Nfttfr L-y S I NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED April 19^5 as Advance Restricted Report L5CI3 INVESTIGATION OF EFFECT OF SIDESLIP ON LATERAL STABILITY CHARACTERISTICS II - RECTANGULAR MHWING ON CIRCULAR FUSELAGE WITH VARIATIONS IN VERTICAL-TAIL AREA AND FUSELAGE LENGTH WITH AND WITHOUT HORIZONTAL TAIL SURFACE By Thomas A. Hollingvorth Langley Memorial Aeronautical Laboratory Langley Field, Va. NACA |£~ 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-S DOCUMENTS DEPARTMENT Digitized by the Internet Archive in 2011 with funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/investigationofeOOIang NACA ARR No. L5C13 RESTRICTED NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE RESTRICTED REPORT INVESTIGATION OF EFFECT OF SIDESLIP ON LATERAL STABILITY CHARACTERISTICS II - RECTANGULAR MIDWING ON CIRCULAR FUSELAGE WITH VARIATIONS IN VERTICAL-TAIL AREA AND FUSELAGE LENGTH WITH AND WITHOUT HORIZONTAL TAIL SURFACE By Thomas A. Hollingworth SUMMARY Power-off tests were made in the 6- by 6-foot test section of the Langley stability tunnel to determine the variation of the static lateral stability characteristics with vertical-tail area, fuselage length, and wing dihe- dral. Two NACA 23012 rectangular wings with rounded tips and dihedral angles of 0° and 5 were tested alone and in combination with three circular fuselages of different lengths. The wing-fuselage combinations were tested as midwing monoplanes with and without a horizontal tail and with variations in vertical-tail area. The results are presented as curves showing the variation of the static- lateral-stability slopes with angle of attack, and the rolling-moment, yawing-moment, and lateral-force coeffi- cients with angle of yaw. The results indicated that the wing-fuselage inter- ference on the slope of the curve of yawing-moment coef- ficient against angle of yaw C n , and on the slope of the curve of lateral-force coefficient against angle of yaw Cy was small and remained practically constant over the unstalled angle-of-attack range. In the high- speed flight range, the wing-fuselage interference on the vertical tail was small and, in the normal flight range, was not appreciably changed by an increase in fuselage length or vertical-tail area for the sizes Investigated . RESTRICTED NACA ARR No, L5C3-3 With no vertical tail, increased fuselage length caused a negligible change in C n . for the fuselage lengths tested. For the complete model, C increased approximately linearly with fuselage length. The slopes C n and Cy,, increased approximately linearly with vertical-tail area. For the system of axes used, the slope of the curve of rolling-moment coefficient against angle of yaw Cj. increased with vertical-tail area at negative and small positive angles of attack out decreased at large positive angles of attack. The results also indicated that increased dihedral angle slightly decreased the rate of change of C n , with vertical-tail area but had a negligible effect on v the rate of change of C n - with fuselage length. An appreciable increase in C~ . was caused by the end-plate effect of the horizontal tail on the vertical tail. INTRODUCTION The trend toward greater speed and higher wing loadings and the increased consciousness of the importance of satisfactory flying qualities have resulted in additional attention being given to handling characteristics in air- plane design. Mathematical equations and convenient charts for predicting the lateral stability characteristics are given in reference 1. In order to use this material, however, it is necessary to know the stability derivatives, which vary with each airplane configuration. A seri.es of investigations has therefore been undertaken in the Langley stability tunnel to determine the variation of both the static-stability and rotary-stability slopes with various -airplane parameters. The present investigation is a continuation of the investigation described in reference 2 except that, for the present tests, the fuselage was equipped with a rectangular wing in the midposi tion. The purpose of the investigation, which was conducted in the o- by 6-foot test section of the Langley stability tunnel, was to determine experimentally the effect, with the propeller off, of vertical-tail area, fuselage length, wing dihedral, interference, and the presence of the horizontal tail on NACA ARR No. L5C13 the static lateral stability characteristics. P geomet- rically similar model has been tested in the Langley 7~ by 10-foot tunnel (reference 3) and the data may be used to correlate the results in the two tunnels. Tests were made of a model that had dimensions pro- portional to those of the average airplane. The ratios of fuselage length to wing span and of vertical-tail area to wing area investigated were taken to bracket the range commonly used on present-day airplanes. APPARATUS AND MODEL The tests were made in the 6- by 6-foot closed- throat test section (adjusted for straight flow) of the Langley stability tunnel. A three-view drawing of the model tested, which was constructed of laminated mahogany, is given in figure 1. Figure 2 shows the model mounted on the three support struts for tests in the tunnel. The two rectangular wings used for the tests have dihedral angles of 0° and 5° and, in elevation, the maximum upper- surf ace section ordinates are in one plane. Each has an aspect ratio of 6.I4. and an area of 3°1 square inches, which includes the portion inside the fuselage. The NACA 23012 profile is maintained along the entire span. The fuselage is of circular cross section and was constructed as described in reference 2. Its dimensions are presented in table I. With the shortest tail cone attacked, the fuselage is geometrically similar to that of reference 3- Five interchangeable vertical tails and the horizontal tail were made to the IJACA 0009 section (fig. 1). Their dimensions are presented in table II. TESTS The wings with dihedral angles of 0° and 5° were tested alone at angles of yaw of -5 and 5° over an 1+ NACA ARR No. L5C1J angle-of-attack range from -10° to 20°. The model combina- tions tested are shown in table III. Model combinations were tested at angles of yaw of -5° , 0°, and 5° over an angle-of-attack range from -10° to 20° and at angles of attack of -0.2° and 10. 3° over an anrle-of-yaw range from -30° to 12°. Tests in which the angle of attack was varied were run at a dynamic pressure of 65 pounds per square foot, which corresponds to a Reynolds number of approximately 838,000 based on an 8-inch wing chord. Tests in which the angle of yaw was varied were run at a dynamic pressure of 65 pounds per square foot at an angle of attack of -0.2° and at ko pounds per square foot, which corresponds to a Reynolds number of about 5^6,000, at an angle of attack of 10. 3° to minimize the possibility of compressibility effects at large angles of attack. The rolling-moment data are not presented for a few tests, because the tare readings were inconsistent. PRESENTATION OF DATA The results of the tests are presented in standard NACA coefficients of forces and moments. Rolling-moment and yawing-moment coefficients are given about the center- of-gravity location shown in figure 1. The data are referred to the stability axes, which are a system of axes having their origin at the center of gravity and in which the Z-axis is in the plane of symmetry and perpen- dicular to the relative wind, the X-axis is in the plane of symmetry and perpendicular to the Z-axis, and the Y-axis is perpendicular to the plane of symmetry. The coefficients and symbols used are defined as follows : C L lift coefficient (L/qS w ) Z-q drag coefficient m/qS w ) C-cr lateral-force coefficient (Vqs w ) Cy slope of curve of lateral-force coefficient against angle of yaw (dCy/d\!/J Cj rolling -moment coefficient (L'/qoS^ NACA ARR No. L5C13 Cj slope of curve of rolling-moment coefficient ^ against angle of yaw fdCj/c^ C n yawing-moment coefficient (ll/^^yj) C n . slope of curve of yawing-moment coefficient V against angle of yaw (bC n /bty\ A-. increment of C n , or Cy. caused by wing-fuselage interference A ? increment of (V or C v caused by wing-fuselage interference on vertical tail tail-volume coefficient D force along X-axis; positive v/hen directed downstream Y force along Y-axis; positive when directed to the right L force along Z-axis; positive when directed upward N yawing moment about Z-axis; positive when tends to retard right wing L' rolling moment about X-axis; positive when tends to depress right wing q dynamic pressure (tP v ) V free-stream velocity p mass density of air S w wing area (2.507 sq ft) b wing span (I4. ft) T dihedral angle, degrees I tail length; measured from center of gravity, which is assumed t o be IO.I4.O inches behind nose of model on center line of fuselage, to hinge line of tail surface NACA ARR No. L5G13 S v vertical-tail srea a angle of attack, degrees ii angle of yaw, degrees The static-lateral-stability slopes c„ , C 7 , and Cy . were obtained from data measured at \J/ = ±5° since the yaw tests showed that the coefficients had an approximately linear variation in the range of angle of yaw from 5° to -5 • In order to indicate the validity of this procedure, the slopes obtained from yaw tests at \Jr = 0° are plotted with tailed symbols in the figures. The accuracy of C n , On, and Cy was determined experimentally to be about t'o.OOO^, ±0.0008, and ±0.001, respectively, at a dynamic pressure of 65 pounds per square foot. The average experimental accuracy of n , C Z| , ana Cy, is about ±0.00010, ±0.000l6, 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-wall effect by the following formulas: 5 Aa = 57-5%. ~C L =0.609C L (deg) c ACr, = 6_ Oj - O.OlOoC/ where 5 jet-boundary correction factor at wing (O.I525) C cross-sectional srea of tunnel (36 sq ft) Both corrections are additive. ¥.0 jet-boundary correc- tions were applied to Cj, n , and Cy The correction to Cy is within the experimental error, whereas the corrections to C n and C7 would be subtract! ve and equal to about 1 percent. NACA ARR No. L5C13 7 The C T and C n data were corrected for the support- strut effect: no corrections were applied to Cy, C ? , or C n since previous results indicated the magnitude of these corrections to be small for this model and support system. The values of At and A~> for n . for the model 1 d. n\j/ without wing fillets were obtained by the following formulas : V^lT ^ -/c +c n) \ v wing-fuselage combination I ^wing ^fuselage] ^n^n \l/ I "' n \!/ ^complete model V y win| is + ^ +A 1% fuselasce with hor. and vert, tails on 7 The values of A, and A-, for C v may be obtained in J- <- ity the same manner. The method used to obtain A-, and Ap the same as that of reference Ll. The following formula (by which the value of C- for the complete model is n \i/ obtained) is an example of the application of the incre- ments A]_ and k^' c n Ur = G n^ + c n,|, * *vn.ng ^fuselage with hor. and vert, tails on + *1% + *2% The interference between 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 shown in figure $. The lateral-stability slopes n and C, r for the wing 3 NACA ARR No. L5C13 are presented in figure [|.. The data presented in the figures are summarized in table IV. DISCUSSION The static-lateral-stability slopes C end C v remained practically constant over the unstalled angle- of-attack" range (figs. 9 to ly , 15, and 16). With the system of axes used, the center of pressure of the vertical tail varied with respect, to the X-axis. At negative and small positive angles of attack, the center of pressure was above the X-axis and, therefore, the side force on the vertical tail caused a oosltive increment of C 7l . The opposite was true at large positive angles of attack, since the center of pressure of the vertical tail was below the X-axis. The jags in the curves of lateral-force, rolling- moment, and yawing-moment coefficients noted in figures 0, 15, l6, and 18 can probably be attributed to vertical- tail stalling. Interference Effects The increments caused b"" wing-fuselage interference A-, and by wing-fuselage interference on the vertical tail A_ were computed by the equations previously given. The fuselage data (with and without tail surfaces) used in these computations were taken from reference 2. The other data were obtained from the present investigation. The magnitudes of A,C,~ and i r v are small and 1 n \|/ x - \ir remained practically constant over the unstalled angle- of -at tack range (fig. ?)• T^- e change in tha magnitude of these quantities with fuselage length was within the experimental accuracy for the fuselage lengths tested. Both AjCy, , and AoC,, varied aopreciably with angle of attack but their magnitudes were small in the high-speed flight range. (See fig. 6.) c eplaci vertical tail 2 by vertical tail h. (a i|S-percent NACA ARR No. L5C13 9 increase in area) orly slightly changed the magnitude of these quantities in the normal flight range. As indicated by previous experimental data (reference tj.), the varia- tion of A^C- and AqC v with the fuselage lengths tested was generally small in the normal flight range. Effect of Horizontal Tail Theory indicates that the presence of the horizontal tail increases the effective aspect ratio of the vertical tail and thus increases C and C,, (reference S)- A pronounced increase in these quantities was obtained in the present investigation by the addition of the horizontal tail. This increase diminished somewhat with a positive increase in angle of attack. (See figs. 7 and 8.) A correlation of the results of previous airfoil tests in the Langley stability tunnel indicates a value of 0.105 for the section lift-curve slope of an NACA 0009 airfoil. by substituting this value for the theoretical section lift-curve slope of 0.109 in equa- tion ([}.) of reference 6 and by the use of figure 5 i n reference 5> an incremental increase in C v of 0.0010 was computed for the end-plate effect of the horizontal tail on vertical tail I4.. An average experimental increment of 0.0010 was obtained for the model with a dihedral angle of 0° and 0.0015 for the model with a dihedral angle of 5 . The end-olate effect of the horizontal tail on C, amounted to about 1° of effective dihedral. With the vertical tail off, the magnitude of the static-lateral-stability slopes was not appreciably affected by the addition of the horizontal tail. (See figs. 9 and 11. ) Effect of Changes in Fuselage Length Within the scope of the present investigation, a negligible increase in C„ , was obtained by increasing the fuselage length for the model with no vertical tail. (See figs. 9 to 11.) For the complete model equipped with vertical tail I4., the increase in C„ with fuselage 10 NACA ARR No. L5C13 length was approximately linear and fairly constant over the unstalled angle-o.f -attack range. (See figs. 12 to ll\. ) The variation in C 7 and C Tr was small both with and without s vertical tail for the fuselage lengths tested. Effect of Changes in Vertical-Tail Area The increases in C n and C v with vertical-tail area were approximately linear and the magnitudes were nearly constant over the unstalled angle-of-attack range. (Gee figs. lk to 16. ) As would be expected, at negative and small positive angles of attack, C 7 increased with vertical-tail area whereas, at large positive angles of attack, Ci decreased with increased vertical-tail area. Effect of Changes with Constant Tail Volume In figures 17 and 18 the result of changing the fuselage length and vertical-tail area in such a manner as to hold the tail volume constant is shown. The configu- rations tested in which the tail volume remained constant are shown in table V. Data from figures 17 and lS are cross-plotted in figure lli.. All the vertical tails tested had an aspect ratio of 2.15> The slope D r should remain approximately the same with constant c 'tail volume. The small experimental variation is possibly caused by interference or might be explained by the arbitrary manner in which the tail- volume coefficient was defined. The values of C 7 and C, r are dependent mainly on vertical-tail area and are practically independent of tail length (fig. 1I4. ) . For the range of variations giving constant tail volume, the change in 3^ was not more than about 0.0002, which is equivalent to about 1° of effective dihedral. NACA ARR No. L5C1J 11 Effect of Changes in Dihedral The slope C v generally was slightly greater for V = 5° than for T = 0°. (See figs. 9 to 13, 15, and lb. ) iA'itr the vertical tail off, the change in C n| with dihedral angle was insignificant (figs. 9 to 11) but, with the vertical tail on, C n was slightly larger for T - 0° than for r = 5° (figs. 12 to l6). Figure ill shows that increased dihedral angle slightly decreased the rate of change of C~ with vertical-tail area but had a negligible effect on the rate of change of C- with fuselage length. The change with dihedral angle of wing-fuselage interference and wing-fuselage interference on the vertical tail was small. Comparison of Data from Langley 7 - 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~ by 10-foot tunnel for the investigation of reference 3» The test Reynolds number, based on the wing chord, was about 619,000 for the Langley 7~ by 10-foot tunnel compared with about 838,000 for the Langley stability tunnel. The effective Reynolds number, however, was about the same since the turbulence factor for the Langley 7" by 10-foot tunnel is 1.6 compared with less than 1,1 for the Langley stability tunnel. Data taken from reference 3 were converted to the stability axes and the angle of attack was corrected for tunnel-wall effect in order to make the data comparable with data from the Langley stability tunnel. Figure 19 shows that satisfactory agreement was obtained for all three static-lateral-stability slopes. In both tunnels the model, when yawed, tended to roll violently at the stall. CONCLUSIONS The results of tests of a model consisting of a rectangular midwing on a circular fuselage with variations 12 NACA ARR No. L i ;Cli in vertical-tail area and fuselage length with and with- out a horizontal tail indicated, for the range of con- figurations tested, the following conclusions: 1. The wing-fuselage interference on the slope of the curve of yawing-tnoment coefficient against angle of yaw G-n , and the slooe of the curve of lateral-force coefficient against angle of yaw C v was small and remained practically constant over the unstalled angle - of-attack range. In the high-speed flight range, the wing-fuselage interference on the vertical tail was small and, in the normal flight range, was not aporeciably changed by fuselage length or by an increase of about I4.8 percent in vertical-tail area. 2. The end-plate effect of the horizontal tail on the vertical tp.il appreciably increased C„ , and C v ,- n^ i\j/ Good agreement was obtained between experimental and commuted values of 0, r . 3. Increasing the fuselage length with no vertical tail resulted in a negligible change in C n for the model, both with and without a horizontal tail. r or the complete model, the increase in C n was approximately linear with fuselage length. The changes in the slope of the curve of rolling-moment coefficient against angle of yaw Ct, and in C Y with fuselage length were small. u. The increases in Gv, and Cv with vertical- tail area were approximately linear. For the system of axes used, an increase in vertical-tail area increased G 7 at negative and small oositive angles of attack but the opposite was true at large positive angles of attack. 5. Increased dihedral angle slightly decreased the rate of change of C n . with vertical-tail area but had a negligible effect on the rate of change of n , with il \J/ f u s e 1 a g e length. Langley ^en-orial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. NACA ARR NO. L^CIJ REFERENCES 1. Zimmerman, Charles H. : An Analysis of Lateral Stability in Power-Off Flight with Charts for Use in Design. NACA Ren. No. 589, 1937- 2. Fehlner, Leo p . , and I'acLachlan, Robert: Investi- gation of Effect of Sideslip on Lateral Stability Characteristics. I - Circular Fuselage with Variations in Vertical-Tail Area and fail Length with and without Horizontal Tail Surface. NACA ARR No. li|E25, iskh. 3. Bamber, M. J., and House, R. 0.: Wind-Tunnel Investi- gation of Effect of Yaw on Lateral-Stability Characteristics. IT - Rectangular N.A.C.A. 23012 Wing with a Circular Fuselage and a Fin. NACA TN No. 730 , 1939. I4.. Recant, Isidore C, and Wallace, Arthur R. : Wind- Tunnel Investigation of Effect of Yaw on Lateral- Stability Characteristics. Ill - Symmetrically Tapered Wing at Various Positions on Circular Fuselage with and without a Vertical Tail. NACA TN No. 825, i9l.1l, 5. Katzoff, S. , and Kutterperl, William: The End-Plate Effect of a Horizontal-Tail Surface en a Vertical- Tail Surface. NACA TN No. 797, 19hl. 6. Jones, Robert T. : Theoretical Correction for the Lift of Elliptic Wings. Jour. Aero. Sci . , vol. 9j no. 1, Nov. I9I4.I, pp. 3-10. NACA ARR No. L5C13 iU TA3LE I FUSELAGE DIMENSIONS Fuselage Fuselage length (in. ) Teil-ccne lenrth (in. ) Tail lenpth, I (in.) Tail length I Wing span D Short 52.25 9.£5 20.07 0.1+18 Medium 37.05 lU.65 21+.87 .518 Long la. 85 19.15 29.67 .618 TABLE II ^ail-sup^ace Li E7s:o:; r Tail surface Designation Tail area ( sq in. ) (1) Tail area "."ing srea Aspect ratio Vertical 1 IO.83 0.0500 2.15 Do— 2 25.78 .0659 2.15 Do 3 28.37 .0786 2.15 Do k 35.16 .097*1 2.15 Do- 5 U6.20 . 1280 2.15 Horizontal 6^.21 .178 3.99 "Area measured from root chord at center line of fuselage, NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 15 NAC;\ ASH No. L'-'.C >-■; TABLE III MODEL COMBINATIONS TESTED [T = 0° and 5°] Horizontal tail Vertical tail Fuselage Variable On Off Short, medium, and long a 1 2 3 h 5 2 Long * u 5 Medium h h Short Off Off a and ^ Off Long k Short NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ARR No. L5C13 TABLE IV PRESENTATION OF RESULTS 16 Figure Description of figure Data "O resented 5 6 7 8 9 10 n 12 13 Lift and drag curves for repre- sentative model configuration Slope of yawing -moment and iateral-f orce coefficients for NACA 23012 rectangular wing Effect of wing-fuselage inter- ference Effect of wing-fuselage inter- ference on vertical te.il End-olate effect of horizontal tail End-olate effect of horizontal tail Effect of changing fuselage length (no tail surfaces) effect of changing fuselage length (no tail surfaces) Effect of changing fuselage length (horizontal tail on; vertical tail off) Effect of changing fuselage length (horizontal tail and vertical tail I4. on) Effect of changing fuselage length (horizontal tail and vertical . tsl Lik_QnJ G-r and C D as f(a) % and C y ^ as f(a) Al°n^ and - \j/ A 2 C nv(/ and A 2 Cy as f(a) Cn r Cty, and C v , as f (a) C n , C z , and Cy as f(\|/) %' C V and C y as f(a) C n , Cj, and Cy as f(\|/) c n^> c l$ and C Y . as f(a) c n r O lr and C Vl as f(a) ^n ' ^ J ' a1 " 1 '^ C Y as f(ii/) NATIONAL ADVISORY COMMITTEE FOF AERONAUTICS 17 NAG A ARR No. L5C13 TABLE IV - Concluded PRESENTATION OP RESULTS - Concluded Figure Description of figure 1 Data presented 11+ Effect of changing fuselage length C Y as f ~~ 15 Effect of changing vertical-tail area 2 n,'/ G lf and C v , as f(a) 16 Effect of changing vertical-tail c n» c l> and area G y as f(^) 17 Effect of changes with tail volume G n r G k\,, ^id G Y y as f(a) constant IS Effect of changes with tail volume C n , Ct,, and constant C Y as f(i') 19 Comparison of data from Langley ^n,'.' c Z-,..» an( l stability and Langley 7- by 1 10-foot tunnels C Yx{ , as f(a) NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ARR No. L5C1J B CO Eh *^l & O > PI H <^ E-i Eh CO & O o & > CO o M Eh ID es H I o o p; W P O >l g CO Iw ro |,Q ^ -P C CD ■r-J CJ •H O- C— F~ S 3 .— i o > | rH «H OS Eh >| S CO |co ^ CO CO -3- *vO CTn CD © r- CO Lf> £h f- rr- !>- ^O «S cO O o o • • • H bO o •H c a •H eh "H P *\ £ -p c bO CO CO co CO C P\ rH r-t r-^ (D CO -3" LT\ ■>£> H bO • o • • H c •rl jH En © tu) CO -p | rH u •H bO o 13 s CO a CD o CO hH & rH CO O rH •rl iH _^f KA t\] -P CO f-> -P CD ►> ►~ -- ,....- CO o H >H O C !SI w o M K +J CO 3 >> cu cu bo sx c P. CT> ■H J 3 a* c CU -H rH CO cu -.-J -a w o cu S -r> bo Ij C O •H S-j OJ > NACA ARR No. L5C13 Fig, Fus&loge o Long o Long o Long *■ Medium « Short Hor/7-Onto/ toil Off On On On On Vert icol toil Off Off 5 5 5 -4 4 & l& Angle of attack j cC , d&g figure J .- Var/of/'or? of //'ft or^f drap coeff/c/er?fs w/fh an?/e of affac/f for representat/ve mx/e/ ror?f/^4/raf/ar?s ■ f 7 , f"> ■*■ , 0° ) $ , 63- //>/s? ff ■ Fig. 4 NACA ARR No. L5C13 v© N & |S Q Q -5S Qi S ^ ^ ^ > 1 ^ <*> . ^ 1 It 3 1 s NACA ARR No. L5C13 Fig. 5 "0 « u Q) Q 5; Si I I i ^\* V. .N 5 <£ Fig. 6a NACA ARR No. L5C13 *> N CO ^ i ^ 1 s 5 1 3 s: ** ^ $ ty NACA ARR No. L5C13 Fig. 6b O ^ ^ S^ <0 I 9 to v. O r?fs w/f/? anq/e of yaw ■ Sf>orf fuse/ope w/fh vert/ca/ fa// 4 • NACA ARR No. L5C13 Fig. 8b tf »» 5 "K s 1 * ^ £ * ft » 5 .$ c < X t> Ar/g/e of yow , ?" , deg Cb) r , 0° ; oc , /O.J° i a , 40 /b/sa ft r~/gure 8 . - Cont/nued . Fig. 8c NACA ARR No. L5C13 Vert/'ca/ to// .02 5« 5 -.02 -.04 I! -20 -/O /O Ang/e of yow , 2" , o/eg (c) r , J ° } cc , -0-2° ; a , 65 /6/j-g ft. F/'gure <3 . - Cont/nued . NACA ARR No. L5C13 Fig. 8d Hor/zorifo/ Vert /'cor/ fa// to// Off Off o Off 4 r 0n- _ 4 V $ <0 -.04 Angr/e of yaw , >" , otep' (d) T , 5° ; ac , /O.J° - y o , 40 /6/sg ft f~/gure <3 • - Conc/ua/ea/ . Fig. 9 NACA ARR No. L5C13 & * & I I I I * * | IS NACA ARR No. L5C13 Fig. 10a v .02 * - S| o /O o -t- Fus&toge. Shori Long ■02 s -30 -20 -/O /Inp/e of yaw ? & , dep /O Co) r, 0° } or, -0-s2° } g, 6 J /6/sg ff. f/pare /O- - £ff&cf of c/7ar?p/r?g fuse/ope /e/?gfh on vor/af/on of yow/ng-mo/nenf , ro////?p-/7?o/7?enf , and /of era/- force coeff/c/enfs w/f/> or?g/e of yow • Hor/zonfo/ ono* vert/ca/ fo//s off • Fig. 10b NACA ARR No. L5C13 * £.02 H ^ "-.02 £° -5 I s Short + Long :^rf Fuse /oge -+- 02 -.02 V * r>l *» «0 $ fl% Q $ "K ^ 1 *. &l <0 k:s V N. <»> * Ang/e of yaw , 7^ , deg (b) r , 0° ; a , 10.3° ; 9 , 40 /fo/sq. ft. F/gure /O . - Conf/nued . NACA ARR No. L5C1S Fig. 10c Ang/e of yaw , If , deg fc) r , 5° ; a , -O.Z.* ; £ , tf\2" /6/vT? /V. ^/gure /O • - Conff/iueo' . Fig. lOd NACA ARR No. L5C13 Ang/e of yaw , IF , deg Cd) r , S°; a , /0.3° ; q; ,40 /b/sg ft. F/gure /O . - Conc/uded • NACA ARR No. L5C13 Fig. 11a Fig. lib NACA ARR No. L5C13 I H \ t >-> £ c -^ 5 o > ^ * 1 s 03 1 i 1 NACA ARR No. L5C13 Fig. 12 VO & 8 2* ^ *> r • 4 S ~o q, tt /A/ 'so ff ■ F/gore /J '•- Fffecf of crporpg/r?*? fuse /ope /erppff? or? vor/of/or? of yow/r?p-r77or?7er?f , /-o///r?^-rr?orr?er?r , a/vcr /o/erof force coeff/c/e/vfs w/fh or?g>/e of yow ■ f/or/zorpfo/ fo// one/ werf/co/ to// 4 on • NACA ARR No. L5C13 Fig. 13b •k ^ , V *J X ?5 //?/j~g_ D/'/iedro/ fdeg) O 5 5 5 .04 .08 Sy/ Sw JZ ■/6 Cb) cc , ZO° . F/gure /4 • - Concluded . Fie. 15a NACA ARR No. L5C13 NACA ARR No. L5C13 Fig. 15b Q> N V ^ C> Q 4 O f* i s -I tf ' C5 w -K »-^ * * ^ ^ T y i Q ^J o °0 I Q Q> ' G" Fig. 16a NACA ARR No. L5C13 -.PO -30 -PO -/O O /O Ar>g/e of yaw , 1r~ ? deg fa) r , 0° } v-}> s t ^4 y V NAT MMMin ONAL Al ~± FOR f IVISORY ERONAUl ICS -30 -20 -/O O /O y4n 9 , 40 /6/sg ff. Fig. 18c NACA ARR No. L5C13 Fuse^/age. l/o^rtica/ toil Shor-h 4- Mad/urn 3 Long /SL -SO -Z0 -/O Ar>g/e of yaw y V 7 a/eg fc) r , 5° ; cc , -O.Z° ; q , 66 ib/jq ft f/Qture iS ■ - Continued . NACA ARR No. L5C13 Fig. 18d .04 V ^ s V $ Q •k $ i s e* 5 s £ ^ ^ >.