ARE No. L5H09 P^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED November 19'4-5 as Advance Restricted Report L5H09 LATKJLEy FULL-SCALE TDHNEL INVESTIGATION OF THE FACTORS AFFECTING THE DIRECTIONAL STABILITI AND TRIM CHARACTERISTICS OF A FIGHTER -TYPE AIRPLANE By Harold H. Sveterg^ E-ugene R. Guryansky, and Roy H. Lange Langley Memorial Aeronautical Laboratory Langley Field, Ya. s. NA<5X WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advalce 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 - 109 DOCUMENTS DEPARTMENT Digitized by tine Internet Arclnive in 2011 witln funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/langleyfullscaleOO NACA ARR No. L5T-109 RESTRICTED NATIONAL ADVISORY COr,ir/:ITTEE FOR AERONAUTICS 1 ADVANCE PES TRIG TED REPORT LANGLEY FULL-SCALE ''lUKNEL INVESTIGATION OF THE FACTORS AFFECTING THE DIRECTIONAL STABILITY AND TRIM CHARACTERISTICS OF A FIGHTER-TYPE AIRPLANE By Harold H. Sweberg, Eugene R. Guryansky and Roy H. Lange Tests were made in the Langley full-scale timnel of the GrumTian XF6f-I!. airplane in order to investigate the factors tiipt affect the directional stability and trim characteristics of a typical fighter-type airplane. Eight representative flight conditions were investigated in detail. The separate contributions of the wing-fuselage combination, the vertical tail, and tlie propeller to the directional stability of the airplane in each condition were determined. Extensive air-flow surveys of sidewash angle and dynamic- pressure ratio along a line coincident with the rudder hinge line were made for each condition investigated to aid in evalxiating the slipstream effects. The data obtained from the air-flow surveys v:ere also used to investigate m.ethods for calculating the contribution of the vertical tall to the airplane directional stability. The results of the tests showed that, for the condi- tions investigated, the directional stability of the air- plane was smallest for the gliding condition vjith flaps retracted and v^as greatest for the v/ave-off condition with flaps deflected ^0° . The variation of sidev/ash angle at the vertical tail with angle of yaw was destabilizing for all conditions investigated. Propeller operation increased the magnitude of the destabilizing sidewash but, at small angles of yaw, also increased the dynamic pressure at the vertical tail sufficiently to make the combined effect stabilizing. The lateral displacement of the slip- stream with respect to the vertical tail at angles of yaw larger than aporoximately ±10*^ caused a reduction in the contribution of the vertical tail to the airplane direc- tional stability at positive angles of yaw -and an increase RESTRICTED NAG A ARR No. L5HO9 at negative angles of yaw. Flaio deflection tended to increase the directional stability of the airplane regardless of the condition of propeller operation. The rudder deflection required for directional trim Vi/as greatest for the wave-off condition wdth the flaps deflected 5*^^. The large changes in the directional trim of the airplane resulting from propeller operation are primarily due to the effects of the slipstream on the wing-fuselage combination and on the vertical tail and are onlj'- secondarily due to the direct effects of the propeller forces. INTRODUCTION The importance of the effects of propeller operation on the directional stability and trim characteristics of an airplane is well knov-n. Past experience has shown that the directional trim is usually critical for a take- off or ].ow-speed climb convjition in which high propeller- thrust and torque coefficients produce large increments of yawxng-moment coefficient. For such conditions, a p?Mot may often find that, because of the large trim^ changes involved, he has insufficient rudder control and is unable to msintein the desired heading. The directional stability is usually lowest for a condition of high angle of attack and low power, during which the contribution of the vertical tall to directional stability is lowest because of the low slipstream velocity and the relatively large loss in dynamic pressure due to the fuselage and canopy wakes. An-s.lyses have been made in the past of wind-tunnel data on directional stability and control (references 1 and 2) but these analyses were based m.ainly on the results of scattered tests of a large numiber of airplanes and airplane models and did not ixiclude any systematic test results showing the effects of propeller operation on the directional stability and control characteristics of a single design. In particular, only m.eager data were available to show the effects of propeller operation on the air flow in the region of the vertical tail. In order to obtain some systematic wind-tunnel-test data relative to these effects, an investigation was conducted in the Langley full-scale tunnel on the Gruimnan XFbF-q. airplane. The investigation included measurements of the NACA ARR No. L5H09 directional sta^ollity and control ch'nracteri sties of the airplane for a \vide rarigs of fl?".^,ht conditions. tor each flight condition investigated, tests were made of the complete airplane, of the airplane without propeller, of the aii'plane without vertical tail, and of the airplane without both propeller and vertical tail. The separate contributions of the propeller, the vertical tail, and the wing-fuselage combination to the airplane directional stability and trim could thus be evaluated. In addition to these force tests, measurements were made of the djnam.ic pressure and the angularity of the air flow at the vertical tail. Particular attention was given to these air-flow measurenients inasmuch as the available data on this subject are vei-y limited. Y ¥. B L S ^L lift coefficient (L/q^S) G^- lateral-force coefficient (Y/qoS) Cj. yawing-moment coefficient (N/qQSb) T^ thrust coefficient (^Tg/2qpD2') Qi^ torque coefficient (^^/Zq^lPj L force along Z-axis; positive vihen acting upward Y force along Y-axis: positive when acting to the right N moment about Z-axis: oositive when it tends to turn nose to right Te effective propeller thrust fx^-, - Xw Xp resultant force along X-axis with propeller operating *tD X' force along X-axis, propeller removed Q propeller torque D propeller diameter (I5.O8 ft) h NAG A ARR No. L5i:^09 S wing area (55^ -Q. -f 't ) S^ vertical-tail area (I9.O sq ft as defined in text) I distance from airplane center of gravity to quarter-chord point of mean vertical-tall chord, measured parallel to fuselage refer- ence line (19.5 ft) b wing span ()4.2.£3 i't ) h^ span of vertical tail surface {k--2.^ ft as defined in text) C4- section chord of vertical tail \i/ angle of yaw, degrees; positive with left wing forward a angle of att'ick of fuselage relerence line relative to free-stream direction, degrees 5|- angle of flap deflection, degrees 5„ angle of rudder deflection, degrees; positive ¥7hen trailing edge of rudder is moved to left (3 propeller blade angle at 0.75 radius or angle of sideslip, degrees '^ sidewash angle, degrees; positive when flow is fromi right to left when airplane is viewed from rear CT average sidewash angle along rudder hinge line weighted for chord and dynamic pressure, degrees I 1 / ^ d'4/ (a tr . — 1 / c^ -^ adb. .' \^^ (qA-o)av St 4 ' ^° V rate of change of average sidewash angle with angle of yaw q local dynamic pressure q free-stream, dynamic pressure MAC A ARR No. L5H09 q/q rstio of local d^mairiic pressure to free-streair dynatrlc pressure ( q/O-o ) average dynairlc-pressure ratio along rudder ^^ hinge line weighted for chord Y. indicated airspeed C„ rate of change of C„ with respect to -^ , per 4^ n degree Cy . rate of change of Cy v/lth respect to 1 1; , per degree /dC. / w rate of change of vertical-tail normal-force do. / /t coefficient with angle of attack, per degree n r degree rate of change of C„ with respect to o-p, per (?\r- rudder deflection at zero yawing-moment coef- ficient, degrees C^Y/r =0 lateral-force coefficient at zero yawing- -"n ^ moment coefficient Subscripts t t vertical tail p propeller s slipstream av average NAG A ARR No. L5HO9 AIRFLA^'S AND APPARATUS Tests were made of the Grumman XF6F-ii, which is a low mldwing single-place fighter airplane weighing about ll,kOO pounds a:id equipped v.l th a Pratt 5- '"Oiitnsy R-2800-27 engine rated at 160O horsepower at SliOO rpm at an altitude of '^'JOO feet. The rear portion of the fuselage is wedge shaped, and the gap between the rudder and fin is sealed. The maximum rudder travel is i55'^» -^ three-view drawing showing the principal dimensions and areas of the airplane is given in fig-jire 1 and photographs of the airplane mounted in the Langlsy full-scale tunnel are given in For some of the tests, the vertical tail was rem.oved and the gap left by its removal was faired to the contour of the fuselage bj a sheet of aluminuiTi. A sketch showing the tail fairing superimposed on the vertical tail surface is given in figure 5, which shows also the principal The air- flow measurements were obtained by means of the com/Dined yaw, pitch, and pltot-statlc tube shoiAOi in detail in figure h. Photographs of this instrument mounted in position for the air-xlow measurements are given in figure 5* METHODS AND TESTS All the tests were made with the airplane landing gear retracted and the cowling flaps closed at a tunnel airspeed of approximately 60 miles per hour, which corre- sponds to a Reynolds n^jmiber of approximately a., 580,000 based on a mesn wing chord of J.SO feet. The ailerons and elevators were locked at 0'-' deflection for all the tests and the landing flaps were locked at ^0° v/hen deflected. Fo attempt was made to daolicate the "blow- up" characteristics of the landing flaps. The directional stability and trim characteristics of the airplane were obt allied for th3 tight representative flight conditions outlined in table I. Directional- stability measurements . - The directional stability characteristics of the airplane, for each flight condition, were Investigated by measuring the forces and moments on the airplane at approximately 5*^ increments NACA ARR No. L5H0g 7 of angle of yaw between ±15'^> which was the maximuin yav/- angle range possible with the present airplane-support setup in the Langley full-scale tunnel. For each of the eight conditions, tests m'ere made of the airplane v.dth the propeller both removed and operating and with the vertical tail surface both removed and in place. Direct io nal-trim measurements . - The directional trim character! s'tics of the airplane were determined from rudder-effectiveness tests. Only four of the conditions listed in table I were investigated; namely, the landing, the Vifave-off, the gliding, and the low-speed climb (V,- = 9S mph) conditions. Rudder-effectiveness tests also were made for similar conditions v^ith the propeller removed. Air— flow measurements . - Surveys of the velocity and angularity!- of the air "flow in the region of the vertical tall V'jere made for all the conditions listed in table I. At each angle of attack, surveys were made for propeller- removed and propeller-operating conditions at angles of yav; of approy.imately 0°, ±5° 3 ±10°, and ±15°- The surveys were made with the vertical tail surface replaced by the tail fairing and consisted of measuremients taken every 6 inches along a line coincident with the rudder hinge, line and extending from approximately li inches above the tail fairing to approximately 12 inches above the top of the vertical tail surface. (See fig. J.) Povjer-on te sts.- For the power-on tests, it was desired to simjulate the variations shov;n in figure 6 of thrust and torque coefficient with lift coefficient for constant-power operation at sea level. It v/as found that these relationships could very nearly be produced with a constant propeller-blade-angle setting of 2J4..0° m.easured at the 0.75 radius; hence this blade-angle setting was used for all the tests with the rrooeller operating. A comparison of the variation of thrust coefficient with torque coefficient for constant-power operation and for the •oropeller vdth a blade-angle setting of 2L.S° m.easured at the 0.75 T'Sdlxis is shown in figure 7* tb^r- the idling- power conditions, the engine was run at the lowest speed, considered possible ( 7OO rpm.) without fouling the engine spark plugs. The thrust and torque coefficients thus obtained for the idling-power conditions were 0.01 and 0.005, respectively. 8 NAG A ARR No. L5H09 ■ Accur ac y of test results .- The accuracy of the results of the force tests is shown by the scatter of the test points. The accuracy of the combined yaw, pitch, and pltot-static tube is estimated to be about tO.25 for the jaw- and pitch-angle measurements and about iO.OlqQ for the G^T^amlc-pressure measurements. Deviations of the test results from zero for apparently symmetrical condi- tions are probably due to differences in the airplane on the tv^o sides of the plane of syirimetry and to asymmetries in the tunne 1 flow . RESULTS AND DISCUSSION The data are given in standard nondimensional- coefficient form v-ith respect to the stability axes and the cent er-of -gravity location shown in figure 1. The stability axes are a system of axes having their origin at the center of gravity and in •;iiiich the Z-axis is in the plane of symm.etry and perpendicular to the relative v/ind, the X-axis is in the plane of symmetry and peroen- dicular to the Z-axis, and the Y-axis is perpendicular to the plane of symmetry. The presentation of the test results and the analysis of the data have been grouped into two m^ain sections. The first section gives results showing the directional stability character! st*. cs of the complete airplane for the various flight conditions investigated and an analysis of the effects of the wing-fuselage combination, the vertical tail, and the propeller on the airplane direc- tional stability. The results of the air-flov; measure- ments in the region of the vertical tail also are included in this section. The second section presents rudder- effectiveness data from which the directional trim char- acteristics of the airplane have been determdned. DIRECTIONAL STABILITY The results of the force tests made to determine the directional stability characteristics of the airplane for each of the eight test conditions listed^in table I are given in figui'e 8. Each part of figure 5 shows curves of Cv-^ and Cy against xj/ for one specific flight NACA ARR No. L5H09 9 attitude for the coiTiplete airplane, for the airplane va th the propeller removed, for tbe airplane with the vertical tail removed, and for the airplane with both the propeller and the vertical tail removed. No test points are shown in figure 8 for the propeller- removed data, inasmuch as these data were obtained from faired curves. Values of C„ and C-^ for the complete airplane in each flight attitude inves-tigated are given in table I. Before a detailed discussion- is -presented of the various factors. that affect the directional stability characteristics of the airplane, a few of the outstanding trends indicated by the test resxilts of figure 8 are listed as follows: (1) The- directional-stability -oaram-eter C- at sm.all angles of yaw (between ±5 ) is sm.allest for the gliding . condition vdth flaps retr-acted.. For this con- dition. Or, - -0.0001==>. (2) The directional-stability param^eter, at sm.all angles of yaw, is largest for the^high-pov/er condition with flaps deflected (wave-off ..conditi.on) . For this condition, Cv, = -0.00l[(.7. (5) For the conditions with high thrust coefficients, the directional stability decreas.es at angles of yaw greater than ap-oroximately 10° and. ..inc-r eases at negative ■angles of yaw greater than- approximately -10 . (Ii) F'lap -deflection tends to -increase the airplane- directional stability. Effects of ""'ing-Fuselage Combination and Vertical Tail v;i th Propeller Removed Wing-fuselage combination .- Values of C^ , a-^d Gy , for the wing-fuselage com/oination" are shown plotted in figure 9 as a function of angle of attack for flaps retracted and flaps deflected S0° . These values of C„ , and Cv , 'vvere obtained from the results shown in fig- ure 8 for the airplane with the propeller and the vertical tail rem:0V6d. The variation of yawing-m:oment coefficient 10 i'ACA ARR No. L5H09 v;ith angle of yaw of the wing-fuselage combination with flaps retracted is unstable for the angle-of-attack range investigated. Incressing the angle of attack, however, decreases the unstable yawing -rr.oment variation of the wing-fuselage combination. A further decrease in the unstable yawing-moment variation occurs with flap deflec- tion and causes the wing-fuselage combination to become stable at angles of attack greater than about 8°. This increase in stability with increasing angle of attack and flap deflection is probably due partly to an increase in directional stability of the wing alone v/i th increasing angle of attack (fig. 8 of reference 3) and partly to an increase in the directional stability csused by a favorable effect of the wing-fuselage interference (figs. [|. and 5 of reference i^) . Thxe variation of lateral-force coefficient with angle of yaw for the wing- fuselage combination is positive for the range of angle of attack and flap deflection investi- gated. Increasing the angle of attack and deflecting the flaps decreases the rate of change of lateral-force coef- ficient with angle of yaw. Air- flow surveys .- The results of the air-flow raeasurem.ents for the propeller -removed conditions are given In figure 10, which shows the variation with height above the fuselage along the rvidder hinge line of the sidewash angle a and the dynarai c-pressure ratio q/^Q for angles of yaw of approjcimately O"-', ±5°> ±10° 5 and tl5°. l^'eighted average values of the sidev/ash angle and dynamic-pressure ratio along the rudder hinge line are given in table II. The surveys ^fig. 10) show that, for this airplane, the variation of average sideY'ash angle at the vertical tail with angle of yaw da/d\i; vjas, in general, positive (destabilizing). The data show that the direction of flow from the fuselage wake and air beside it (region in which share loss in dynamic pressure occurs) is strongly destabilizing. Inasmuch as the vertical-tail chord is largest near the fuselage, the effect of the flow in this region on the contribution of the vertical tail to the airplane directional stability should predominate. The flow above the fuselage wake appears, in most cases, to be slightly destabilizing for negative angles of yaw and to have little effect on the stability at positive angles of yav;. Increasing the angle of attack or deflecting the flaps tends to increase the destabilizing effect of the NACA ARR No. L5HO9 11 sidewash. These results are, in general, contrary to the results published in reference 5* which indicate that the sidewash is usually stabilizing lor lov.'-wing airplanes. The discrepancy nay be due to the fact that, for the present series of tests, the horizontal tail and canopy v:ere in place and the rear portion of the fuselage was wedge shaoed ; whereas the tests of reference 5 v/ere made on a smooth circular fuselage with no horizontal tail. The data given in table III show that the dynainic- pressure ratio at tne vertical tail has its mlnimuiri value at small angles of yaw and increases as the angle of yaw is increased in either direction. For any given angle of yaw, the contribution of the vertical tail to the air- plane directional stability is directly proportional to the dynamic-pressure ratio at ttiat angle of yaw. At small angles of yav.' (between ±5°) the vertical tail lies directly in the path of the fuselage and canopy wakes and hence q/q^ for these oondi ti ons reaches its minimum value.. As the angle of yaw is increased in either direction, the vertical tail moves av;ay from the fuselage and canopy wakes and q/qg increpses. Inasmuch as the fuselage boundary-layer and canopy w-akes increase with- increasing angle cf attack, the loss In q/qQ at the tall increases with increasing angle .of -attack. Vertical tail . - Ej'.perim.ental increments of yawing- moment and lateral-force coefficients due to the vertical tail v;ere obtained from the data of figure 8 for the propeller-removed condition-s and are shown plotted in figures 11 and 12 for all the airplane attitudes investi- gated. Figures 11 end 12 show also increments of yawing- momient and lateral- force coefficients due to the vertical tail that were comiputed on the basis of the results of the air-flow surveys. The force-test data show th'^t the contribution of the vortical tail to the airplane directional stability is lov7er in the yav;-angle range between -5 and 5° than at the^ higher angles of ya-'v and, in addition, the -contri- bution of the vertical tail decreases with increasing angle of attack and flap deflection. Numerical values for the sloTDes C^ ■ and C^^ are eiven in table III. The trends., shewn by tnese results are in agreement with the conclusions drawn from, the results of the air-flow surve vs . 12 NACA ARR No. L5FO9 An analysis has been Trade of the results of the air- flow surveys and. the force tests in order to investigate methods for computing the contribution of the vertical tail to the airplane directional stability. The incre- ments of yawing-moment and lateral-force coefficients due to the vertical tail are given by the following expres- si ons ^''nt = -V-af /, ('" - <'av)(. \°-^/^ d\l/ 3 b (5) and t t '' Equation (5) shows that the contribution of the vertical tail to the airplane directional stability is directly proportional to the derivative of (y - a^^^ f q/q^) with respect to the angle of yav. The term {ii - '^g^-^r) {'i/'io) > which is designated the air-flow factor, is shovm plotted in figure I3 , and average values of the slopes ^{'^ - '^av;U/qo;av between i, = -5" and M; = 5° are given in table III. This table indicates also the effect ill NACA ARR No. L5HO9 on the contribution of the vertical tail to the airplane directional stabil.ity of the decrease in the derivative of the air-flov' factor v^ith angle of attack and flap deflection. For test conditions with flaps deflected 50 the destebilizing effect of the sidewash and the loss in q/'-Iq "^s sufficient to reduce the contribution of the vertical tail to the airnlane directional stability by about ^0 percent of the value that would be obtained /. ^(^-^avll^Aolav j-i — ^ were equal to 1.0. The com- dV narison given in table III of the values of C^ S-J^'^- ^t O^r obtained fror:. the force tests and calculated from equations (3) and (L) by use of the air-flow-survey data and the correction factor of 1.55 ^^^ the jeortietric tail aspect ratio shows fairly good agreement betv;een these slopes . o Effects of Proosller Operation The total increments of yawing-moment and lateral- force coefficients due to oropeller operation are given in figure iL for each of the conditions investigated. These increments v/ere obtained from the experimental data slotted in fisure 8 and are the differences in 0-^ and G-r for the complete airplane with the propeller operating and the propeller rem.oved. For the airplane vjith fhaps retracted (fig. Vu.{a.)), propell&r ooeration wss destabilizing at angles of yaw from about -10^ to 15°; the instability was greatest at Inrge positive angles of yaw. At angles of yavv^ betm-een -10° and -15° > propeller operation gave a stable variation of AC^ against i|/. J^one of the effects of propeller ooeration y^'as prooortionel to the oov;er applied or to the thjr'ust coefficient; in fact, at small angles of yaw (between ^5'^)? the Instability caused by propeller operation was about the same for all conditions, regardless of the thrust coefficient and angle of attack. The effect of propeller operation on the directional stability of the airplane with flaos deflected 50° at sm.all angles of yaw (fig. lij.(b)) ^--as , in general, to increase the st^^bility for the wave-off condition, to decrease the stability for the landing NACA ARR No. L5H09 15 condition slightly, and to cause no appreciatle change in the stability for the landing-approach condition. The average increase in directional stability due to propeller operation for the wave-off condition (rated power, Tq = 0.51), 9t angles of yaw between t5°, was very large ACn^ = -0.00105V \ ^ / Tlie effects of propeller operation on the directional stability characteristics of the airplane can be con- veniently considered under the following groups; (1) Direct effect of the propeller forces on the airplane directional stability (2) Effects of the propeller slipstream on the contribution of the wing-fuselage combination to the airplane directional stability (5) Effects of the propeller slipstream on the contribution of the vertical tail to the airplane directional stability Direct effect of propeller forces .- J/Iethods for computing the direct effect of the propeller forces on the variation of lateral-force and 7/ awing -moment coef- ficient with ancle of yav/ are given in reference 7* The dashed lines shovm in figures I5 and lb are increments of and Cy due to the propeller forces that were calculated by equation (7) of reference 7- (The pro- peller side-force factor was 99'2.) The calculations show that the direct effect of the propeller forces is to decrease the airplane directional stability for all conditions investigated. This effect is greatest for the low-speed climb condition ^Ct = 1«59) T = 0.51"^, for which the decrease in directional stability due' to the isolated propeller is O.OOO5S. Effect of slipstream on wing-fuselage combination . - The effects of the propeller slipstream on the lateral- force and yawing-moment variations with angle of yaw of the wing-fuselage combination m.ay be indirectly obtained from the experimental results. The Increm.ents of C^t^ and Cy due to propeller operation for the airplane with vertical tail removed, increments which were obtained from the force tests, are shown by the solid lines in l6 " MCA ARR No. L5HO9 figures I5 and 16 for each condition investigeted. These increments include the direct effect of the propeller forces and the effects of the passage of the slipstream over the wing-fuselage combination. The difference between the solid and the dashed lines in figiires I5 and I6 are therefore presumed to be due only to the effects of the slipstream on the v/ing-fuselage combination. The data shov/ that for all conditions with the flaps retracted, at angles of yaw between i5"> the slipstream effects on the wing-fuselage combination caused destabi- lizing variations of yawing -moment coefficient with angle of yaw. At the low thrust coefficients this effect was small; at T- = 0«51> however, the slipstream caused a destabilizing increment of C^^ of about O.OOOli?' For P the flaps-deflected conditions, the directional stability of the airplane was not changed appreciably by the slip- stream effects on the wing-fuselage combination for angles of yaw between 5" ^^-id -15 tiut was considerably decreased for angles of yaw between 5 ^^'^ 15°* Effect of slips tream on air flow in region of vertical tail . - The results of the surveys vdth the propeller operating are given in figures 17(a) to 17(e) for the fiaps-retracted conditions and in figures 17(f) to 17(h) for the conditions with flaps deflected ^0'-^ . ?;eighted average values of the sidewash angles and the dynamic- pressure ratios at the vertical tail determined from these surveys are given in table TV. For all conditions investigated, the variation of the average sidewash angle at the vertical tail with angle of yaw v;as generally destabilizing ^positive da^j^^/d-A . The destabilizing effect of the sidewash appeared to increase with thrust coefficient and angle of attack and to decrease with flap clefleotlon. (See table IV.) The most Important factor contributing to the destabilizing effect of the sidewash is the flow from the fuselage boundary layer, which exists in the region in which, for the present airplane, the vertical-tall chord is largest. The destabilizing sidewash in the region of the fuselage boundary layer was smaller in magnitude for the flaps-deflected conditions (figs. 17(f) to 17(h)) than for the flaps-retracted conditions (figs. 17(a) to 17(e)). The data show that the air flow at the vertical tail in the region above the fuselage boundary layer is NACA ARR No. L5H09 I7 dependent on the conditions of propeller oneration. As the thrust coefficient increased frotr: one condition to another, the sidewash in this region became increasingly negative (flov frorr left to right when airplane is viev/ed from the rear). This effect may be accounted for by the slipstream rotation. The vertical tail vas in the region of the rotating flow from, the uoper half of the propeller, which for right-hand propeller operation caused the air to flow from left to right. A further effect of the pro- peller rotation was a lateral d:l splacem.ent (toward the right) of the sliostreami in the region of the vertical tail due to the tangential- velocity components of the rotating flow. The result was that, as the airplane was yaved nose left (negative yaw), the vertical tail tended to move into the center of the sli.pstream and the side- wash became increasingly negative; as the airplane was yawed nose right, however, the vertical tail tended to move away from the center of the slipstream and the side- wash becam.e decreasingly negative. These tendencies indicate that increasing the slipstream, rotation tends to increase the destabilizing effect of the sidewash. The effect of the increased dynaric pressure at the vertical tail due to the propeller slipstream y;as to increase the contribution of the vertical tail to the airplane directional stability, inasm.uch as the average sidewash v/as never large eno^jgh to cause the contribution of the vertical tail to be destabilizing. Surveys (fig. 17) shoved that the disposition of the slipstream at the vertical tail was such that the maximum dynamic pressure occurred at the sections near the m.iddle of the tail for zero angle of yaw and at the sections about one- third of the tail height above the top of the fuselage for other angles of yaw. The dynamic pressure v.-as a minimum at the bottom, of the vertical tail as a result of the large dynamic-pressure losses due to the fuselage and canopy wakes. The displacement of the slipstream with respect to the vertical tail, ?.s the angle of yaw is changed in either direction, can be observed from the dynamic-pressure measurem.ents . The results (fig. 17 and table IV) show that the dynamdc pressure at the vertical tail is highest for negative angles of yaw and is lowest for positive angles of yaw. These results indicate that the contribution of the vertical tall to the directional stability of the airplane with the propeller operating will be greatest at negative angles of yav/. Effect of slipstream on vertical tail .- Sxperim-ental increm.ents of lateral-force and y awing-moment coefficients 18 KACA ARR No. L5HO9 due only to the effects of the oropeller slipstream on the vertical tail surface were obtained from the data of figure o. The increments, which are the difference between the increments of C^ and C^ due to the vertical tail with the propeller operating and with the propeller removed, are shown in figure l3. In general, these results substantiate the conclusions drawn from the air- flow surveys in regard to the effects of the propeller slipstream on the vertical-tail contribution to the air- olane directional stability. The variation of AC^, n+- v/ith angle of yaw is such as to decrease the airplane directional stability at high positive angles of yaw and to increase the directional stability at high negative angles of yaw. Except at T^ = 0.01, at which the effects of the slipstream are small, the directional stability is increased for all conditions in the low- yaw-angle range (between ±5'"') as a result of the slip- stream. This stabilizing effect of the slipstream at small angles of yaw increase? as the thrust coefficient Increases. The total increments of C^^ and Cy due to the vertical tail are given in figures I9 and 20 for the con- ditions with the prooeller operating. These increments were obtained from the data of figure 8 as the differences between the prooeller-operating results with the vertical tail installed on the airplane and with the vertical tail removed. Also shown in figures 19 and 20 are increments of 0^^ and Cy that were calculated from equations (1) ^ t and (2) by use of the air-f lov;-survey data with the pro- peller operating and the effective lift-curve slope of the vertical tail determdned from the data for the propeller-removed conditions. Curves showing the varia- tions of the air-flow factor with angle of yaw for the propeller-operating conditions are given in figure 21. The agreement between the calculated and the experim^ental results shown in figures 19 and 20 is good. Experimental values of the slope CA-y - o^^)(q/q^) av d-V which is used in equations (5) and (i;.) for calculating the contribution of the- vertical tail to the airplane directional stability, are given in table V. These values Bhovj that the effect of the vertical tail in increasing the airplane directional stability is greatest for the conditions with the highest thrust coefficients and decreases as the thrust coefficient decreases. NAG A ARR No. 15 HO 9 I9 Numerical values of C^ and Cv obtained from the "^t ^^t force tests and calculated from equations (5) and (i(.) by use of the air-f low-survey data and the tail lift- ' curve slope previously determ.ined are also given in table V. The satisfactory agreement between the results Indicates that little change in the effective slope of the lift curve of the tail occurs as a result of the propeller slipstream. DIRECTIONAL TRIM The results of the rudder-effectiveness tests are given in figures 22(a) to 22(c) for the airplane with the flaps retracted and the propeller operating to simulate a gliding condition and two low-speed climb conditions and in figures 22(d) and 22(e) for the air- plane -.■vith the flaps deflected 50° and the propeller operating to sim.ulate a landing and a wave-off condi- tion. The results of the tests with the propeller removed are given in figure 25 for the airplane with flaps retracted and with flaps deflected 50°- '^he more important results of the rudder-effectiveness tests are s\imm.arized in figure Zh, which shows curves of dCj-j/d5 j f^r\r, -f.) a^d (^Cy\„ _^ plotted against angle of yaw for each condition investigated. All the values of the slope dCj-^/d5p were measured at zero rudder deflection as a basis for comparison. For the propeller-removed conditions, dC,^ /d5„ reaches its minimum value near zero angle of yaw and increases as the angle of yaw is increased in either direction (fig. 2I4.). The dynsjnic -pressure losses at the vertical tail are greatest at zero yaw, and the losses decrease as the angle of yaw is increased in either direction. For the propeller-operating conditions, the rudder effectiveness Increases as the thrust coef- ficient increases from one particular condition to another because of an increase in the dynam.ic-pressure ratio at the vertical tail (fig. 2I4.). Tor all the con- ditions investigated with the propeller operating, except the gliding condition with flaps retracted, dC]ri/d5p attains its maximum value at high negative angles of yaw and its minimum value at high positive 20 NACA ARR No, L5H09 angles of yav^f (fig. 21).): the dynamic pressure at the vertical tail reaches its maximum value for high negative angles of yaw and reaches its minimum value for high positive angles of yaw. An analysis of the test results showed that the values of dCj-^/d5 are very nearly directly proportional to the dynam.ic -pressure ratio at the vertical tail. The rudder deflections and angles of sideslip required to trim simultaneously the airplane yawing moments and lateral forces for each condition investigated were determined from, the data of figure 2I4. and are given in table VI. For the conditions with the propeller removed, the data show that the values of 5„ and p for zero yaVi?ing-moment coefficient are small. For the conditions with the propeller operating, the data show that the rudder deflections required for directional trim are greatest for the two low-speed high-power conditions. (See table VI.) These deflections, however, are con- siderably lower than the m.aximi:un available rudder travel on the Grumman XF6f-1| airplane. The data show that the amount of rudder deflection required for directional trim In any condition is primarily dependent on the effects of the propeller slipstream on the vertical tail and on the v/ing-fuselage combination and, to a lesser degree, on the direct effect of the propeller forces. The increments of C,^ and Cv at zero yaw due to the effects of the slipstream on the vertical tail, the effects of the slipstream on the wing- fuselage combination, and the direct effect of the pro- peller forces are given in table VII for the wave-off and low-speed climb conditions. Of the total incremient of Cv^ at zero yav-r due to propeller operation for the low-speed-climib condition, 77 percent was due to slipstream effects and 25 percent was due to the effects of the pro- peller forces. For the wave-off condition, 98 percent of the total increm.ent of C^ at zero yaw due to propeller operation was caused by slipstream effects. The curves in figure 2I4. of (5-p)p _p, against ^\! , besides indicating the rudder deflections required to trim the i-irplane yawing m.oments, are a measure of the airplane c?lrectional stability. The conclusions regarding the airplane directional stability character- istics, which are derived from these results, are sub- stantially the same as those derived from, the curves of llkCL ARR No. L5?I09 21 figure 8 showing the variations of Cp against \J/ for 5^ = 0. S U II K A E Y OF R23ULTS Data are presented of measurements T^ade In the Langley full-scale tunnel on the Grurnran jCF^F-U. airplane to investigate the f&ctors affecting the directional stability and trJT, characteristics of a typical fighter- type airplane. Although these datti are quantitative for this particular airplane, the trends are believed to be generally applicable to reasonably similar airplanes. The results are summarized as follows: 1. For the conditions investigated, the value of the directicnal-stabili ty parameter Cv., at angles of yavv- between +5'^ ras sir.allest for the gliding condition v'ith fla-os retracted {C^. = -O.COOIS) and was largest for the wave-ofi condition with flaps deflected 5^'^ (C^., = -Oo00lii7). 'Alth the values raeasured in the low-yav;- angle range used as a reference, the airplane directional stability for the conditions v'ith high thrust coefficients v^'as decreased, at large positive angles of yav/ and v/as increased at large i^^egative angles of yaw. 2. For the XFbF-U airplane, the 'variation of average sidev.'ash angle at the vertical tail -."ith angle of yavi/ was generally sj ch as to decrease the contribution of the vertical tail to the airplane, directional stability. Propeller operation increased the ir.agnitude of the destabilizing effect of the sldev;ash but, at snail angles of yaw, also increased the dynamic pressure at the tail sufficiently to r;ake the cornbined effect stsbilizing. J. rne v'ing-fuselage co'Tibination with flaps retracted, was directionally -■ons table for the ang;le-of -attack range investigated. Increasing the .angle of attack and deflecting the flaps decreased the unstable variation of yawing -moirient coefficient with angle of yav; of the wing- l;.. For all the conditions investigated with the flaps retracted, the contribution of the propeller decreased the directional stability of the air-nlane at small angles of yaw. With the flaps deflected SO'-', at 22 NAG A ARR No. L5HO9 small angles of yaw, the contributj.on of the propeller increased the alrolaiie dj^rectional stability appreciably for the wave-off condition, decreased the airplane directional stability slightly for the landing condi- tion, and caused no appreciable change in the stability for the landing-approach condition. 5. The propeller slipstream increased the contri- bution of the vertical tail to the airplane directional stability at small angles of yav;. As a result of the lateral displacement of the slipstream with respect to the vertical tail, the contribution of the vertical tail to the airplane directional stability was greatest at nega- tive anjjles of yav>' and was srr-Hllest at positive angles of yav'. 6. The destabilizing contribution of the v;ing- fuselage combination to the directional stability of the airplane for the conditions with the flaps retracted, at angles of yav; between ±5'^, v;as increased by the effects of the propeller slipstream. The directional stability of the airplane for the conditions v/i th the flaps deflected 5''^° v-as not changed appreciably by the slip- stream effects on the v.lng-fuselage comibination at angles of yav; between 5° and -1^^ but was considerably decreased at angles of yaw betv^een 5° ^^Q- ^-5°' 7. The amount of rudder deflection required for directional trim, is prlmiarily dependent on the slip- stream effects and only secondarily dependent on the direct effect of the propeller forces. Of the total increment of yav/ing-m.oment coefficient at zero yaw due to propeller operation for the low-speed climib condition, 77 percent was due to slipstream; effects and 25 percent was due to the effects of the propeller forces. For the wave-off condition, 9^ percent of the total increm.ent of yavdng-iroment coefficient at zero yaw due to propeller operation was caused by slipstream effects. The wave-off condition, at a lift coefficient of 1.39* required the largest amount of rudder deflection for trim (5^ = -18.5*^). 8. A comparison of the results of the extensive airflow surveys vvith the results of the force tests made possible the determiination of a value for the effective-lift-curve slope of the veitical ta:l; tlrds value permrltted NAG A ARR No. L5H09 23 calculation of the ccntribut5.on of the vertical tail to the directional stability of the airplane within accept- able limits. Langley Merrorial Aeronautical Laboratory National Advisory Comrrittee for Aeronautics Langley Field, Va. REFERENCES 1. Pass, H. R.: Analysis of 'Aind- Tunnel Data on Direc- tional Stability and Control. NACA TN No. 775, 19i!.0. 2. Irf-lay, Frederick fl. : The Sstimation of the Rate of Change of Yav-ing Moment v/itn Sideslip. NACA TN No. 656, I95S. 5. Shortal, Joseph A.: Effect of Tip Shape and Dihedral on Lateral-Stability Characteristics. NACA Rep. No. 5^3, 1935.' L. Recant, I. G. , and Wallace, Arthur R.: V.'ind-Tunnel Investigation of Effect of Yav; on Lateral-Stability Characteristics. IV - S^/nime trie ally Tapered Wing with a Circular F'uselage Having a Wedge-Shaped Rear end a Vertical Tail. NACA ARR, March 1914.2. 5. Recant, Isidore G. , and -"allace, Arthur E. : '.''ind- Tunnel Investigation of the Effect of Vertical Position of the Wing on the Side Flov in the Region of the Vertical Tail. NACA TN No. 8oli, 19al. 6. Katzoff , S. , and Futterperl, William: The End-Plate Effect of a Horizontal-Tail Surface on a Vertical- Tail Surface. NACA TN No. 797, 191x1. 7. Ribner, Herbert S.; Notes on the Propeller and Sliostream in Relation to Stability. NACA ARR No, Lkll2a, 1914]^. 2h <; cc w o W o CM g w '^ ^ o NA cr. 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K OS 53 ?3 V^ c 5 15 Q qj 0) o CO I •-{ >> (U iH bo a a _] dJ •4-3 C 3 o E 0) C — u •H I Ix. to X C (d E E 3 u i-l 3 bo NACA ARR No. L5H09 Fig. 2b ■d 0) T3 • 3 s rH o o p ; 2X li 8^ NACA ARR No. L5H09 Fig^ \j > t- a => O z 55 o -» 9 Z UJ 9 1^ !5^ o I 1 10 I On NACA ARR No. L5H09 Fig. 5a CO ' 1 (d (U i^ •T3 o •H j (U > t4 3 CQ ITS (U 3 bo NACA ARR No. L5H09 Fig. 5b Q. O E-i •a i-i a c o o in 3 bo Ix, NACA ARR No. L5H09 Fig. 6 I O 'O I ./Or- -n ■ 1 .12- . — — t^A - — uo ,— — • ■ ^ — " — - " D/f i- ■^ _^ -" Uh- — -" ' -- -' _- - rT-- - " n _ ^'^ _ .8 .6 .4 .2 O Rafed power /600 0.6S rofec/ po^ver I040 f^n^ rpm 2400 I960 ~^ " ~^ ~ ~ ~ ' r 7=' ^ ■^ ^ -H •^ ^ ^ ^ ^ ^ — — ■ - ^ - ,> ^ -- — r^ ■^ ^ - ^ ^ '' r-^ ^ -- " r^ NAl lONAL ADVISORY "I — _ L L _l — 1 —1 _l -J — 1 LJ _i —J Ul —J Li — 1 J O .2 .4 .6 .8 /.O 1.2 U Lift coeff/c/enf , C^ /.6 /.a 2.0 22 r/gure 6 . ~ Ca/cu/oted var/af/ons of 7^ anc/ Q^ w/th C^ for con^tonf - power operaT/on at dea /eve/ . 1 NACA ARR No. L5H09 Fig Horsepomr Rated power 1600 0,65 rated power 1040 A -2^.5° measured of 0,73 R Engine rpm Z400 /960 Rated power J .Z .3 ,4 Thrust coefficient^ 7^ Figure 7. - Comparison of mnotion of T^ wfh Q^ for constant- po^er operation and for the propel ier with blade ancjie fixed at 24.Q° Qt 0.75 radius. NACA ARR No. L5H09 Fig. 8a ^ 8 o I -.^ "" ^-^ 1-^ ^ ;^ j^^ ^ — J ir; ^ ::^ , - " r- .^ i^ ai ^ ^ ^ - ^ ■^ 2 " " r»" 3^ ■^ » =^ ^ 3 - _ _ _ v5 ■♦J" c: s:: c r -" — - - Zj-t-^ — — 1-^ zz rq -^ - _, CD- J-" -< — . "-- ^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1 1 1 1 1 1 1 M 1 1 1 -/s -10 IS -s O S lO Angle of yaw y Y, ^^9 (a) Climbinci affifude- oc,/0°; df,0\ Figures.- Directional- stability ctiaracteristics of XF6F-4 airplane; rudder locked at O" when on airplane. NACA ARR No. L5H09 Fig. 8b o I 2 -•' •-1 -.1 ~ --- -€ A in --1 ;-La ^ ;::;^ K — r^ r-; r-- ^ iB u ■ " r* ^ ^ 51 r- _— -- *; S( , ^ -■ --" 'X == -■ f^ — ■ •^ r^ f _ _ •J) o .04- .oz c ^ -.04 Vertical fail on Off ._. ,. On Propelk _ . . Off PrdpGller cff il Power Lift coefficient , C Rated ITc =0.1 1) 0.43 '- RatscllTc=o.iQ .43 Propeller off .4-0 .40 e~ ~ -- -- -- _ ^ —^ a -- -- -- ^ -\ r= — =: =^ — _ ^ — - =3 ■^ f= ;^ 'A s = s \ V ~- ~— ~^ - -« ^_ — — _ . Lt — — — — — — — = ^ — — - = ^ NATIONAL ADVIiUKT -j .. ( :ov ' """''' *MI 1 1 1 — 1 tt run - HI 1 .... J -15 -10 -s s 10 15- Anc^le of yavj j y/ , de ___ J— ^ — — -e J— — ■^ K __ \— ^ — — —* , _^ r- — -C ...1 1 - -IS -JO -5 S 10 Ancjle of yavj , ^ ofey 16 (d) Climbincj affil-ude ; oc,fe.3°; <5f,0\ Figure 6.- Continued. NACA ARR No. L5H09 Fig. 8e X. I I I I I 8 I :^ 2 ./ -I -.2 -^ ^<^% i^S^i -i= =- =^ ^ ^E" ^ " — —-^^^'-'^^ --"^n^^"^ m-'''^'-' f Z^^ TT — > Vertical tai I Power L/ft coeff/c/ent , C^ On Off On Off Tr=0.0/ ^ Tc=0.0/ Prope/hr off Propeller off 0.83 .83 .83 .33 .KJH- " 02 8- ■ - - ^ 3 ^ 5^ j^ ^. ■>- ^ ^ f^ ^ =i t 3 = H }^ >«-j :j Q- — _l 1- -= = = E — — — ■^ i— = — ■*< >S "^ — — -■t) 02 NATIONAL ADVISORY 1 n/t _ CO MH 1 1 tt \o R A 1 tHl NA Ul Cb -15 -10 -5 5 10 Angle of yo\A/ , '^, deg 15 (e) Glidini^ affifude. 00,9.2" ^ 6f^0\ Picture 6.- Continued, NACA ARR No. L5H09 Fig. 8f G^ I I I .2 .1 -.1 -.2 jl T -^ ^■^ '^ '' 2 ^ ^'^ ^''ll-'' -^^^:^''- ' - ^ = g^s2^^ --''','- '' i* ^ " ' '■' "" ^"^ ^7 - ^^^^ a^^^ ^^ Vert/ca/ fa// Poi^er Lift coefficient , C^ I I ^ On 0.65'Rared(i:=0.33) Off .65Rated(^^0.33) On Prope//er off Off Propeiler off .04 .02 -.02 -.04 — 1 ^ > -c s. .. - - _ ^ "" ^ ^ "" ~ .^ J fi _H ™ ^ =* _ _ _ -S L _ — , =1 _ -1 1- = u t "" — — -. ^ D: = ^ — = =3 c - — — - - — — — "" — 6 — — — ^ T ._ ■ 1 .. 1 -/5 -/O -5 5 10 15 Angie of ya^ , "t , cfeg i.37 (f) Landin(j- approach aHifude^ cc,S.6°; 6^,50" Figure 3.- Continued. NACA ARR No. L5H09 Fig. 8g I -/ -z 3 -^■^ J ^"r" -i i-" ^p:-"^-'" ^^ ,^^ zT^- ""^ ^-^^^ ' _. — -r^'' ^-^ "-''^ ^e^ ^^-^ ^ -^ ^^^ T^ v5 o -In. i Vertical tail On Off- On Off Power Lift coefficient^Cj_ Rated Oh = 0.5l) Rated (Tc = 0.^/) Propeller off PropeZ/er off 04 NATIONAL ADVISORY COMMITTEE FM AERCHIAUTICS ■" 0/ ^ ^ "^ - ~_ „ , ~ - -- - -- _ _ . , _ a. ■ — - -. .,^ -1 "" a ■^ -J P -' b'^ i 1— _i- — — — — i - ^ ■P^ ■ _-i L- -^ "" 1 ^ -J Si- — ■i -«— ■ — 1 — — -^ .04 _^ ^_ __ u _ u _ _ -IS -10 -S 10 IS An(^le of ya^ ^ y, de-■— ^s3P- ""^ - ' '^^ ^^T"' 1 --Z -t^^^^ i^^ i ± Vertical tail Po/^er Lift coeff/c/'ent , C^ ^ On 7c = 0.0/ Off Tc = O. Of On Prope//er off Off PropellGK off /.S8 /.6Q /.J6 /.66 ^H- 02 — — ^ --- — - - — , _ — {_ ^ ^ &- — 1 =t F= — ■^ =; -r= Z^ — 9,- — — t T~ ' — ^ ^ K ^ ::0 02 NATIONAL ADVISORY COMMITTEE FOd AERONAUTICS D/t . II Ml -/5 -10 -5 5/0 Ang/e of ya\^ ^ >^^ deg '5 (h) Landin(j afii^ucJe ; oc,//.<3°j Sf ^SO" Ficjure 8 .- Concluded. NACA ARR No. L5H09 Fig. 9 O ^ .01 ^ '.01 /a /6 ^. Flaps retractecf 'Ffaps def/ecTed 60 ,00/ ^c -00/ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ,F/ape retracted ■\ Flaps deflected 50' 4 6/2/6 Ang/e of attack , oc ^ cteg Figure 9,- Coninbut/on of wmq-fu^elaqe combination to Cn cind Cy , l/erf/cal fail and propeller'^remov^c/, ^ n NACA ARR No. L5H09 Fig. 10a US Iff _lL-j r iiQ , ■"-^•~ .,.-^ra Cj ■§r- % « ^ . ^j &i^ -'i r_j E^S- NACA ARR No. L5H09 Fig. 10b iiiilLlLU J: ;y-ferrttil: 5h »i>4v= #@r^" I — -=2^77273^:. s 4i- ."X:^; 1 -■ '-\ I NACA ARR No. L5HC9 Fig. 10c ,43 1 ^ S" .. I ^:i ;si^-t:-:SL bt.J ji ^a-^u: a ? NACA ARR No. L5H09 Fig. lOd n 1 1 I ' I > i I I M ! I I I I r m m ■H-- JlIMI I I K'tAM \'i I' I I l.n .IT r r 'l rm I ' ■ ;^.-'~ .1 >i *, >-: -^ J T^'?'''" :rrw -vdf-jpofp TSL!l^rM^^3M:Z i?:7iaa^ -si(6uu jy^cjjSJff -&-!^-- > o c rwr" ppr.j/o-j- /fbj: ' /Doirjs^^oaiTvqtxxjSsu ift: kks!iMi!M^ khvkmXli.iMi i,:J::.i .S&i::. t. tli i.J .ri:.1 t.ri,.. I..: L ::i . i:.! J :t..x i ..t, t.,:::J,:.J. L,. l. .i3j J.:.Lu .,i.± J.L.faE NACA ARR No. L5H09 Fig. lOe t "^ ..3:1: f 4- .£r. vts:. 'ws' -©- ' — - e--& &:.■ -_ -— -1 - -T- -., l^f— -- ^ — ^ rs': <YDZvy"'Sc^/:j'/'iq/ ifi±H3ii±rn:fe±: iiiiteiiHiiit:; NACA ARR No. L5H09 Fig. lOg m FFFFF 1 ij-r- ■i! k' riy^rf'r t: W1_^"__ --^--;=^ '3-TlSiP ■"-ji?*^"'' erH& — %- ._Os»S.. '::Xs::.'::^- m W 'J I^M -pcfijro^ //^ '^^ ^^a^' ... . ^J^ i 1 , ■- ^ s' 1 "1 J __ - ?--'- :z;3:::^ u ■+"^ it t-l'=H^ ff tnro NACA ARR No. L5H09 Fig. 11a force - fes f do fa Ca/cu/a^eof from surveys iSeg) C^ /P.3 /.04 -.02 -15 -10 -5 5 Ang/e of ya/v , '^, c/eg /O /5 .83 .60 .40 .23 (a) Sf, 0°. Fiqure II ■ - Comparison of /ncremenf of C^ due to verficol toll as deferm'med from the force- tesf data and from the ojr-f/ov^ surveys. Propeller removed- 1 NACA ARR No. L5H09 Fig. lib d^ F'orce-^esf do to Co/cu/aTecf from surveys I I c I .^a -.02 a. (deg) Ct_ 11.8 1.56 =^ =», 1 ■=3 r ~ — ~ = ^ = ..\fi — — . ~i = _J i^C " " ' -^ ~ ~ ~^ ^ ^ .^ — ~- =^ rr= ~_ - 4.."^ ■ - ~ ::r- :::<; ~- -- 02 - L _ U-l 1— < _ , _^ 1 -/5 -10 -5 5/0 /5 Ang/e of yaw , 1^ , cfeg (b) Sf , S0\ Figure II .- Concluded . NACA ARR No. L5M09 Fig. 12a Force -f est data Calculated from surveys ^t A> -1 >^ V ./ *^i i>. ^ di ,^1 ^ •k -I .S .1 V, .0 Vi s^ -1 ^ .1 e n V, < (^ ~^ v -./ tO) -k J c p> L ^v: oc (deg) C^ 12.3 1.04 9.2 .83 -15 8.9 .80 8.4 .40 ~^ "■ ~ " _ _ — ~ - , , __ ^ — — ■ NATIONAL ADVISORY ~| . CO MM TT tb KX ) A IR( N« UT CS .J 10 .23 -10 -5 O 5 10 /J Angle of yoi^ , H^ ^ deg (a) Sf ,0° . Fiqure 12..- Compari.5on of incremeni of Cy due to vertical fail as determined from file force -test doio and from the air-flow surveys- Propel le r re mo ve d . 1 NACA ARR No. L5H09 Fig. 12b Force -test data Ca/cu/ofed from surveys I I I .(J 8 I I I o ./ 7/ ./ o J -J -/5 — — — -— »_u — _-3 — — — - _- _= =„ ^ ^ ~^ — "" ^ _ _ -_ — :- — _ ; . = — — -^ — — " - - ^ — — 1 1 NATIONAL ADVISORY j _ _ U- CO _ 1H L 1 1 tc w _ .. J (deg) C^ 11.8 /.J<5 5.8 /.// 4.9 L04 -10 -5 6/0 Ang/e o5 ya\^ , 1^ , deg /5 (b) Sf-.SO" Figure 12.,- Concluded. NACA ARR No. L5H09 Fig. 13 15 10 6 ^ -5 I I 16 -10 lO -15 5 -5 -10 cc -15 7 Z I y IZ. z. '/ ^z V A ^I ~ T- n^^ y^ .2 Flaps def/ecfed 50' T /. 1^1 -K' y T- — & — r T^^- k I 8 I (deg) /.O 3.4 6.9 3.2 /2.3 ^ ^>i »vV V h!P Is (^1 "b -V ^-; V ^ !<■ V 'i) ^ 1 ^ <.^ ^) ^ ^ ■K ,^ .H) Al S § ■^ s> . t* \ /^orce-fesf data Cq Icuhfed ( Prope//er a/one) oc .deg) Q 7^ I £3 1.39 0.61 92 .S3 .01 8.9 .96 .30 3.4 .43 .11 1.0 .24 .06 r^ATlONAL ADVISORY COMMITTEE FM AEDOMAUTICS' I I M I I I I I Z -15 -10 -S O S 10 IS Angle of yaw , ^ , deq (a) cSf,0°. Figure /s:- Exper/menta/ and calculated effects of propeller operat/on on the variation of Cy with T for the a/rpiane with vert/cat ta/t removed. NACA ARR No. L5H09 Fig. 15b Force -fesf data Calculai-ed ( Propel /er a /one) I .(J I I I I "■ " // __ ^-^ _ _ r- ^ = „ _ _ = =: S; L ^ .« - = — - - — - — — — - — — ./ -.1 oc Cdeg) Ci_ 7^ ii.Q 1.58 aoi S.8 1.37 J3 ./ " ^ " ^ 4 r— — — - 1 - — - ^ ^ — 1 '' J — — -./ .-• ^ ^ 1 , 1 _ -A _ 1 1 4.9 139 ^1 -/S -/O -SO s /o An^/e of ya w, f , dea (b) 6,, 50\ Figure /s: - Concluded. /S NACA ARR No. L5H09 Fig. 16a I o c: D^ ■DZ .01 -.02 .02 -.01 .02 -.02 .02 -.02 Force -fesf data Co/culQi-ed ( Prope//er a/one) oc (deg) Ci_ ^ 12.3 1.39 0.51 9.2 .83 .01 '3.9 .96 30 '3.4 .43 .11 ID NATIONAL ADVISOR r COMMITTEE FN AERONAUTICS J I — 1 i 1 1 \ I I 1 u .24 .05 -15 -10 5 5/0 /S Angle of yaw , Yi deg (a) d^,0°. Figure /6. - Exper/menfa/ and co/cu/otec/ effect's of propeller operof/on on fhe var/af/on of C^ with) y* for the airp/one w/th vertico/ Ta/7 removed. NACA ARR No. L5H09 Fig. 16b I .o o -.02 OZ Force -fesf da fa Calculoi-ed ( Propel /er abne) oc (deg) Q ^ -.02 .02( 11.8 I.5Q 0.01 5.Q 1.37 J3 'Z.g 1.39 .51 ^ NATIONAL ADVISORY COMMITTEE FW AERONAUTICS J— I 1 1 1 — I — > — r 1 1 ^ -15 -10 -5 5/0 Angle of yaw , '^ , deg (b) df,50°. F/gure i(>. - Conc/uded . 15 NACA ARR No. L5H09 Fig. 17a y- -tS ■nzi-.. W ■ Q ^:---^±Li-:xf ■m^- m NACA ARR No. L5H09 Fig. 17b NACA ARR No. L5H09 Fig. 17c p iDij'fTrj'P qt.^^^ iJ^fis*^— +--- ^ — r^ M^ NACA ARR No. L5H09 Fig. 17(1 1 « . .- --, i-| ^ j£^:,::;iT^x^|T-t:t^-^^-x '' -^ - - , -r, ■^" ^'^ ^I ^ "It ±: .X -:i ""±13 '""^4:54:"" '■!■-' 'r~i \ ' , ! ! 1 ' 1^ i ' I 1 1 ^ f. - . r--4_ ; JT T 4jr.+-.. +^± V ' 1 "■ ^ 1 ' TT i H ! IJ I ^ : ^ ^ K . ''Mr *? 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T : : 1 \ ' 1 ' ^— '1 ' 1 ! : ' i , ! , ^ ' - i " ; » \ 1 ' ■ i >' 1 ' ' J ^ ,^ ^ ■ \l, i_;s-,, ! l X ^ r .....,j-^^ L^x5xspx f V -^X- "t ■ i ■- -*+-" '*- 1 i fT"^ r T 1 1 ^ 1 X^ ^ 7 Tjr - I ' i 'j ' i ■ J ' I •■ \ ' ' [ ^ ' \ ' L I ' , J- , . , ,j , . , . r . , 3 .,XX; j [-stjRr^-7*^^<;- -4--'^ "^7"^ " "^ ' ' ' ■ • '^ ' i ; i x'' ' X ■ ' ' '^ ' ' 1 i ^' "L^j-X t xdlinS^ ' X '^X-i.j^'V ilXX .X •xu,4-iX^^Xi(i+ii-4^i4sy-i-isryit "^T; -TT-;- -, 1 |-;- --ir- -X- -j--^.- .-pi-X-^.-,--;- - - 4 ^ t _ L.]_4- + - xiIiiE-i-x^i: "is-; xr:i^: +tL^±J-X^=i:^?S:S^^^4:^±::| 1 1 TT - y 1 \ r\\ (3 ! '^ i 1 1 Q 1 1 ' ' 'T ' \ 1 . ^ . '"X-Vu^p-i'D -- 1 -\-pi ■f-' :-■-- ■- ■■;■- •-+'■ -^- 1 -i - r- 1 r "7"X'" r " ■ 'X ' ""' "^ ;lXrf^ ill." :':.-': i ■■ i 1 ' ' : ! ^l i '•"(Xcx:. - "" ______ 1 ;:■;.. :i;y^vMS5A/ .; x::Xh.-1=X'^"xr-r ''p^o 1 ■ ^ / vfJ 1 s,; VI ^ <;;:; ^>' . ^ ^ ; ■ i ih,Mi1^^^-^fK,P' xtigt:- iptT-V'^^^^^tTi^T^tT^'" .:. , 1 L ' X 1 .....: S ,..■ _.l L.''!?-y/»< <^-?'jy'y^:^'i ■'(:;'..(_ ^ .lI .1 ^- ^jm,- /opfija/t-^. , 1 .,:... , ■ ^ i !■ 1 : (_. i ; , ' , ..-;.] :.! ' j'/yBiw • _..i..i i • f- -.«: 1 10 b« ^ "^ 5 ^ .5 o ^ ^ v: -,'? ■!- X -10 -16 " ^ - ^ ^ \ * / \ y f' y V / -- / / ^ f / / / __ — / / ^ ' 4 / ' / / f f y ^ / / <• y I / f / / / / / f / / / / / / / / / '-' , / / / / / / / / / / / / / / ^■' ^ 1 / /■ / 1 t ^ ( 1 A / 1 / 1 / ^ / / / / I / / t y / ^ y 1, ' / / / / / / / / /, / / ^ f / ''-y V / / -y / / NATIONAL ADVISORY COMMITTEE FOfi AERONAUTICS - - __ 1 1 1 1 1 1 1 1 1 ' (deg) -8.9 -3.4 -1.0 -3.2 -20 -15 -10 -5 5/0 /5 Angle of yayv , "^^ deg (a) df, 0°. Figure 2!- Var/ar/on of a/r-f/o\// factor with angie of yaw . Prope//er operat/hg . NACA ARR No. L5H09 Fig. 21b gc^_de^ Power Uft coeffic/ent. C^ — 4.9 RaTed (-^=0.5/) /.33 --S.6 0.65 Rated(rc-0.33) 1.67 11.8 Tc- O.O/ /.58 25 20 On ^ /5 ^^ 10 ^ o -10 5 - -15 -20 -15 XL : : IT -' :;= (C ^^ "^ ^ / ^ ==^~- / y r 7 J / r ^ — L ^ T -fi^ / 7 / / )_ ^ /_ 7 ^ / ■ ■ \— *- ' _^ £_ ^ 1 — L z. - ' ^ ^^ ' ^ 'Z ^^^ _ T~ -p^ 4.'^ ^ t^ 4A t- J^ t /^ -i - Ti' -/■d tt ^^^ % -t 7 '^ J I COMMITTEE FOft AEW)NAUTICS ± 1 1 1 1 1 1 1 J 1 1 1 (deg) 4.9 5.8 //.S -D -5 O 5 10 Ang/e of ya^/v^ T , deg (b) df , 50°. Figure 21 .- Conc/udec/ . 15 NACA ARR No. L5H09 Fig. 22a X ^ v5 -30 -20 .2 OZ ^ -.04 -/O O /O £0 30 y (deg) Angle of yaw, Y, deq -14.6 -9.9 -SI o so 10.0 14.1 J i fe^ M k V (deg) r^ ^ F^ C^ b 1 "^ H fe ^ k ^ ^-/4.6 \^-9.9 X ^ S ^ ^ ^ -SI ^ ^^^^ — 5.0 "-—10.0 NATIONAL ADVISORY COMMITTEE FM AMOMAUTICS -—14.7 -30 -20 -/o 20 30 o /o Rudder def/ecf/on , Sr ? deg {a) 6lid/ng cond/t/on ; 1^,0.0/ ; Oc ,9.2°; Figure ^^ . - Variot/on or C^ and Cy with d^ for Qeveral angles of yaw. Prope//er operating. NACA ARR No. L5H09 Fig. 22b O I \1 Angle of vaw,Y,c/eQ ^ — -/4.6 ^ -9.9 -30 -ZO -10 O 10 ZO 30 Ru dder def/echon 7 <^ ^ de^ (b) Climbing cond/rion ; rated power (1^=0. 30) lOc ,6.Q°; C^,0.96 ; df,0° Figure 2,Z . - Contin ued . NACA ARR No. L5H09 Fig. 22c V. I \1 .01 % r -.01 -.04 r06 Anqle of (jaw^ y, decj NATIONdU ADVISORY COMMITTEE FOB MRON»UTICS -30 -20 -lO O /O 20 30 /Judder def/ecfion t Sr 7 c/eo (c) Climbing condition ; rated power (Tq'^O.S/); Figure 2.2 .- Continued. NACA ARR No. L5H09 Fig. 22d ^ I ^^ Ancjie of yaw J Y> deq -9.9 -6.1 5.0 10.0 ^ -30 -20 -/O O /O ZO 30 Ru c/der def/eof/on , Sr , deg id) Wa ve -off cone// f ion : raTea poi^/er (7^=0.^/)) Figure ^2 .- Continued . NACA ARR No. L5H09 Fig. 22e S ^ ^ N ^^ Anc^le of yaw, Ijf^ deq -14.6 -9.9 -J.I 5.0 10.0 14.7 ^ -30 -20 -lO O /O 10 30 /Rudder def/ec f/on ; c5r ; deg ie) Landing condition ; 7^,0.0i jccj /.8°; C^^ i.58; Figure 2Z.-Conciucie(d. NACA ARR No. L5H09 Fig. 23a ^^ x-2 I o -./ I I .OA Ol -.01 -.04 Angle of yaw , \/, deq -14.6 - 9.9 -SI SO 10.0 14.1 '30 -20 -/O J 10 20 R.U dder def/e c f/on ^ 6r 7 dec^ Figure 23. - Var/af/on of Cr, and Cy with d^ for severaf angies of yo^ . Propeller removed . NACA ARR No. L5H09 Fig. 23b I ^ Anqle of yaw, Y, de(^ -I4.(> -9J> -SI S.0 10.0 14.1 -JO -20 -/O O 10 20 JO /Judder def/ecf/on , 4- j de(^ ib) OC , Id.O "'; C^ , /.08; 6f, O^. F/'gure ^3.-Cont/nued NACA ARR No. L5H09 Fig. 23c o^ 5 -5; v.^ Ani^/e or yaw , Y, d^^ -14.6 -9.9 -6.1 6.0 10.0 14.7 ^ -30 -20 '10 lO 20 30 /? a dder def/e cf/on 7 ^n de(^ ic) a.se"; c^,/-09 ; 4;, ^o"". Figure 33 .- Continued . NACA ARR No. L5H09 Fig. 23d Anc^le of yaw , Y, decj -14.6 -9.9 -S.I SO 10.0 14.1 -30 ^ -20 -10 O lO 20 Rudder def/ecf/on , Sr y de^ id) a, 1 1.6°; C^, /.66; df ,^0°. 30 Figure Id . - Concluded . NACA ARR No. L5H09 Fig. 24a O ./ CK' ,9.2 ; C^,0.83 ;, propeller removed 0C,9.2°;C^,O.83 ; 7^,O.OI 20 10 ^^ o iO -20 ~ ~ " " "n ~ ^ ~~ N ^ ^ ^ ^ ^ -J ^ ^ ^ --- ^^^ ^ - -^ ^ \ \ nI \ \ _ _ _ ., \ -.002 :§^ -.00k o — — - - _ J .- _- - — - ~ IT — = = \ ^^^ ...-r.,. MM 1 1 tt 1 -IJ 15 -10 -5 O 5 10 Angle of yaw , T , deg (a) Gliding condition ; df,0. Figure ^4.- D/recTlonoi tr/m characteristics of the XF6F-4 airplane . NACA ARR No. L5H09 Fig. 24b ./ ? O ^ >^ -./ -2 20\ 10 -10 -20 — :*-' ' — --,^^1^ __ • oc J 9.2 ; Q , 0.83 ; propef/er remo ved ■Oc,&.3°jC^, 0. 96 ; ratecf pOi^er(l^=0. 30) — ■ ~ ~ \ V ^ ^ ^ ^ ■" ^ •s ■" ^ . -^ ^ - ^ ■~- ■^ '■ -- - -- - — -. ^ ^ ^ 1 — 1 K -.002 ^^VS -.001 -^1^ ^ ~ _ _ _. — - - — •- — ~ -^ - -- . " ~ ~ " - NATIONAL ADVISORY _ CO HM Tl Et K) ! A tuu N< HI Cbl o -15 -10 -5 O 5 10 Ang/e of yaw , ^ , deg (b) C/imb/ng cond/T/on ; df ,0 . Figure ^'f■ .-Continued . /5 NACA ARR No. L5H09 Fip. 24c o i^ ^ cN ^y i< -/ -.2 10 ^ o -10 -20 -30 — ::* --— — — —^ '— —, .^i— — , — ^ --^ — 0CJ3.0 ; C^,/.06 ; prope/fer remove(::f 0Cjl2.3°; C^,/.SQ; rated power (l^-=0.5/) o II ~ -^ -^ \ N N -V ' ^ s -^ s -~~ ^ > ^ ^ -■ ^ ^ >, ~ — ■ -\ — _ -V ~ -_ N — 1 V ^ \ , -.005. -.00^ ' " - __ _ — _« " ^ — J - "^ ■^ i ^' J - - ^ — — — — 1 1 1 — NATIONAL ADVKORY _ L u ^ u , _ _ _ _J CO ' " MM M tt »U « A itH JN< U> ii.:i J -15 -10 -5 O 5 Ang/e of yoi/v ^ i/r, deg ic) c/imbing condition ;6f,0°. Figure 24.- Continued. lO 15 NACA ARR No. L5H09 Fig. 24d 2 J ^ O c -.2 2a ' ' / / ^ / __ -" __ — '- __. ' 1 " __^ U- --1 . 1 . ^ ^ ^ ' ^ ^ _^ ^ ' _ _ _ ^ _ _ ^ /(9 o II -zo -30 (X , S.6° ; Ci_ , 1,09 jprope/ler removed / ■ ■ .. .-. .. ^, ■^ ^ \ ^^ ■^^ "■ — — _ ^ ^ ^^^ ^ "^^^ 'X- ^-^^ ^^ -^ \^ s 5 ■^- \ ^ ) -.003 -nnc ^ -nn, - _ „ -1 _ — - - , ^ ^ 1 - ~ -- — — -_ ^ — — \— ■ — — ' — — — ■" " - c NATIONAL ADVISORY ] _ u u u u L _ _ ^ « J -/5 -10 -5 5/0 Ang/e of yaw , 'f' , deg id] Wave -off condition ; cff, '50° Figure 24. - Continued . /d NACA ARR No. L5H09 Fig. 24e ./ — 1 ^ o II ^ ' '' — ^O^ _ T^ ^ ' -* -~ ^ — ■^ ^ ^ '' __ -^ ,^ -^ -1 - L_ iD /O ii -10 -20 -30 — (X,//.8 ; C^,/.66; props/ /er remo\/ed — oc,//.8°; C^, /.^6; 7^,0.0/ s 1 N s s ^ "v, \ ' ^ v ^ .^ ^ ^ — — ^ -J ^ ^ ~ ~ ^ ^ --^ ^ "s s \ N \ \ \ N \ \ N ^ . S -DDP .L/\yc ^f \^ -.00/ >s _ ~ - _ , _ - — _ . _ , "" ' — H ^ r - ^ ^ ^_ _ _ ^ _ _ _ LJ _J [_l I—I 1 1 1— 1 LJ 1 1 1 1 L_ -/5 -10 -5 O 5 10 Angle of yaw , y^ , deg {e)Land/ng cond/t/on ; d^^ 50 .' Figure £4. -Concluded. /6 UNIVERSITY OF FLORIDA !iiiiiiMiiiiiiiiiiiiii{iiiiiir 3 1262 08104 965 1 UNIVERSITY OF FLORIDA DOCUMENTS DER^RTMENT 120 MAR3T0N SCIENCE LIBRARY RO. BOX 117011 GAINESVILLE, FL 32611-7011 USA ^ 1 'I