c: C5 . V) ^ 00 \; y Ci O U I Fig. 8a NACA ARR No. L5C13a Hor/zonfc^/ Vett/co/ fa// fa// -ZO -/O Ang/e of yav\/ , fa) r , O (yeg> o/i /^/gure 8 . - £ffecf of /lOf^/zonfa/ to// ysurface vor/af/on of yaw/ng - mo/rje/^f ■> ro///ng - /r)o/^enf , and /of era! - force coeff/c/enfs w/f/o ong/e of yoi^ . S/)ort fuse/age y q , 65 /b/sg ff . a , /0.^' NACA ARR NO. L5C13a Fig. 8b ^ ^ >. <)> ^ .■N. ^*v ^ V ^ Q >- O Hor/ zon to/ Ve rffCa/ fat/ fa// . Off Off o Off 4 + On 4 0* I Ang/e of yow , 1/^ , c/eg ft) r , 6\ /^/gure 6 . - Conc/uded . Fig. 9 NACA ARR NO. L5C13a Q w < 5§> '\ .^ ^"^ ^t ^ <>^ ^ l^-b .ooz -.004 •00£ ^/ ^ ■0/ ^n ^ o Short X lor?p n S/7orf O O o L//7e of co/?sf^/7f to// i/o/i//77ej0.0^07 Sy/S, H' (a) a ^ 0°, Figure 14 .—E-f-fect of ahang/nq fuseJcrge /engfh on var/afion of /otero/-sfabi/ity s/opes ^h^j ^ixy) one/ ^y^y ^'^^ Mertical- tail area. Hor/zonta/ to// on 5 c^, 65 /b/sq ft. NACA ARR No. L5C13a Fig. 14b D/hedro/ ideg) -- 5 Fc/jff/oge C, yy^ ^/, -■00^ -.004 .00£ .0/ l/r>e of (Tonsfaof fa// vo/i/^e, 0-0407 C n. o S/yorf + Med/c/rr? Long a Shorf o Med/u/7? A lo/7g D//7edra/ (deg) O J S •04 .06 •/£ '16 (jb) a,/0°. F/gore /4 - Co/7c/c/ded ■ Fig. 15a NACA ARR No. L5C13a ^ ^ 8 X I o .0 ^ ^ S^:^ ^ 'o iv <; c NACA ARR No. L5C13a Fig. 15b bO ^ ^ ^ ^ 8 -N ^ % -^ X ^ ^ o V ^ ^ k^ b 1 ^ o 1 1 1 ^t k Fig. 16 NACA ARR NO. L5C13a ^^/73^ -00^ -•004 -.006 ■04 .08 ./£ •/6 Syy^ Sy^ Figure 16 .- Increnoent of /af-ero/-sfalj////'y s/ope ACn-y, cac/se>d by yert/cal - to// area. (Cp. for mea//urr?-/engfh fusehge w/fh hor/zonfo/ to// and wing with 0° d/hedra/ subtracted from ^^ for comp/ete mode/ to obta/n ^Cn^) q , 65 tb/sq ft . NACA ARR No. L5C13a Fig, 17a ■♦s ^^ ^ ^ «\ 5- >. ;> r !^ ^ ; *si -30 '20 -/O fO A/7p/e> of yaw , >^ ^ de^ fo) r , 0^. ^/gure /7 . - Effecf of cha/7g//^p ^er^/co/ - to// area or> var/af/on of yow/ng - noomenf , ro///ng /nomerjf , and /aferal -force coeff/c/enfs w/fh angle of yaw . A/fed/um-Zength fu^e/oge w/fh fior/zo/ifa/ fa// ; a , /O.Z" -, q , 66 fb/sq. ft. Fig. 17b NACA ARR No. L5C13a V . 02 V." ,^ » V i (b ^ o-Z>^ >- <0 Jo ./o l/ert/ca/ fa// -.zo -30 -ZO Ang/e of yaw , Ih , c/e g (d) r /^/gure /7 . - C onc/uc/ed - NACA ARR NO. L5C13a Fig. 18a ^ <\ ^ C» Qi Q) ^ r « ;V v: rO ^ 1 <») ^ ^ i § «> 5 fa// S .02 -m -30 -ZO -/O O /O Ang/e of yaw ^ 'X' 5 deg (a) r , *. F/gure /9 . - Effecf of 'Se^/era/ cornb/nofions ha\//n^ consionf fa// volume o/o yor/afion of yav\/tng - moment , /-o/f/ng- moment ^ancf /atero/ -force coefficients w/fh ong/e of yc/w 0.0407-^ honizonta/ to// on ; ac i /O.Z' To// vo/ume j q.,6S/dfsg ft. Fig. 19b NACA ARR No. L5C13a v ^■UA. ^ ^ » V3 $ ^ ^ •^ S-^>a •5 >^0 <0 5 -^0 ta// 4 J .oz ^^ Ang/e of yaw j 9^ , cfsgf (b) r , 3\ f/gure /9 . - Conc/uded . NACA ARR No. L5C13a Fig. 20 Q) I S .1 (J F ° • 6 ^ -^ u< «; Jii ?5 l~ ^ c ^ & s ^ ^^ < s: o Q *=» ^ bk nT ^ I 8, j^S I- UNIVERSITY OF FLORIDA 3 1262 08104 942 ..J .(UA ^ u... UN. S DEPARTMENT 20 MARSTON SCIENCE UBRARY KO. BOX 117011 GAINESVftlE. FL 3261 h70t X USA UNIVERSITY OF FLORIDA DOCUMENTS DEPARTMErviT 120 MARSTON SCIENCE UBRARY P.O. BOX 11 70 11 GAINESVILLE. FL 32611-701 1 USA