K/Afa-L')?- ^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED May 1924.14. as Advance Reetricted Report Ll|.E25 INVESTICJATION OF EFFECT OF SIDESLIP ON LATERAL STABILITY CHARACTERISTICS I - CIRCULAR FUSELAGE WITH VARIATIONS IN VERTICAL-TAIL y^SEA. AND TAIL LENGTH WITH AND WITHOUT HORIZONTAL TAIL SURFACE By Leo F. FehLner and Robert MacLachlan Langley Memorial Aeronautical Laboratory Lan^lcy Field, Va. NACA !h 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. DOCUMENTS DEPARTMENT L-12 Digitized by tlie Internet Arcliive 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/investigationofeOOIa ■^kCk ARR ITo. L4E25 RESTRICTED ^^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE RESTRICTED REPORT IIWESTIGATION OF EFFECT OF SIDESLIP ON LATERAL STABILITY CHARACTERISTICS I - CIRCULAR FUSELAGE WITH VARIATIONS IN VERTICAL-TAIL AREA AND TAIL LENGTH WITH AND WITHOUT HORIZONTAL TAIL SUTvFACE Ey LeQ-:F. Fehlner and Robert MacLachlan SUI.1MARY The results of tests of a circular fuselage with various combinations of tail lengths and vertical tail surfaces with and without the horizontal tail surface in the 6- by 6-foot test section of the NACA stability tunnel are reported in the form of diagrams of vari- ation of coefficients of lateral force and yav/ing moment with angle of yaw and angle of attack. The results of these tests indicated that the change in the unstable yawing mor.ient of the fuselage alone due to increased tail length did not appreciably affect the yawing moment of a fuselage and vertical- tail combination. The addition of a horizontal tail increased the efficiency of the vertical tail in normal-flight attitudes and in the region of negative angles of attack. Existing methods of computing tail effectiveness gave results v/ithin ±7 percent of the measured values for the cases computed. INTRODUCTION Desirable qualities for the lateral stability and control characteristics of an airplane are dependent on the set of stability derivatives peculiar to the airplane. The stability derivatives can be changed by changes in airplane parameters, such as vertical- tail area, horizontal-tail area, and tail length. RESTRICTED NAGA ARR ^lo . L4a25 Extensive tests to determine the changes in stability derivatives effected by uniform changes in airplane parameters have been made with a model geometrically similar to the model used in the present investigation. Included in these tests were the effects of cowlings, of wing positions, and of the presence of a vertical tail (references 1 and 2). Reference 1 is mainly con- cerned with lift and drag characteristics, whereas reference 2 deals with the effects of yaw on the lateral stability characteristics of a rectangular wing with a circular fuselage and vertical tail. The present investigation, in which the model was tested without wings, is an attempt to determine experimentally the basic changes in stability deriva- tives caused by uniform changes in vertical-tall area and tail length and by the presence of a horizontal tail. Because a geometrically similar model has been tested in the LMAL 7- by 10-foot tunnel (reference 3), the data may be used for correlating the results in the two wind tunnels. The tests v;ere made in the NAG A stability tunnel and included an angle -of -attack range from -10*^ to 20° and an angle -of -yaw range from 12*^ to -30° with various combinations of three fuselages of different lengths and five vertical-tail areas with and without a horizontal tail surface. Two com.binations of the model parts used in the present tests are geometrically similar to the model used in the LMAL 7- by 10-foot tunnel for the tests of references 2 and 3. APPARATUS AND MODEL The tests were made in the IIaCA stability tunnel 6- by 6-foot closed-throat test section with the regular six-compcnent balance. The principal dimensions and the arrangement of the parts of the model used in the investigation are shown in figure 1. All the model parts are made of laminated mahogany. Figure 2 shows the model unassembled, and figure 3 shows the model mounted on the model support. The horizontal strut supporting the model does not rotate in pitch with the model. The vertical struts rotate in yav/ with the model and remain alined with the relative wind. NACA ARR No. L4525 Th3 fus3lag3 is of circular cross section. Its length can ba changed by tb'i use of three in-csr changeable tail cones. 'Vhjo the ?/n.orr.O'it cf these tail cones is attached, the fuselage is p;8cnietricall7 similar to the circular fuselage described In rel'srence 1 and to the model used for tre tests reported in reference 2. The coordinates for the mediuir. and long tail cones were obtained by extending the abscissa for the length of the short tail cone according to the formula X-L = X^ + Xo(c - 1) sin ^ X^ where Xq abscissa for original length a length of portion to be distorted c ratio of original length of portion to be distorted to final length of portion distorted Xn abscissa for final length of distorted portion The ordinate s corresponding to X]_ are taken as those corresponding to Xq from which X]_ was computed. The tail lengths, the lengths of the three fuselages and tail cones, and the ratios of the tail lengths to the 48-inch span of the proposed wing are given in table 1. Five geometrically similar vertical tail surfaces were made to conform to the NACA 0009 section. In plan form they are representative of the vertical tail surfaces used on the average airplane. The geometric aspect ratio of each vertical tail is 2.15. The horizontal tail surface was made to conform to the KACA 0009 section. Its geometric aspect ratio is 3.99. The niombers by which the tail surfaces are designated, their areas, and the ratios of these areas to a proposed rectangular wing area of 561 square inches are given in table 2. TESTS The model combinations tested are given in table 3. Angle-of -attack tests for each model combination were made over a range from -10° to 20° at angles of yaw NAG A ARR No. L4E25 of -5^, 0°, and 5°. Angle-of-yaw tests for each model combination were made over a range from 12- to -30° at angles of attack of -10°, 0°, 10°, and 20°. The dynamic pressure for the tests was 65 pounds per square foot, which corresponds to a velocity of about 160 miles per hour. The Reynolds number based on an S-inch wing chord was about 888,000. RESULTS The results are presented as standard nondimensional coefflclaats based on the dimensions of a rectangular wing proposed for the model. The following symibols are used herein and the senses are defined relative to a person vi/ithin the airplane facing the direction of motion: Cy lateral-force coefficient (Y/qS^) C^ yawing-mioment coefficient (N./qS^b) Y lateral force (positive to right) N yawing moment (positive when right wing tip tends to miove rearward) q dynamic pressure p air density M V tunnel-air velocity _ d Cy % = J47 ^1/ angle of yaw, degrees (positive when right wing tip has miOved rearward) a angle of attack, degrees (positive when tail has been depressed) I tail length b wing span (48 in.) S^ vertical-tail area NACA ARR No. L4E25 S_ horizontal-tail area S^ wing area (361 sq in.) Af aspect ratio of vertical tail surface Figure 4 shows the system of axes used in the measurement of forces, moments, and angles. The axes are fixed in the model for all changes in angle of yaw. for changes in angle of attack, the X-axis remains in the plane in which it was located at a = C^ . The axes intersect the model at the assuriied center of gravity, which is 10.40 inches behind the nose. The lateral- stability derivatives are computed, for the range of angle of attack, from measurements of lateral force and yawing moment at angles of yaw of is^; the variation of the forces and moments with angle of yaw is assumed to be linear over the ±5^ range of angle of yaw. Angle-of -yav; tests were made to check the linearity of the curves of Cy and C^ against angle of yaw in the 15^ angle-of-yaw range. The slope of these curves shows that the variation of the forces and moments within the angle-of-yaw range of ±5° is linear except at high angles of attack. The measured slopes of these curves are plotted with tailed symbols in the figures. The measurements of lateral-force coefficient Cy are considered accurate to ±0.0012 and of yawing-moment coefficient Cn to ±0.0005. The angle-of-yaw measure- ments are accurate to about 0.06°, and the angle of attack is accurate to about 0.1'^. A model geom.etrlcally similar to the NACA stability tunnel model was tested in the LMAL 7- by 10-foot tiinnel and the results of the tests were reported in refer- ence 3. The model consisted of the short fuselage, vertical tail surface 4, and the horizontal tail surface and was tested in the LMAL 7- by 10-foot tunnel at a dynamic pressure of 16.37 pounds per square foot, which corresponds to a velocity of 80 miles per hour. The Reynolds number based on a 10-inch chord was 619,000 and the turbulence factor was 1.6. The model tested in the NaCA stability tunnel is eight-tenths the size of the model tested in the LMaL 7- by 10-foot tunnel and was 6 NAG A ARR No. L4E25 tested at a dynaTiic pressure of 65 pounds per square foot corresponding to f- velccity of 160 miles per hour and a Reynolds numbar of 658,000, The turbulence of tr.e air stream in the NACA stability tunnsl Is not known tut is believed to be low^^r than that of the LMAL '7- by 10-fjct tunnel. Variation of Cy,, and On,, v/ith a for tne similar models are shown in figure 5. Values of Cv ^i"-d Cn agree well for the two sets of data. The maximum discrepancies occur at high angles of attack in the region of the stall. In order to check the data obtained in the NAGA stability tunnel, a tem.porary one -component spring balance was installed to measure the yav/ing moment due to sideslip. The model support consisted of a cylindrical rod fixed perpendicular to the top of the tunnel by a tripodal wire stay. The model was supported in the same position in the tunnel as on the regular tunnel balance except that it was inverted. Such an arrangement was expected to give altogether different interference regular support. Figure 6 sho".;s the with ii thus obtained for a typical of attack of 0^ and compared with the model en the regular support in the NACA stability tur^nel and in the LIvlAL 7- by IC-fcot tunnel. The two sets of data obtained in the WaCA sta- bility t\innel check each ether. The data obtained from the LLiAL 7- by 10-foot tunnel check the slope from the NaCa stability tunnel within 8 percent. This difference in slope is the same as the difference shown in fig- ure 5 for Cy. at an angle of attack rf 0°. The source of the discrepancy is not obvious from the data tut m.ay be the differences in the deflection of the models, angularities of the air stream, or model-size to jet- size ratios. The results are presented in the form of curves that show the effects of changes in fuselage length for fuselage alone in figures 7 and 8; of changes in vertical-tail area, figures 9 and 10; of changes in tail length, figures 11 to 13; of adding the horizontal tail surface, figures 14 to 16; and of changes in tail length and vertical-tail area with constant tail volume, figures 17 and 18. The data plotted in the various figures and the model combinations used in obtaining each plot ai-e summarized in table 4. effects f r; :ro the variation of Cn case at an an gle KACA ARR No. L4E25 " DISCUS3ICN Effect of Changes in Fuselage Length for Fuselage Alone Th9 effect of changes in fuselage length on the stability of the fuselage alone is shown in figures 7 and 6. The derivative Cy v/as increased by an increment of 0.0007 by increasing the value of Z./b from 0.418 to 0.618. The absolute magnitude of this increment is small compared with the magnitude of derivatives for the fuselage with the vertical tail surfaces tested. The derivative Cn is very slightly if changed by the change in fuselage length relative to the magnitude of the derivatives for the complete model. Although theoretical analysis indicates that the unstable yawing moment of the fuselage alone varies with fineness ratio and volume, this variation has not been detected herein because the magnitude of the variation is of the same order as that of the experimental accuracy. Effect of Changes in Vertical-Tail Area The effect of changes in vertical-tail area (hori- zontal tail on) is shown in figures 9 and 10. At an angle of yaw of -10° and at an angle of attack of 0°, changing Sf/Sw from a value of 0.0659 to 0.0974 increased Cy hy an increment of 0.019 and C^ hy an increment of 0.0097. Throughout the angle -of -attack range, the same change in Sf/Sw increased Cy hy an increment of about 0.0014 and Cn by an increment of about 0.00098 V The values of Cv and Cn decrease v/ith angle of attack; the decrease for a change in angle of attack from -5° to 5° is 0.00043 for Cn , with vertical tail J/ surface 2 and 0.00048 with vertical tall surface 4. The decrease in Cv . for the same decrease in angle of attack and for either vertical-tail area is 0.0012. ru:Ck ARR No. L4E25 Effect of Chanres in Tail Length Ths effect of changes in tail length is shown in figures 11 to 12 for the model with the horizontal tail surface and vertical tail surface 4. The change in Cy due to changing l/b from 0.418 to 0.618 is small and probahly negligible for cases in which the lateral force is largely the contribution of the vertical tail surface The effect on Cy as shown in figure 12 a-pears - if inconsistent but is small and therefore probably negligible . The yawing moment due to sideslip increases with tail length. This increase in C-p^ Increases with ^V up to shortly after the stall of the vertical tail sur- face. At values of Uf beyond the stall, Cn is increased about 0.01 by an increase in tail length of 0.1. Changing the value of l/b from 0.418 to 0.518 causes an increase in C^^ of approximately 0.0007 throughout the angle-of-attack range. An increase in l/b, however, from 0.518 to 0.618 causes increases of 0.00059 and 0.00046 in Cni, at angles of attack of -5° and 5'^, respectively. Increasing the angle of attack decreases C^ • For the short, medium, and long tail lengths, the decrease in C^ is 0.0001b, ■J; 0.00037, and 0.00050, respectively, for an increase in angle of attack from -5° to 5^. The effect on On,, and Cy. of changing the vertical-tail area and tail length is shown in figure Ic . The m.odel, in this case, is at an angle of attack of 2° and is equipped with the horizontal tail surface. Increases in vertical-tall area produce regular increases in both Cy and Cn, . Increases in tail length produce regular Inci^eases in Cn excent for the extremely small values of Sf/S^ for which the directional instability is of the same order of magnitude as for the fuselage alone. i'or all practical values of Sf-Zs^,,, therefore, increasing FACA i'.RR No. L4E25 tail length increases th3 dirscticnal atability as measured bT Cn , • Effect of Horizontal Tail Surface The removal of the horizontal tail surface decreases the efficiency of the vertical tail surface in all attitudes except at large angles of attack. (See figs. 14 to 16.) For the long fuselage, at -5° angle of attack, the value of Cy, -s decreased by an increment of 0.001 by removing the horizontal tail surface v/hereas, at 5° angle cf attack, Cy, is not decreased. The effect on Cy,, in rreneral, is the same magnitude for the short-fuselage and vertical-tail combination. For the long fuselage, figure 15 sbov/s a large effect on 0^ that varies from a decrease of 0,00090 at an angle of attack of -5° to a decrease of 0.0C039 at an angle of attack of 5^. The corresponding decreases for the short-fuselage and vertical-tail com.bination are 0.00054 and 0.00020. (See fig. 16.) By removing the horizontal tail surface, the efficiency of the vertical tail surface is therefore decreased by an amount that varies with angle of attack. The decrease is slightly greater for the short-fuselage and vertical-tail combination than for the long-fuselage and vertical-tail combination. Effect of Changes with Constant Tail Volume The effect of changes in tail length and vertical- tail area with tail volume held constant based on the dimensions given for the m^odel is shown in figures 17 and 18. The derivative Cv increases as the vertical- tail area increases and as the tail length decreases. The value of the derivative Cn , theoretically should not vary Vi?ith changes in tail length and vertical-tail area if the tail volume is held constant. The variation of Cn,, , measured experimentally, is sm.all over the range of angle of attack from 4° to 12° but is appreciable at negative and at high positive angles of attack. 10 NACA ARR No. L4E25 Comparison with Theoretical Variations The experimdntal results have been compared with theory by m.eans of accepted simple calculations that Involve a m.inimum of anticipation for the desired results The first of these calculations can be made from the expression of the variation of lateral force with side- slip, which can be written as Vfuselage V Vf where { Cy ) i^ the experimental value obtained \ '"^/fuselage in this inves tia;atlon. N f = 2tt s^ A^ + 2k (2) and f denotes vertical tail surface. The constant K is given in reference 4 as 0.875 for an elliptical span- wise loading. If the spanwise loading of the vertical tail surface is assumed to be elliptical for the purposes of analytical treatment and if the model dimensions are used as previously given, values of Cy for vertical 'J/ tail surfaces 2 and 4 are 0.00248 and 0.00515, respec- tively, according to equation (2). The corresponding experimental values com.puted from the data according to equation (1) at 0'^ angle of attack and with a horizontal tail on the long fuselage are 0.0038 and 0.0054. The theoretical relation then underestimates the value of Cy by 9 percent for vertical tail surface 2 and 5 percent for vertical tail surface 4. Similarly, Cj^ may be written (s\ = - ^^^, Theoretical values of Cni for vertical tail surfaces 2 and 4 on the long fuselage are -0.00215 and -0.00317, respectively. The corresponding experim.ental values for the model with the horizontal tail surface are -0.00226 and -0.00350. The theoretical relation then under- estim.ates the value of Cv, for combinations with the NACA ARR No. L4E25 11 long fuselage by 5 percent for vertical tail surface 2 and 4 percent for vertical tail surface 4. The theoretical values of C^i for vertical tail surface 4 in combination with the medium and short fuselages are -0.00267 and -0.00215, respectively. The corresponding values of -0.00277 and -0.00207 were obtained experimentally for the model with the horizontal tail surface. The theoretical relation then underestimates 0^,1 ^'or the medium tail length and vertical tail surface 4 by 4 percent ana overestim^ates C^ for the short tail leng-th and vertical tail surface 4 by 4 oercent. If. the value of Cv comouted according to H' equation (2) is increased by 2 percent, the resulting values of Cv, and Cn, ar^ within 7 oercent for the cases analyzed. This 2-perc3nt increase in Cy. may account for the influence of the horizontal tail surface, the influence of the fuselage, or any peculiarities of flow. The resulting discrepancies, which amount to ±7 percent, are slightly less than twice the limits of discrepancy shown previously between the data from the NACA stability tunnel and the LIv:.-i.L 7- by 10-foot tionnel. CONCLUSIONS The results of tests in the NACA stability tunnel of a circular fuselage with various combinations of tail lengths and vertical tail surfaces with and without the horizontal tail surfaces indicated the following conclusions : 1. The effect of the change in the \Hastable yawing m.oment of the fuselage alone due to increased tail length on the variation of yawing moment with sideslip was negligible. 2. At an angle of attack of 2^, the vertical tail surface in the presence of the horizontal tail surface produced values of lateral-force derivative Cy,. and yawing-moment derivative Cn,^ that were within 7 per- cent of the estimated values. 12 LTACA ARR No. L4E25 3. The addition of the horizontal tail surface increased the efficiency of the vertical tail surface. The increase in Cv, varied frora C.OOl at an anpile of attack of -5"^ to at an angle of attack of 5^; and the increase in C^ varied from 0.OC090 at an angle of attack of -5° to 0.00039 at an angle of attack of 5^. Langley Memorial Aeronautical Laboratory National Advisory Cornnittee for Aeronautics Langley Field, Va. REFERHTCES 1. Jacobs, Eastman N., and Ward, Kenneth E.: Inter- ference of Vising and Fuselage from. Tests of 209 Combinationr in the rl.A.G.A. Variable-Density Tunnel. NAG A Rep. No. 540, 1935. 2. Bamber, Ivi. J., and House, Ro 0.: V/ind-Tunnel Investigation of Effect of Yaw on Lateral-Stability Characteristics. II - Rectangular N.A.C.A. 23012 Viiing with a Circular Fuselage and a Fin. NACA TN No. 730, 1939. 3. Wallace, Arthur R., and Turner, Thomas R.: Wind- Tunnel Investigation of Effect of Yav; on Lateral- Stability Characteristics. V - Symmetrically Tapered Wing with a Circular Fuselage Having a Horizontal and a Vertical Tail. NACA ARR No. 3F23, 1943. 4. Higgins, George J.: The Prediction of Airfoil Characteristics. NACA Rep. No. 312, 1G29. WAG A ARE No. L4E25 13 TABLE 1 FUSELAGE DIIvTENSIONS Fuselage FuselCLf-e length (ill.) Tp.il-cone length (in.) Tail length, V (in.) (1) Tail length 7 Wing span * Short c? • OD 20. C7 0.413 Medivun 37.05 14.65 24. e? .518 Long 41.85 13.45 29.67 .618 "Tail length measured from center of gravity, assiiir;ed to be 10.40 in. behind nose of the model, to hinge line of tail surfaces. NATIONAL ADVISORY GO'lJITTEi FOR AERONAUTICS TABLE 2 AREAS OF VERTICAL AND HORIZONTAL TAIL SURPaCES 1 Taiirsurface DsGignaticn Tail area (sq in.) Tail area '.'/ing area Vertical Do Do Do Do Horizontal 1 2 3 4 5 10.83 23.78 28.37 35.15 46.20 64.21 0.0300 .0653 .0786 .0974 .1278 .178 NACA ARR No. L4^25 14 X Eh o (H CO w Eh o n r-( <_ M CQ O o o o ■ 1 Oj (D O rH O •H .H ^ Cm -P ;d Cm iH Cvi to si^ lO Cm rH W fj ■ '-0 r-i of 1 1 1 1 1 1 1 1 1 1 +J 05 1 1 1 1 1 1 1 1 1 1 C iH O 1 1 1 1 1 1 1 1 1 1 O .H 03 a c o O c O O o o O d W Ci tH (D Ti '■d 'd Ti ■d 'd 'd -d nd ■d V1"^ 1 1 1 1 1 1 1 1 1 1 1 1 r { O CO 1 1 1 1 1 1 1 1 1 1 ra L_. t t , , , , , 1 , bO g 1 1 1 1 1 1 1 1 1 o:< p 1 1 1 1 1 1 1 1 1 rH •H O o o o bO o o b O d xi 'd ■o •'Z.J -d r-r 'd •d ■d 'd X! n I"-' 1 1 1 1 o 1 1 t 1 1 1=^ 1 1 1 1 .-) 1 1 1 1 1 5h 1 ' 1 ' ' 1 1 1 1 C* o •r-t -P d r' 0.1 ■o ^ lO c£j i> ■X) a> o rH (M •H rH 1 — 1 rH H fH rH rH r-\ w CM CM -S n o a r-i d CI O rH O '-0 r-l '.S 1 1 1 1 1 1 1 1 1 -P f) 1 1 1 1 1 1 1 1 1 c; H o 1 1 1 1 1 1 1 1 1 O "H Kj <♦-( o O o C o o o o o o !SJ 05 Ch Ci-i ■o Ti 'd O -d ■'d ■d Xi ■d Tl •,H -P in o 1 1 1 1 1 1 1 1 1 ^ ^ 1 1 1 1 1 1 1 1 1 O »-0 1 1 1 1 1 1 1 1 1 1*H (D fcO jz; crt -p 1 1 -p 1 1 1 1 1 3 rH f-. 1 ■:0 1 f-, 1 1 1 1 1 •rl (D o o C o o o o o o o Tl CO f'' TJ O ^d .G '^d ■d TJ ■d 'd .P. :^ 0-) 1 ,4 1 W 1 1 1 1 fin I 1 • 1 1 1 1 c o •r-! -P cd iH Oi « ^ lO -.D £> CO a> o rH i'^ rH rH •M .Q B O O o tH ;>H til' O tr. rO O •H « > W Q < PC ►J o < fc o f-" M '■-- Fh Fh < IH O o :...CA ARR No. L4E25 15 TrtELS 4 FR^SENTrtTIOi: OF RlSULTS Plot Model con-blnation Figure Cy and C^ against a 9 ac; Cv, against ^1/ 9 ^6 Cy and Cr^ against ^ 1 and 3 7 Cv and Cv, against a 1 and 5 8 Cy and Cn against \i/ 19 and 21 9 Cy and Cn against a 19 and 21 10 Cy and C-r^ against \!/ 9,15, and 21 11 Cy, and C^ , arainst a Cy, and C^ , against -£ at a = 2*^ 9,15, and 21 12 5 to 22 13 Cv and Cn against t' 4 and 21 14 Cv, and C„ against a 4 and 21 15 Cv and C- against a 2 and 9 16 Cy and C^ against \!/ 9,14, and 1^ 17 Cy and C^ asrainst a 9,14, and 19 — ■ 1 16 ^Also shov^'n in this figure are results from Llvl^iL 7- ty 10-foot tunnel for modal with di-nensions geometrically sirailar to ir.odel combination 9 tested in NACA stability tunnel. TT •.T' lOI^-AL AD YI 30RY C0M'ITT''3j; for AERONAUTICS NACA ARR No. L4E25 Fig. 1 .Z80RU6.7ZO NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. .603-' \.468 R .053 R .^ \'6.09&' 1 1.^60 R [*-7.4l6'H I 1467-' 0.7II"R-- H4.l^r' 0.991 '-' Fi<^ure I.- The circular fuselage, /ert/cal and horizontal taili, and tail conei with the pnncipol dimenjionj for o^^embly. NACA ARR No. L4E25 Fig. I/) > t- a 3 i2 o > '^ l< -J 9 Z UJ 2 hi ii o o O E o I 0) 3 bo NACA ARR No. L4E25 Fig. 3 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. Figure 3.- A/loo/el mounted on model support. NACA ARR No. L4E25 Fig. 4 > c V E 0) • o en w c E o 4J c o c •N •~" ^0 -o -V- 4J c 2 5G . s^ ^ o 1 ^" u k_ D O^ Fig. NACA ARR No. L4E25 u > t- i5 o > 2 < z o o ^ ^ <;> i t ^ « ^ "*s,^ c: ^ ^^ Si Is: 1 t -^ 1 1 1 >^ ^ ^v o < ^ ^ ^ •• a C )J « O vo ^•H © (J-P O « d ■C3 >«-! C-H ^ d h « n rHTj © r-t C >^^ M C -P etf 3 ^ 4J 43 r-i • •H rH _d- >»^ a) T=- •P 0« 4J OS •H 4J C (-t 0) O • o-> •H M 43 *H -H (U cd O (h II 03 13 ■p o « « K «-i| » -\ 60 < B CO |(Z) b a o o i% -=h "N 0) JC 6 "m -p « O 043 ■f- fi«- Gd *3 V( cd . iH Q) C «^ f-« Q~> O C -P CO S 9 6 C > 000 ^e OVi • m^ 1 1 03 1 o-o rA • H a -4- ir\ aJ • ^ e Xi u > II S)cJ-cy '^\Xi 00 I <\1 I NACA ARR No. L4E25 Fig. 6 bi ,02. o o c Q) I >5 .01 ".01 -.02. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. \ \^ \ V 1 3—^'^ k ^ k \ k ^ \ 1 k ^ o Stability funnel rec^ular balar ^ 5 fabilitv tunnel sonna balant ice \ L \ ^ 7- -by h 0-foc >ttun nd bi aianci » '30 -2.0 "10 Ar\Qle of yaw , ^ , deo 10 Plgxire 6.- Comparison of data obtained with NACA stability tunnel regular balance, NACA stability tunnel spring balance, and LMAL 7- ^7 10-foot tunnel balance for rate of change of yawlng- moment coefficient with angle of yaw. Horizontal tall surface on; - = O.I1.I8; vertical tall surface I4., b 1^ = 0.097i^; a '^W = 0^ Fig. NACA ARR No. L4E25 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. -30 -zo Angle of yav¥, \^ , deg ZO Figure 7.- Effect of changing fuselage length on rate of change of lateral-force and yawlng-moraent coefficients with angle of yaw. No horizontal or vertical tall surfaces; a = 0°. NACA ARR No. L4E25 Fig. 8 I ir ijE , 1 J JT^ 4 i H 1 13^ t 1 -r f -ir , ^ \ \ L. i —\ 7 s s 00 00 K \ ^ o 4- N \ uJ i ^ ( 7 ^ Q • > 1- ^J a O 10 3 ■« Z O > a < < _i O < z M> o UJ K- o < X 2 X «> o to o c • U C0 <\J ■p «-t -p o d +> o r^ o cri cU Q) o C UJ "D ^2 -p si -» W*3 C -H • « O -^ O -^c* >»- -icr 60 C «H o C5 c n w "d 3 C r-l 4~ f>-i Qi -^ O 01 W -P *+- OOP > o i^ 4-1 «H > O 1 8.- Effec lllty dorl orlzontal 1 gure 3tab No h T &4 ^ "0 ^A- Fig. 9 NACA ARR No. L4E25 -30 -2.0 -10 10 Ana le of yaw ^ ^ j deo Figure 9.- Effect of changing vertical-tall area on rate of change of lateral -force and yawing -moment coefficients with angle of yaw. Horizontal tall surface on; ^ = O.618; a = 0°. b NACA ARR No. L4E25 Fig. 10 O fhu o o I" o ovl VO u >- K a => O z i5 o > '^ 3< -J ? Z UJ 9 !^ I- ►- II o <>J 00 "^3 S3 O I I T §» o u c« O 4J « * o • O bO (3 O 4 -P H » I C 00 u n II o CO bO C « a) o x; > ■P (m 01 o > ■P u c « t3 «-l H-P 1 rH • w^ o^ f-t « ■p c n u c o f> o (0 p ■P C o M O U) O tl] &4 :? ^A. Fig. 11 NACA ARR No. L4E25 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. -2L0 "10 An^/e of yaw , ^ dei Figure 11.- Effect of changing tail length on rate of change of lateral-force and yawing-moment coefficients with angle of yaw. Horizontal tail surface on; verti- cal tail surface !+» o" = O.O97I4.; a = 0°. f NACA ARR No. L4E25 Fig. 12 vo ON o u • o o t> ^ II too C 0) > (3 60 H^ ^ ^ O o ^c^ o <0 ctf • o ■****. O 0) 01 CJ-> a> u O-H 7i •< 4-> ■p a a 1 o > r-t - •N 4- (J I I 4) O -J NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. -10 10 f\r,(^\Q. of yaw ) (p ^ deq 20 Figure II4..- Effect of horizontal tail surface on rate of change of lateral-force and yawlng-moment coefficients with angle of yaw. — = O.618; vertical tall surface k, b -i- = O.O97I4.; a = 0°. Fig. 15 NACA ARR No. L4E?5 C) ih ^ r V© «\J I 00 I o « do C • SX o U ai ■p «-i -P o d *j o at u o c to O C o oA T^ o 0) cd at -p c o ^ ■r^ O U O A « o v^ > ■p ■p at o > © •H V I'd • >>. • -P ON o 01 |cO u. o o at -P at o -P (0 > CO SO -H O a> ^ K»|43 D d A :p NACA ARR No. L4E25 Fig. 16 Q) O I* o CM ^ -ci- iH c~- C 60 C7N o c O cd • o o o J3 0) -P II ^^ c •* •HO -^ cd -p o n o -H C at cd at H •HO rH (h •H o at J3 0) ■p 0) «-i > rH O-H Of ^ o -P 0) «H o > ■P a> "H tl w-d • •^ 1 >> 00 • -p rH •sOn-i _=^ rHrH • •H o (D^ ;h 3) II §)w r^\Xi ^ ^ "0 u. Fig. 17 NACA ARR No. L4E25 Figure 17.- Effect of different combinations having constant tall volume on rate of change of lateral- force and yawlng-moment coefficients with angle of yaw. Horizontal tall surface on; a = 0°. NACA ARR No. L4E25 Fig. 18 fh I ^ Q ^ vO tall vol with ang ^1 ■p § t -P o a C o -o cr>> u C cd QJ 60 00 t> C > ^ ^O 8 09 n -^ C o o > '«*- JC ■P -p c o aj aj O c > <5 •r\ -rK O ■t- ■2 ^ " a « 0) -K o "d «H C ^ V 4J -P 01 C -H o «) r-l rH ti -^ -H d) V ^ 0) ^ -P »H 01 rH ^ ■O al