N(\C(^L'l1' ^ AER No. L5AI3 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLy ISSUED February 19'4-5 as Advance Restricted Report L5AI3 THE IKFLUENCE OF VERTICAL -TAIL DESIGN AND DIRECTION OF PROPELLER ROTATION ON TRIM CHARACTERISTICS OF A TWIN -ENGINE -AIRPLANE MODEL WITH ONE ENGINE INOPERATIVE By Marvin Pitkin, John W. Draper, and Charles V. Bennett Langley Memorial Aeronautical Laboratory Langley Field, Va, 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- ** » ' I*' 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 - 191 , 1 Digitized by tine 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/influenceofvertiOOIang VT NACA ARR Ko. LfA13 NATIONAL ADVI30IIY C0?>5?^ITTEE FOR AERONAUTICS 3XOm79^{ ADVANCE RESTRICTED REPORT THE Ii??LUINCS G? VERTICAL-TA^L DESIGN AND DlR^CTTO:^ 0? PROIELLLR RCTATTO:! CN TRIM CHARACTIRI3TICS OF A TWIN-ENGIKE- AIRPLANE MOD'sL WITF ONE ENGINE INOPERAIiyZ By Mai'vin 13 tkiii, Joljn '/■;. Draper, and Charles V. Bennett 3NMMARY Tefrts have been hiade in the Langley free-fllfrht tunnel to determine the influence of rriode of propeller rotation and vertical-tail design upon the trim characteristic? of a model of a tvvin-engirie airplane vtith one engine inoper- ative. The test model was mounted on a trim stand, v/hich allowed freedom in roll and yaw under conditions sim^ulat- ing thore required by the NAGA and the Army Air Force? for asymmetric-pcwer operation in flif'^ht. The seven vertical- tail desiavis tested included three tails of low aspect ratio and of different area, one tm'in tail of low aspect ratio, two tails of high aspect ratio and with different rudder areas, and one all-mo vaole tail of high aspect ratio equipped -.vith a linked tab. All tests vvere made v/ith the flaps dov^n. The tests showed chat the effect of m:od(. of propeller rotation upon tne directional trim characteristics of the model operating with a='yrTmetric power was considerable. Propeller rotation in v.'hich che upper tips rotate cut- board toward the vfing tip (outboard rotation) generally created m.ore -severe out-of-trim ccnditicns than ir^board rot action. The all-movable tJiil design was found to be more effective than the other designs tested in nullifying the effects of aryrrjnetric pover. The conventional tail de- signs with high aspect ratio were mors effective than the designs with low aspect ratio in this respect. The single vertical tails were generally more effective in tr'mming the yawing FiOments created by asy.:itrietric power than twin vertical tails of the same asoect ratio and eoual KAOA ARR Ko. L5A13- area, par-txcul3rl7 when the rudder was free. At p'Tr.all anf;le3 of sideslip, however, the mo?Tients caured b^* a?yirmetric pov/er Vi/ere more readily tr.'.r'ned by deflecting the rudders of the twin tail'^ than by deflecting the rudder of a single tail. The trlrrmlng effectiveness of the vertical tail in- creased al:noet directly V;'lth vertical-tail area but in- creased at a decreasin£^ rate with rudder deflection and chord. When the rudder was free, the addition of dorsal- and vsntral-fin areas permitted increases in the asym- rnetric power balanced by the tail surface at moderate angles of sideslip. INTRODUCTION of one or more engines of xnultieiigine airplanes introduce? a sudden and severe demand upon the directional stability o.nd control of those airplanes. Such failures result In the instantaneous application of large yawing mornents that must be reutralizied either oir the rudder control or by the directional stability of the airplane. In addition, asyranetric power conditions create rc'lling n:3ments that :nust be balanced by aileron deflection in order to inaintaln straight flight. This aileron de- flection creates additional yawing moments that r'^cuire further trimming by the vertical tail surfaces. For multi- engine airplanes, then, the asymmetric power condition generally im.poses the most severe requlrem.ents for di- rectional stabilit:/ and control and to a large extent dictates the design of the vertical tall surfaces of these airplanes . An invef:^tigation has therefore been carried out in the Langley free-flight tunnel to provide data concerning the relative merits of seven vertical-tail designs and two modes of propeller rotation mider conditions of as:/m- metric power. ■ The NACA and Army fl-ying-qualitles require- ments (references 1 and 2) for directional stability and control of airplanes operating with asymm;etrlc power were u'^fed to establish the test conditions. The results of the invcstigati:!n are reported hrrein. NACA ARR No. L5A13 A J^- scale model of a conventional twin-engine air- 20 plane in the r:ed.ixini-bomber class (the North American B-2S airplane) was used in the tests. The model was mounted on a test stand, which allowed freedom, in yavif and roll. The effects of asymmetric power could thus be visu- ally observed from changes in the model attitude. The seven vertical-tail designs studied in this in- vestii^ation varied in either aspect ratio, total tail area, rudder area, or general arrangement. Tests vvere made with rudders fixed and free, and the effects of adding various dorsal and ventral fins v;ere studied with the rudders free. The eff-^ct of mode of rotation of the operating propeller upon the vertical-tail characteristics was investigated for all tail arrangements. All tests were made with th^ fla-os down. SYMBOLS C T lift coefficient flijll] .-, . _ . ^r.. . . . Rollin \ ^ -w / / \ -, T -, . , ^r.. . u f Rollins moment \ C7 rolling-moment coefficient I ttc^, '' ~ /■ _ ^ q-w"w /' Cn yawing -mom.ent coefficient (' ^'sving moment \ Cv^ rate of change of yav/ing-mom.ent coefficient -P with angle of sideslip I - - — D propeller diameter, feet p Gensit;7 of air, slug per c^.'^hic foot V free-stream airspeed, feet per second Vj. , ,.„ vel:)CitY at end of take-off run, feet t,er take-oif ' , ' second V -^ stalling speed with flacs dovm, feet per second m_n _ -c .. ^ X q free-stream, dynamic pressure, pounds per square foot (¥"') w 4 IkACk ARR Fo. L5A13 T^ effective thrust of one engine, pound r^ T) propeller efficiency, percent f-;ros3 v;eifTht, pounds 3y_, v'ing ai-ea, square feet bhp brake horsepower of full-scale airplane simulated by rrodel thp thrust horsepower 63, corrblr.ed aileron cfeflectlon, do.crrees 6p I'udder deflection, positive when trailing edge is to left, degrees 5|. flap deflection, positive vvhen trailing edge is do^A/n, degrees 5 elevator deflect:^ on, degrees e ^ 5m tab deflection of all-movable tail, positive when trailing edge is to left, degrees 1, tall incidence of all-T.ovable tall with respect to center line of fuselage, degrees a angle of attack, degrees a+. local angle of attack of x'-ertica] tail, degi'ees p angle of siaeslip, degrees ^ 2/ \ A asoect ratio of v^'rtical tail (t-^ /^^ V / / St area of vertical tail, square feet S-j-, balance area o'^ rudder, percent ruddci- area S rudder area, square feet bf Epan rf vertical tail, feet b^^v wing span, feet NACA ARR No. L5A13 APPARATUS Wind 'Tannel The tests were r.ade in the Langley free-flight tunnel, a complete description of vhich will he found in refer- ence 5. The tunnel was locked at an angle of pitch of 0° for all tests. Trim Stand All tests were made on a trim stand, v/hich was securely festezied to the floor of the wind tunnel. This stand was so constructed as to allow the model freedom in roll and yaw about the stability axes of the m.odel. The stability axes are a system of axes in which the Z-axis is in the plane of symmetry of the airplane perpen- dicular to the relative wind. The X-axis is in the plane of symm.etry perpendicular to the Z-axis. The Y-axis is perpendicular to the plane of sjTrm.etry. The origin of the stability axes is at the center of gravity of the air- plane, which for the present tests vvas located on the fuselage center line 25 percent of the mean aerodynamic chord behind the leading edge. Photographs of the model mounted on the trim stand are given as figure 1 and the construction of the stand is Illustrated in figure 2. Figure 2 shows that bearing A permits freedom in roll and bearing B permits freedomi in yaw. A calibrated coil spring was inserted in bearing A to provide stability in roll. This alteration m.ade pos- sible the measurem.ent of unbalanced rolling moments as a function of the angle of bank and thereby facilitated the trimjning of these m.oments by means of aileron deflection. Both bearings A and B were equipped '."'i th ball bearings to keep frictional effects to a minim.um. The trimming fin shown in figure 2 -r&s added to the trim stand to neutralize the drag yawing moments caused when the wind was on by the forvi'ard struts at an angle of yaw. Since this fin area was such that the trim stand was in complete equilibri'.mi of yawing momorts [Cj^ = C| ever the yaw range tested, the trim stand did not affect the directional stability characteristics of the m;odel. NACA ARR No. L5A13 Model The model used in the investigation was a _!_- scale 20 nodel of the North American 6-28 airplane. A three-view drawing and a photoe-raph of the riodel are given as figures 3 and 4, respectively. The model was equipped v.ith 2 four-blade propellers having a diameter of 8.£0 inches and set at an angle of pitch of 20°. Pov;er was furnished hj a direct-current controllaiDle-speed electric ir.otor rated, l/s horsepower at 15,000 rpm. The left propeller, which was kept in- operative during the tests, was so niounted as to windmill freely. The right propeller, which v;as used as the oper- ating propeller for all tests, was geared to the motor at a ratio of 1:5. Provision was made for reversal of the direction of propeller rotation. The model v;as equipped with partial-span slotted flaps (fig. 3), which were de- flected 45° for all tests. Sketches of the vertical-tall designs used in the Investigation are shov/n in figure 5 and sketches of the dorsal- and ventral-fin areas utilized in the rudder- free tests, in figure 6. Tail 2 represents the original vertical tail surface of the full-scale airplane and is considered typical of conventional vertical-tail design. The dimensional characteristics of this tail were varied to form the other vertical-tail designs. All vertical tails were constructed of the NACA 0012 section. In order to maintain similitude of hinge -moment character- istics as far as practicable, all rudders were of identical blunt-nose balance type with a balance area 12.2 percent of the rudder area. This type of rudder is of negative float- ing tendency and trails with the wind v/hen free. The dimensional characteristics of the full-scale airplane are given in the following table: Vv'ing: Area, sq ft 675.90 Span, ft 72.61 Aspect ratio 7.80 Root chord, in '. . - 161.13 Tip chord, in 67.00 Mean aerodynam.ic chord, in 120.09 Root section NACA 23C17 NACA ARR No. L5A13 Tip secti or. NACA 4409R Percent chora line with zero sv/sepback 33 Sweepback at leading- edge, deg 4.2 Dihedral angle , deg 2 Incidence, deg « 3 Georr.etric twirt (washout) , deg 2.5 Taper ratio 2.4:1 Fuselage J Length, ft 54.5 Section Circular Frontal area, sq ft 36.5 Horizontal tail: Total area, sq ft 183.20 Span, ft 26.35 Aspect ratio 3. 94 Dihe dral angle , deg . Stabilizer setting, deg 1.50 Length frorn hinge of elevator to center of gravity of airplane, ft 28.90 Elevator balance area, sq ft 10.63 Elevator area behind center line of hinge, sq ft 53.00 Vertical tail 2: Total area, sq ft , 74.90 Span , ft 10 . 68 Aspect ratio ] . 54 Length from hinge line of rudder to center of gravi ty of airplane , ft 27 . 40 Pin area, sq ft 35.66 Rudder area, sq ft 39.24 Rudder-balance area, sq :t T:' . 14 Rudder area behind hinge line, sq ft 50.10 (Pertinent data for tar'ls 1, 3, 4, 5, 6, and 7 are given in fig. 5.) Aileron (one of tvjo) : Area behind hinge line, sq ft ,,....»..... 20.91 Soan, ft 11.41 Mean chord, in ....,..,,.. 17,0 8 NAG A ARH No. L5a13 Flap: Total flap area, sq ft 30.5 Total span, ft 58.4 Type Slotted SPECIFICATIONS AND CRITZRICNS The NACA and Army flight: requirements for multiengine airplanes operating with asyrmnetric power were chosen to establish the proper test conditions. No separate attempt was made to reproduce the Navy specifications for asym- metric power because of the close similarity between the Navy and the NACA specifications. Specifications for Directional Control • ■ (Rudder Fixed) The NACA and Army specifications (references 1 and 2, respectively) for directional control of airplanes operating with asjmmetric power are as follows: NACA requirement (II-Z) 3 .- "The rudder control should be sufficiently powerful to provide equilibrium of yawing moments at ^ero sideslip at all speeds above 110 percent of the minimum take-off speed 'under the following conditions: a. Airplanes with two or three engines: With any one engine inoperative (propeller in low pitch) and the other engine or engines developing full rated power," Arm.y requirement E-2c ( 1) (c) »•- "The rudder control shall be pov.-erful enough to triia a multi-engine airplane for straight_ flight with less than 10 degrees of sideslip at 1.2 Vq jVo = stalling speed of the airplane, throttles "h [_ ""h —i closed, gear down, flaps in best tske-off condition] when the throttle on an outboard engine is abruptly closed (propeller in low pitch) and the other engine or engines are developing full take-off power. The flaps shall be in the take-off setting, and the gear shall be down...." NACii ARR :^o. L5A13 Specifications icr Directional Stability (R VI. cider Free) The NACA specification ( requirenient (II-?) 4 of refer- ence 1) relating to the rrouiremenbs for directional sta- bility with rudder free under asyiumetric power conditions is as follows: "The yawing moment due to sideslip (rudder free with airplane trlmned for straight flight on symmetric power) should be .such that straight flight can be maintained by sideslippjng at every speed above 140 percent of the mini- mum speed with rudder free with extreme asymmetry of power possible by the loss of one engine.'' Criterion for Vertical-Tail Effectiveness ".mder As^rmmetric Power Conditions Each of tlie specifications previously listed requires the dlrection;vl control or the directional stability of the sirplsne in question to be sufficiently povi/erful to balance the yawing m.oments created by as^Tnmetric power under certain specified flight conditions. It follov/s that the vertical-r.ai 1 effectiveness in flight m.ay be gaged by the maximum amount of as'i^.rrm-aetrlc power which such a tail can balance under the specified conditions. In this invfc stifration, ther-efore, the miaxlmum asymmetric power perrrdssible under the airspeed and trim conditions speci- fied by the Army and the NACA. v/as used to evaluate the effectiveness of the vertical tails tested. It should be observed that the flight specifications require that straight flight ">t complete equilibrium of lateral forces and momeats be maintained. In order to maintain such eauilibrium In flignt, the ailerons miust be deflected so thot the rolling moments caused by asymm.etric powtr ar-; balanced and the airplane assumes an attitude of bank, which nullifies the side force created by rudder deflect", on and/or angle of sideslio. Inasm.uch as an atti- tude of bank dots not affect the trim req\iirement s of the vex^tical tail surface, no attempt '.vas m^ade in the tests to simulate the balance of side force by angle of bank. Aileron deflection, howevr-r, directly afftcts dii^ectional trim by virtue of the yawing mom.ents created by such de- flections.. Gonsequeritly , the ailerons wore so adjusted 10 NACA ARR No. L5A13 for a] 1 tests as to naintain complete balance of aero- dy-nanlc rolling moirents and thereby to simulate flifrht cond-itj'ons correctly. TESTS Test Conditions "^he test lift coefficient was established from con- sideration of the specified airspeeds in bhe Amy and NACA requireirents . These values were converted to 11ft- coefflcient forms as follov/s: The NACA requirer.ent (II-S) c a (rudder fixed) speci- fies an airspeed equal to 1.10 times the take-off speed. If '^''t glee -off -^ assumed equal to l'2V,jj^j^, the airspeed requirement for this specification is equal to ■^•^^^min* If the maximum lift coefficient of the E-SS airplane is assumed equal to 2.0, the specified lift coefficient cor- responding to 1.22V . Is defined by the expression mm which equals 1,15. In a similar m.anner, the lift coefficient necessary to satisfy IIAGA requirement (II-P) 4 (rudder free) was found to be 1.02. The lift coefficient necessary to satisfy the Army requirement (rudder fixed) was calculated as 1,39. Because it was as- sumjed that slight changes in lift coefficient would not affect the model test results if the correct values of thrust coefficient were used, all tests were r'.m at a con- stant ancle of attack of 5°, which corresponded to a lift coefficient of 1.10. All tests v/ere run at a test velocity of 40 feet per second, which corresponds to a tert Reynolds number of 128,000 based on the mean aerodynamic chord of 0.503 foot. The aileron deflections for all tests were adjusted to provide equilibrium of rolling moments. Test Procedures Rudder fixed .- In the tests with rudder fixed, the model was mountea on the scand with the rudder deflected in the direction that counteracted the yaw caused by NACA ARR No. I5A13 11 apymjrietric power. ^yeasureiuents v^ere then taken of the maxi- m\xr, a:r;ount of asyrrinetric thrust the rudder would balance at an.'xles of yaw of 0*^ and 10° for rudder deflections of 0°, 5^7 10'^, 20°, and 30°. Rudder free .- The tests with ru:";der free were made by measuring the ar.ount of a.=';/mmetric ti.rust and anf-le of yaw produced by asynmetric power for various angles of yaw up to the an.o:le at which directional instability was encountered Tests with rudder free wez'e made of the model with each of the following vertical-tail arrangements: (1) Vertical tail alone (2) Vertical tail olu.s dorsal fin a (Z) Vertical tail plus dorsal fin b (4) Vertical tall plur ventral fin a (5) Vertical tail plus vsntral fin a plus dorsal fin a The absolute dorsal- and ventral-fin areas required for each test were determined from the percentages of the vertical tails being tested given in figure 6. No tests were made to determine the influence of aixciliary fin area upon the characteristics of tw-in tail 4. P ovv-er cal c ulations .- Tl-e thrust coefficients that were obtained in the tests ci the m.odel were converted to the simulated as^>Trjnetric brake horsepovcer of t;:e full-scale airplane by m.eans of the rel8.tionship vv, thp bhp = -^ TgV or /.-\5/~' K^ -w 2 bhp - -^^-;c-: (1) Tel 5f:0r]p 12 NACA ARR No. L5A13 The full-scale propeller efficiency r^ 'va? assun-ed to -?e equal to 0.75 for the calculations. Values of wing loading '■■'V'/s,,. and propeller diameter D v;ere obtained from the full-ccale character! stica of the North American E-28 airplane and were equal to 47.5 pound? per square foot and 14.7 feet, respectively. The value of the mass density of air p was chosen as 0.00258, which is its value at sea level mider standard abmospheric conditions. SifDsti tution of these values in equation (1) yields the relate onship T thp = 9900 (2) Cl" '' The values of Cj in equation (2) are those coi^respond- ing to tne airspeed specified in the Army and the NACA requlreinents and were deteririned as shown in the section entitled "Test Conditions." Substituting these values of lift coefficient in equation (2) yields the expressions defining the conversion of model thrust coefficient T^ to the estimated full-scale brake ?aorsepov,-er, which are: For n^dder fixed. KACA requirement Army requirement For rudder free, FACA requirement bhp = 30eOTc (3) bhp = 6C70Tc (4) bhp = 962CTc (5) RESULTS AND DISCUSSION The aata obtained in the investigation are plotted in figures 7 to 18. Figure 7 sho-z/s the rolling -moment coefficients produced by the ailerons used in the tests. Figures 3 to 10 present the values of the asyminetric- thrust coefficient balanced by means of rudder deflection. NACA kRR No. 15^13 13 Figures 11 to 13 ['ive the vilv.e.s of the asymr.etric-thrust coefficient balanced by the yawed model v^'ith rudder free. Data showinr the influence of dorsal- and ventral-fin areas upon triin characteristics are presented in figures 14 to 13, The test data in figures S to 13 vvere rearranged and converted to values of full-scale brake horsepower in figures 19 to 24. An index to all figures 3S presented as table I, Effect of Mode of Propeller Rotation The mode of propeller rotation in which the upper blade tips r.ove toward the fuselage is henceforth designated inboard rotation. The rotation in which the upper blade tips move out tov>ard the wing tip is designated outboard rotation. Almost all conventional airplanes are equipped with right-hand propellers. On inultlengine airplanes, the direction of propeller rotation with respect to the wine tips (inboard or outboard) is therefore determined by the location of the propeller. If the right engine fails, the direction of the operating propeller rotation is inboard and the airplane yaws in a positive sense. For left-engine failures, the operating propeller rotates outboard and the airplane yaw is negative. The results of the present investigation show that use of different modes of propeller rotation caused considerable difference :' n trim characteristics of an airplane operating under asymm.etric power. With onl.y one exception, the data presented in figures 8 to 13 indicate that the use of outboard propeller rotation decreased the values of permissible asyirirretric- thrust coefficient balanced by any given vertical-tail con- figuration and that this mode of rotation would therefore determine the minimum vertical-tail size. The exception occurred when twin tail 4 operated under the Army specifi- cations (fig. 9); in these tests inboard rotation was less favorable than outboard rotation. The difference in asyminetric power balanced by a given tail arrangement with either of the two modes of rotation appeared to increase in m.agnitude with the amount of di- rectional stability and of control being applied. The largest differences occurred at large rudder angles and for tails 6, 6, and 7, which have high aspect ratios. Particu- larly large effects of propeller rotation were observed when the rudder was free. 14 NACA ARR No. L5A13 The magnitude of effect of reversing the propeller rotation ha? been illustrated In figure 19. Thi? figure presents the calculated values of periaissible brake horse- power for both modes of propeller rotation for the repre- sentative rudder deflection^ of 20° (figs. lG(a) and 19(b)) and for that angle of sides] ip at which directional In- stability was encountered in the tests with rudder free (fig. 3S(c)). This angle of sideslip was between 10"^ and 12° for almost all the conditions tested. The results presented in figure 19 shov/ that the difference in as;;nn.- metric power balanced by the vertical tail for Inboard and outboard rotation was about 4C0 horsepower for most conditions and was as large as ICOO horsepower for some. The effects of changing the direction of propeller rotation appear to be ex'plained by the data of reference 4. Refer-Ence 4 concludes that use of inboard propeller rotation with the flaps down caused the slipstream to con- verge toward the tail and thereby increased the contri- bution of the tail surfaces to directional stability for small to moderately large angles of yav'. This slipstream displacem.ent would result in a beneficial effect of in- board rotation upon the trimming action of the vertical tail surfaces, particularly for twin tail 4, which under TACA specifications (3 = 0°) appears to be partly im- mersed in the slipstream jet. Reference 4 also concludes that outboard rotation causes the slipstream- jet to di- verge. Consequently, this m.ode of rotation increases the tail effectiveness at large angles of yaw but is less satisfactory in this respect than the inboard mode of rotation for other angles of yaw. This reasoning explains the favorable effect of outboard propeller rotation upon twin tail 4 when operating at an angle of sideslip of 10°. At this angle, ov:ing to its original lateral displacement, this tail lies within the slipstream. The data obtained in the tests indicate that for twin-crginc airplanes equipped with single vertical tails and cor.ventional right-hand propellers, the failure of a left engine will imtpose the miore severe flight conditions. For airplanes eouipped with twin vertical tails, hov/ever, the failure of a right engine should prove more critical to the fulfillment of the Army requiremients. Similarly, it may be reasoned that use of propellers rotating in- board on both wings (symmetric rotation) would be advpn- tageous for airplanes equipped v/ith single fins both to improve tail effectiveness and to m.ake the handling of controls similar regardless of tire location of the NACA ARR Mo. L5A13 15 inoperative engine. Conversely, syirjnetric outboard rotation should be favorable for airplanes equipped with tv.-in fins. Effect of Vertical-Tall De?ign Effect of vertioal-tail area .- The effect of varying vertical-tail area was obtained frorr a study of the test data for geometrically similar tails 1, 2., and 3. The data for these tails with rudder fixed vifere converted to values of full-scale brake horsepower and plotted in figure 20. The data of figure £C(a) shov/ that increasing the vertical-tail area resulted in increases in the asynunetrlc power balanced by a given rudcer deflection at zero angle of sideslip, The.'^e increases, however, are not directly proportional to the increase in tail (rudder) area, as might normally be expected; this lack of proportionality Indicates the preoence of secondary slipstream effects upon the vertical tail surfaces. Such secondary effects are probably produced by the sidev/ash angles generated at the tail by inflow into the slipstream jet as well as by the more direct effects of slipstream velocity. Further Investigation, however, is required to establish a complete explanation of these secondary slipstream- effects. The data in figure 2G(b) illustrate the favorable effect upon the asymmetric power characteristics of in- creasing the vertical- tail area at an angle of sideslip of IC'^ , These data show that, v\?hen the airplane is side- slipping, the directional stability of the vertical tail surfaces reinforces the action of the rudder control in nullifying the effects of asymmietric power, and higher values of as^Tnm.etric thrust can therefore be balanced by a given vertical-tail arrangement. The magnitude of the effects of cirectional stability can be obtained from, a study of the curve for a rudder deflection of O'^ (fig. 20(b)), which is directly indicative of the rudder-fixed directional stability. These data show that the directional stabilit:/ contributed by tail 1 barely balanced the unstable yawing nomiSnts created by the yawed fuselage-wing comoination, ?toking the tail area larger than that of tail 1 increased the djrectional stability, as v.ould be expected. 16 • NACA ARR IJo . L5A13 The e'f'fect?? of iricreas: ng tril area noted "in the tests math rudder fixed were also observed in the tests Vvith rudder free. Fipure 11 illustrates the influence of tail area upon the rudaer-free trim characteristics of the model operating under asymmetric power. In this figure, the data inJ-icate that freeing the rudder of tail 1 was cestabi li.'^ing, as would normally be expected since the rudder type employed had a negative floating ratio. Because of tiie slender m.argin of stability associated with tail 1, the destabilizing action of freeing the rudder was sufficient to cause directional instability. Making the tail area greater than that of tail 1 increased the directional stability contributed by the tail surfaces sufficiently to overcome the destabiliz- ing effects of the fuselage and, consequently, perm^itted increases in the asjrmmetric thrust balanced by the vertical tail surfaces. £o mp ari son of twi n^ tail and s in gle tail . - T v; i n tall 4 may be directly compared '.vith tail 2 inasmuch as both tails ver? of the same aspect ratio and equal area. Be- cause the twin trll was located almiost directly in the slipstream, the twin tail was more effective than the single tail at zero and small angles of sideslip. Figure 21 sho.vs that the influence of po^ver at p = 0^' made tail 4 almost as effective as tail 3, a surface of equal aspect ratj.0 but possessing 50 percent greater area. At angles of sideslip greater than 0°, however, tail 4 was less effective than tail 2 with the rudder fixed at an angle of sideslip of 10^ and with the rudder free (fig. 22), These data confirm, trends noted in the past and indicate that the directional stability contributed by a twin vertical tail la less than that contributed by a single tail of the sam.e aspect ratio ana equal area. The in- creased directional stability achieved by use of the sin.srle tall is partly ascribed to the favorable end-plate effect of the horizontal surfaces upon the load charac- teristics of ■ the vertical surfaces. In addition, the single tail has but one ro^t juncture compared wi th two for the twin tail and therefore is less affected by inter- ference effects. It should be noted that the curves for tail 4 for rudder free (fig. 22(b)) do not pass through the origin but fall above and belov; It depending on this m^ode of propeller rotation employed. These curves indicate that reversing the propeller rotation altered the sldewash caused by the propeller sufficiently to reverse the local angle of attack of tail 4 at small angles of sideslip. NACA ARR No. L5A13 The rer-ults of the te.'^^ts Inuicate that choice between single and twin vertical tails would, aepend largely upon the pilot's handling of the controls following a sudden engine failure. If the rudder c~'ntrol can he applied he- fore the airplane reaches a moderately large angle of sideslip, the twin-tail den'ign should be more suitable; otherw-'ise, the single vertical-tail design would be preferable . Effect of increasing a-^pect rat ic- The effect of increasing aspect ratio was deterrriinedi from a comparison of the data obtained with tail 6, a surface of twice the aspect ratio of tail 2, with corresponding data for tails 2 and 3, These data are shown in figure 23 and indicate that doubling the aspect ratio of tail 2 has approximately the same effect as increasing the area by 50 percent at the same aspect ratio (tall 3). This effect is in close agreemer.t with the wind-tunnel force data of reference 5, which shov;; that doubling che aspect ratio of a surface from. 1.5 to 3.0 increased the lift-curve slope from 2.2 to 3.1. For a given rudder configuration, such a change in lift-curve slope v/ould result in an increase in tc^tal tail load, or trimming effectiveness, equivalent to that obtainable by appro.xim.ately a 50~percent increase in area, C om car i_s o n of c onventional t ai 1 and all-m.ovable tail wi th linked tabT - The question has been raised whether the efficient action of the all-miovable tall reported in refer- ence 6 arose largely from the "all-movable" feature or from the fact that the tall was of high aspect ratio and had the inherent advantages associated with tails of that type. For the present investigation, therefore, tests of the all- m.ovable tall 5 were supplemented with tests of tail 7, which is identical with tail 5 except that tail 7 is of conventional - that is, fixed-fin - design. A comparison of the effects of tails 5 and 7 upon the characteristics of the airplane operating with asym- metric power is shown in figure 24. These data indicate that the all-movable tell is m.arkedly more effective than the conventional tail at zero sideslip with the rudder fixed and with the rudder free. At 10° sideslip and with rudder fixed, however, the all -movable tail was only sllehtly more effective than the conventions 1 tall (fig. 24(a)) . These test resiAlts miay be explained by use of the curves showing typical tall loads (fig. 25) . These curves 18 NACA ARR No. L5A13 ?how the vai-iatior. of tai?. load vith vertical-tail inci- dence and rudder deflection for e. conventional and an all-novable tail. The tab area of the all-mo'/able tail is assumed equal to the rudder area of the conventional tail. The variation of the load with deflection of the all-movable tail is indicated by the dashed line in figure 25. Thi? variation Is due to the linkage betveen the tab and the movable forv;ard surface. The slope of the load curve is determined from the linkage ratio 6iji/i+; which, for the case investigated, was equal to 1,12, If the effect of powei' is ignored, the angle of attack (tail incidence) of the c r^nventicnal tall at ^i = 0^ is also zero. The rudder deflection therefore produces changes in load along a path coincidental with the zero ordinate. For e:>:a:;riplc , a rudder deflection of 1C° produces the tail load corresponding to the load indicated by point a, For the all-movable tail, however, a rudder deflection of 10° causes a simultaneous change in tail angle of attack and tab deflection and produces the load indicated by point b. Consequently, at zero sideslip, the all-movable tail is capable of producing much larger yawing moments v-ith wl : ch to balance the effect of asym- metric power than the conventional tail. At moderate angles of side-^lip (1C° to 15"-^), the con- ventional tail operates in the high-lift region of the lift curve of the ta-* 1 and consequently produces tail ]oads of an order comparable with those px'-oduced by the all-movable tail. The conventional tail may conceivably produce tail loads even greater than those of the all- movable tail because the conventional tail r's unrestricted in the use of rudder. The all-movable tail, hovifcver, is limited for a given linkage ratio to the rudder deflection that produces the tail Incidence at maximum, lift. Further deflection would cause the entire surface to stall. In balancing the effects of esymnptric pov;er, the superiority of the all-mjovable tail to the conventional tail was most msrked in the rudder-free tests. This su- periority can be ascribed to the fact that the hinge- moment characteristics of the all-movable tail force the entire tail tD float against the wind v/hen free (positive floating ratio) and crnsequentl^r Increase the directional stability of the airplane. In considering the advantages of the ail-r.iovable vertical tail over the conventional tail, it should be observed that "snaking" oscillations may be induced by control-surface friction with improperly designed tails having positive floating ratios. (See reference 6.) NACA ARR No. L5A13 19 decreasing rudder chord is shown by the te:^t data for tails 6 and 7 (figs. 10 and 1?). These data show that, although the rudder of tail 7 had only one -half the area and one -half the cliord of the rudder of tail 6, the rudder of tail 7 balanced approxirrately two-thirds as much asymmetric power at zero sideslip and approximately seven-eighths as much power at 10*^ sideslip as the rudder of tail 6. These data are in agreement with conventional trends because it is known that decreasing the rudder cnord increases the yaw- ing mcm.ent per unit rudder area, With rudder free, tall 7 balanced a greater amount of asyiTiiretric povv^er than tail 6, which indicated a favorable effect of reduced rudder area upon the rudder-free direction- al stability. This action occurred because the rudders of tails 6 and 7 are of the type that trail with the v.'ind and so reduce the directional stability when set free. Con- sequently, tail 7, because of its smaller rudder area, created smaller destabilizing mom.ents when the rudder was set free and so balanced a greater amount of as^nnnetric power. Effect of rudder deflection ,- The data obtained in the tests showed that increasing the rudder deflection increased the amount of asymimetric power balanced by the vertical tails at a decreasing rate. Effect of dorsal and ventra l fins.- The data illus- trating the 'effect of adding dorsal- and ventral-fin areas to tails 2, 3, S, €, and 7 are presented in figures 14 to 18. No data are presented for tail 1 because the addition of dorsal and ventral fins did not noticeably lessen the directional instability associated with t'lvis tail arrange- ment. The test data indicated that the addition of auxiliary fin area Increased the directional stability at large angles of yaw and thereby increased the miaxim-um. amount of asymmetric thrust balanced by tne tail surfaces vi/hen the rudders were free. Increases in maxir.iuiii asymanetric thrust of the order of 30 to ICO percent were observed in the tests. The addition of ventral-fin area was generally found to be m.ore effective than the addition cf an equal amount of dorsal-fin area. The use of a com.blnation of dorsal- and ventral-fin areas (dorsal, a and ventral a) -was 20 NACA ARR No. L5A13 pen^rplly rore effec"i"ive than a sin.'^le dorral f;.n of the same total area (dorral b) . A"jleron deflections e ction s req ui red to t rim agymmetr l£ entativo plot of total aileron deflec- thru'-t . - A repi'esL.-. ^^^„ ^^ „,„„^ ...^ ... ..„,.^.> tions required to trjm the rolling raorrents created hj asymmetric thrurt is presented In figure 26. These deflections were alvi'ajrs obtained bv equal up-and-dov-Ti movoF.ents of the ailerons. Calculated values ai-e also presented in figure 26. These calculations were made by using the method presented in reference 7. The calcu- ]ated lift increments created b^/ the operating propeller were rrultiplied by the lateral arm of "the propeller to obtain rolling moments, which v.'ers converted to aileron deflections required to ti-iim by use of the datr. in figure 7, The results presented in figure 26 shov; that, although the scatter way considerable, the test data agreed fairly well with the calculated values and Indicated that moder- ately large aileron deflections woula be required to main- tain straight flight undei- as:\Tiimetric power conditions. OOFCLnSTONS The following conclusion"^ were drawn from trim tests of a twin-engine-airplane m.odel operating under asymmetric power (single-engine) conditions specified by the NACA and Army Ai r Forces: 1. The direction of rotation of the operating pro- peller had an important effect upon the asymmetric power that could be balanced by a given vertical- tail design. Single vertical tails were mo5t effective when the oper- ating propeller was rotating inboard, Tv;in tails, however, were most effective when the operating propeller v/as rotating outboard. 2. An all-movable vertical tail of aspect ratio 3 with, a linked tab was more effective than the conventional tail of the same aspect ratio and equal area in balancing asymmetric pov;er, particularly when the rudders were free. The all-movable tall was markedly superior to the con- ventional vertical tail -^f normal aspect ratio (1.5). 3. The single vertical- tail desiQ-ns generally balanced a greater am.ount of asyrnme trie power than twin vertical NACA aRR xMo. L5A13 tails of the rair.e aspect rat5_o and equal area, particularly when the rudder war; free. At sinall angles of sideslip, however, it was, possible to balance more power by rudder deflection of the twin tails than by rudder deflection of a single tail. 4. Increasing the aspect ratio of a vertical tail resulted in increasing its triirining effectiveness under asymmetric power conditions by an amount proportional to the accompanying increase in lift-curve slope. 5. The tri'Ti.-.iing effectiveness of the vertical tail surface increased almost linearly with the vertical-tail area. Increaslrg .'■udder deflection and rudder chord in- creased the trirrjT'l::r effectiveness of the vertical tail under asymmetric concitlons at a decreasing rate. 6. When the ^udder was free, addition of dorsal- and ventral-fin areas increased the capacity of the vertical tail surfaces to balance asymmetric-power effects at moderate angles of sideslip. Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Lanslev Field, Va. 22 NACA ARR No. L5A15 REPER^NCDS 1. '.7j.].ruth, P. R.: Requirements for Satisfactory Flying dualities of Airplanes, NACA ACR, April 1941 / (Clasfification changed go Restricted Oct. 1943.) 2. Anon,: Stability and Control Requireirents for Air- planes. A/iF Specification No. C-1815, Aap. ?1, 1943. 3. Shortal, Joseph A., and Cstc-rliout, Clayton J.: Prelirrinary Sta^^ility and Goncro] Tests in the NACA Free-Plight V^nd Tunnel and Correlation with Full-scale Fll^lit Tests, FACA TK No. 810, 1941. 4. Pitkin, t.'arv'.n: Free-Flieht-Tunnel Investigation of the Effect of Mode of Prooeller Rotation ' upon the lateral-Stability Characteristics of a Twin-Engine Airplane Model \vith Sinale Vertical Tails of Dif- ferent Size. NACA ARR No. 3 JIB, 1943. 5. Zir.'!ir.erman, C. H.: Characteristics of Clark Y Airfoils of Small Aspect Ratios. NACA Rep. No. 431, 1952. 6. Jones, Robert T., and Kleckner, Harold F. : Theory and ireliriinary Flight Tests of an All-I^ovable Vertical Tail Surface. NACA ARR, Jan. 1943. 7. Pass, H. R.: V.'ind-Tunnel Study of the Zffects of Propeller Operation and Flap Deflection on the Pitching Moments and Elevator Hinge ivlotnents of a Single -Enshrine Parsuit-Tyoe Airplane. NACA ARR, July 1942 ; NACA ARR No. L5A13 23 TABLE I.- INDEX TO FIGDRES Figure Inscription Remarks ^ Photographs of test model mounted on trim stand In Lfingley free- fllffct tvinnel Model with tall 2 2 Sketch of test model mounted on trim stand, which permitted freedom in yaw and roll. In Langley free-flight tunnel 3 Three-vlee drawing o£ w».-5cale twin-engine icodel teaCed In Langley free-flight tunnel with asymmetric power « Photograph of twln-englnc model used in trim teste In Langley free- flight tunnel Model with tall 2 S 1 — Plan-form and dlnenalonal characteristics of seven vertical tails tested on a ^-scale model of a twin-engine alrulane In the Langley free-flight tunnel 6 Various fin arrangements tested with vertical tails on a ^-scale model of a twin-engine airplane In the Langley free-flight tunnel Figure Te«t apeel- tleatlont Taet condition Tall arrangement Operating-propeller rotation^ Cnrre Remarks Tall Dorsal [Ventral 7 2 j Propeller off C[ against 6^ Aileron calibration 8 a NACA (p = 0°) Rudder fixed 1 to 3 Inboard and outboard Tg against 5^ Elrectlonal-control run b NACA (p = 10°) Rudder fined 1 to 3 Tj against 6^ Do. 9 a NACA (p = 0°) Rudder fUed 4 Tg against 6p Do. b Army (p = 10°) Rudder fixed 4 T, against 6, Do. 10 a NACA (0 = 0°) Rudder fined S to 7 7g against 6^ Do. Army (p = 10°) Rudder fined S to 7 (i-j_ ..__■ T^ against 6^. Do. 11 NACA (Rudder freej Rudder free 1 to 3 Outboard Tj against p Dlrectlonal- stablllty run NACA (Rudder free) Rudder free 1 to 3 Inboard Tf against p Do. 12 NACA (Rudder free) Rudder free 4 Outboard T, against p Do. NACA (Rudder fre^ Rudder free 4 Inboard T5 against p Do. 13 NACA (Rudder fre^ Rudder free 5 to 7 Outboard Tc against p Do. NACA (Rudder fre^ Rudder free 5 to 7 Inboard Tg against p Do. 14 NACA (Rudder fre^ Rudder free 2 combinations^ Outboard Te against p Effect of dorsal- and ventral-fin area NACA (Rudder free) Rudder free 2 Inboard Te against p Do. 15 MACA (Rudder free) Rudder free 3 ^= Outboard Te against p Do. NACA (Rudder fred Rudder free J do Inboard Tj against p Do. iRlght propeller operative. ^f7nm*Tlnit1 "n* tested ape tell alone, dorsal a, dorsal and ventral e, Testl*! «. *orsal b. .NATIONAL ADVISORY COimiTTEB FOR AERONAUTICS NACA ARR No. L5A13 24 TABLE I. - INDEX TO FIGURES - Concluded Figure Tejt speci- fications 16 NACA (Rudder free) Teat condition Rudder free Tall arrangement Tall ! Dorsal [Ventral Tall aloB*, dorsal a, dorsal a and ventral » Operating-propeller rotation^ Outboard Curve Tq against p Effect of dorsal- and ventral-fin area Remarks NACA (Rudder fre^ Rudder free All combinations^ Inboard Tg against p Do. 17 NACA (Rudder free) Rudder free do Outboard T(. against p Do. NACA (Rudder free) Rudder free do- Inboard Tf against p Do. 18 NACA (Rudder free) Rudder free do Outboard Tj against p Do. NACA (Rudder free) Rudder free Inboard Tq against p Do. 19 NACA O = 0°) Rudder fixed 1 to 7 Inboard and outboard bhp for 6p = 20= Effect of mode of rotation Army (p = 13°) Rudder fixed 1 to 7 -do- bhp for 6- = 20° Do. NACA (Rudder free) Rudder free 1 to 7 -do- bhp for 10 z < 1- (/) S a: Q H Z O 5 LU Q O cn I LU i NACA ARR No. L5A13 Fig. 3 1734" Propeller diam.j 8.^" Wire landing gear Figure 3. - Three-view drawing of l/zo -sca/e twin- engine model as ijsstecc //? Long/ey -fnse- f//g/if- Tunnd iv/^Ji asym/r^ei-r/c poiver. NACA ARR No. L5A13 Fig. 4 NACA LM^L 27625 NACA IJ*U. 27627 v.. Figure 4.- Twin-engine model of B-2.6 airplane used in trirr tests in Langley free-flight tunnel. NACA ARR No. L5A13 Fig. 5 ^95 Cut-out for horizontal tail Arn^^psrr.^^ \ 550 S^ ffiu x pj ji . ^ Tail 122 5^^ y^/^^n^ 5>6.d^0l^ 3 11.0 12'Z 5U 15^ J_ 16.5 12.2 I.5U 11.0 12.2 T5U ^eo 1 1.0 12.2 ^ II.Q 12.2 3.0 25d^0 2$.0 JIO 12.2 3.0 (one of two) ItLTab o. Tail 5 pivots about 0.27 MAC li of GUI taib /6.^3" behind e.g. NATIONAL ADVISORY iw COMMITIEt FOR AERONAUTICS lQi\5 (a/l-movab^) Tail 6 Tail 7 Figure 5. -Plan -form and dimensional charactehetics of seven vertical tails tested on a 1/20- scale model of a twin-engine airplane /n tha Langley free -f//(^ht tunnel. NACA ARR No. L5A13 Fig. 6 NACA ARR No. L5A13 Fig. 7 i; I— <■ '\ ? \ \, 3 \ M \, \ \, \ K \ \ \ V \ \, \ ^ \ ^ \ \l 8 O s Q CM o <3 O '9 q O 'iUdioijidOD lUddiouz-duiiiOf^ ^ 03 8^ y) c: I NACA ARR No. L5A13 Fig. 8a, b o Operating pmpeJ/er rofat/nQ ouftxxuzi H -Operating propeller roTof/ng /ntjoara t I I 6 .Outtjca/zl _k rofat/orp' I nbc xm 3^/ 'Oti t^ y > /^ ^ i V 4 Oatpcc otcdioi 1 n /! II III /nhmrri mtnt/nn -, y _-^ h / c /- Y / UUTDOOm 1 f \h ^r t-Z/O ty/o '/ / A y ,r 1 (p.) A//iC/i 5p)ecif/ca//or)5 (i^'^O"). / i- y ^< p 1 > / 1 / K ^inpoara - rotdtiort y h ouTDcora u /c /or 1 /^ 20 30 10 30 10 20 3D O 10 flight rurtter defject/onj 6rj cleg (b) Army 5pea-fical/on5 ^=/o). F/guns 3.- Asymme^r/c-potyer chanychrf^Hcs of a iw/n- enp/ne - airplane moPel eQ.u/ppGcL m^ yer//ca/-fa// designs I , Z, and 3. cL suc/i fhaf f/ie ro/t/rxp momsnf^ equal Oj df^dS") <^-0°j o(^S° left prope//er ^/nd- millingj rudder fixed.. national advisory COMMITTEE FOR AERONAUTICS NACA ARR No. L5A13 Fig. 9a, b O- Operaf/ng propeller rofaf/ng oultnord Operating propeller roratinq inboard I I O .32 .Z4 .lb .06 O 1 1 Jnboara / roi ht/ on i ^ / / X p / / / / J / y \^ oa, ^/rx irrt / i rotai/on <( ia)mcA specif icQt/ons ((^^O"). 32 .Outboard fotation-^ to ^o 30 '^ f?/ght rudder deflect ion ^ dr, deg (b)Army 5peaf/caf/on5(/3-/0^). figure 9- Asynnmetric-px^/ver choracfensf/cs c/a tw/n-engine- airplone model eouipped vMiTh vertical- fail design -4. 6c 3ucfi fhafftie rolling moments equal o ,6^45°. 6e-'0: oc--j°-/eff propeller tA/irdrniiling , rudder fixed. NATIONAL AllVisoSy "— . •;n ni.i.(KilAuflcS NACA ARR No. L5A13 Fig. 10a, b O operotinq propeller rotal/ng outboord + Operating propeller rotating inboard Hinge line OflQU ^HinQelim ,oftab Jail 5 Toil 6 lai I deflection L,om , , ^, , dr/L.^i.i2 ' 17ighi rudder deflection, 6 r cleg (b)Army specif/cafiom (.S=/OV' Figure /o - A^ynimctr/c -power characteristics or a fw/n-engine- airplane model eguipped witH verticol-tail designs S, 6, and 7. 6q ^uch tnoff/^e rolling moment5 equa/Oi 6f = 4S] 6^ = O'- oc- 5". le// propeller windmil/ing j roa'cfpr p/xecT. NACA ARR No. L5A13 Fig. lla,b -^Mode/ airecr/onOJ// un^/otDt .32 ^ .16 t § I AADdel d/rectonoJly un^tabJe at Tc-0 1 r r — ^ /■ ^ c /-">_ ( r^ \ A J i (a) O/xroting prcpe//er rotaf/ng outboard. .oc Mode/ di rect/onot 1 i/ unstab/e ot Tc-0 JATK NAL ADVI 50RV .24 ( :(iMH IIU ; FO 1 AEI ONA JTICi .16 C ? ( r^ r/ .08 ? ^ J] C r "( P J ] J /O 20 JO I 5 / d 3 o9 D / o a 3 o /Ing/e ot 3/des//pjf3, deg (b) Operating p/vpe/ter m/at/ng /r}tx)ard. T/gure /t .-/^s/m/m/r/c -pother ct/orocter/st/cs ot a /mn-e/ig/ne- Qjrpanc mode/ egw/uped with i/ert/cat- /q// cfes/g/7S /,2,Qnd 3. 6a sudi that the ro////ig nxmenfs equatO;S^--4S°Se=0''- cc--5l/ett propeller windmilling- rudder Tree. NACA ARR No. L5A13 Fig. 12a, b O ^ Model d/rect/ 0/1011/ unstoD/e — Vertical fojl ojone (B rooe of- fiu>oy 32 .24 ./6 .OB I i5 o I CI o c Se .24 5 -0= = = =^ (a) OperQt/ng prope/Jer rotating outJDoard. 32 .lb .08 O -G NATIONAL ADVISOHY COMMIIlEt FUR AERONAUTICS -to , Q ^ , lo^ ^n JO Angle of sides iip.p, aeq (b) Cperating ,oropeller m/ar/ng //?^oara. f/gore /Z.- As/mmefr/c-porver choracten^^f/c^ of a tt^in-engine- Qirolonp model eguippecl lA/itti vertical -tail design ^^. . 6o so'c/7 r/xjt r/7c rot ting moments eguotO; — _^ \ 3 9 wl ^ f i r t —( \^ y —( ^. I I nnprnling nrncy^JIfir rnlnr/nn n/jlhonr,^. siJ^ICS Ang/e of ^/desl/pj (3 > deg (b) Qperal/ng propeller rofaf/ng inixxjrcL . Figure 13 .- /Isymmelric -pother chamd'er/5f/cs cf a ft^/n-eng/ne- ojrplane model egu/pped i/v/f/i yerfjcoJ-tojl di^/gns ^,6,aria7. cfa sucf) ifnf //le ro///ng momen/s equal 0;Sf--45°6s--0lcc^5° leff propeller mnOm/lling ■ ruMer free . NACA ARE No. L5A13. Fig. 14a, b ^Mode/ a/rectionoJl/ unsToJDie TaiJZ Dorm/d a DoooJ VenM — a a — — a — D — VenfmJ ol K^ V I a I .24 .lb JDS "^ _ — — — — — — . -^ 1 TL' ,^ - '/ti r — J *-rnt • T § {a)Ope/ating p/Tpe//er /abf/ng ou^^xjo/rZ. ■c^ .lb 4^ - — -7^- ^<' >— — -— & V' - — " ■^-^ J ^ j08 ■ /- J ^\ s 1 2 3 O ^ t D 6C k Angle of 3/c/e3lip^ .^, ofeg (t>) Operaf/ng prQOG//er m^a^/ng /n/xxjroL. F/gure /^ . -£/^/ecf o/ ab/:soJ-Qrct vpntroJ-f/n oreos upon t/ie Qsymme^nc-poyver choractenst/cz of a fw/n-engine- ourpiane model egu/pped yyifh vert/ ca/- feu I c/es^gn 2. 6a.^uch that the /xDJ/ing momentz equal O; 8f = 45; 6^-0] ex -- 5°; Jeff prcpe/fer windmill ing ; ruddet free. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICI NACA ARR No. L5A13 Fig. 15a, b Mode/ c/>r£c//onQJ/y ans/alj/e -raz/B Dorsal ti a o- Oorsa/ \/en/raJ 8— -3 a ■ — o a. a. VentroJ a ? I I I .32 c '>- .Z4 J > /■ _„ — - - -^ " .16 / n-r L^ i » JDS V T A i (o) Opefofing propeJ/er rofaf/ng ou/Poa/xX. /a cO ^ Ja , -^O , JD 60 NATIONAL ADVISORY Anq/e of sjdeslipj^^dsg cummittee for aeronautics (b) Opera/jng prope/fer rok3///x/ /nboanX . F/gure 15 . - Pf/ecf of dorsaJ- onol venr/aJ-fJn areas upon the Qsy/vmefnc-poyyer cnorac/ef/^f/cs of a. fwjn-^ng/ne.-curp/a/)e model equjpped w/rh veP/cal-fojl design J. da Such thof /t>e roj/ing nnomenfs equo/ O; 5^=45^, 6e-0] cc= 5° /eft propeller i/^tndmil/ing, rudder free. NACA APR No. L5A13 Fig. 16a, b ■Model directionaJly unstoJble ro//5 o- Dorsil YenfraJ - o. — - a a Ventral Q .40 .32 ./6 All l-i-W^J-i-i t 1, \ . 1 \ s ¥^ k-XXi \ \ i x\x \\\ \ \'is\ r^^ / y ' =L> Alaxima/T) ^ / ■2i^//ab/e — J k f,^ 1 i '' / ~' tt ) k m - 4Q [a) Qpepf/ng propel ler rp/o//nn ouitoa rd.. r % I ,U ' ' I — 1 ^ ' i^ ' — I I A^ Mational advisory Angle cf fia^Sj/0^(3,&/ Wvmittee for aeronautics r2)j Ope far /ng propeller rokji/ng /ndosrct. ^iqure 16 . - Effecf of dorsal- and venfral-f/n areas upon //)e Qsymmefnc - poller choracfensHcs of a fmn - eng /ne- at rpiane model eqa/pped w//h \/srfical-i-ci/l des/gn J~. da such lf)af //le rolling momenis equal 0;6f-.45° 6^-0] cx-S°j left propeller mndm/II/ng) rudder free . NACA ARR No. L5A13 Fig. 17a, b Mcx:lel cf/recf/anoJ// u/xsfad^ rQ//c> o — Dorsal \^nffnl — Ol cl ex. b — Ventral a. Jb .06 0<\ "k- ._ . , m ' \^ ^My- - — -— § W\ ) ^ J I (a)Cpemf/ng prcpe/ler /D/o/ing au^£>oarc2:. .40 .32 .24 .lb ^ .ce> TU /(U CXJ JU i^r chorocter/st/c^ of aiw/n-enginc-ojrp/one nrxJel eguz/Queo/ \Nitt} v^rficoJ-tcul design 6. c5a such ttiat the filing mo/mehh equal O-, Sf = 45° de = 0°,c!: = 5 J /eft propeller wrCrDiWirg ^ rudder f/^. r? / Mi ^1 . — — _ — "** 1 .1 M / — ^ w w 1 1 o a o 3 o 4 O / 'O NACA ARR No. L5A13 Fig. 18a, b Mode/ c//rect/cnQj/y unsfo/Dt ro//r o- Dorsal ventm.1 V- — a — --a a a. — o — Ventral a .40 ^.32 C ^ 24- k ^ o O .40 I ■ — — 1 J '7^ >— — = ■ =- — =— T -♦ <^ 1 T HI - — — i / -^ ^ ] 5 r ,2 p ((X) n/)prnr//ig proppJ/rr /pfrif/rn ry// -hnnrr/. iO , 2Q 3Q, <0 , SO 'Oes/// (D) Opera fjng pmpe//er rofaf/ng joboord. ngur&l8.-£'ffoof o/aorWil'Qrxi venrraJ-f/n arsas upon the asymme/rjc- pO)/ver c/iarac/enstic5 of a fyi//o-eng/ne.-Qirplane rncdel eQu/ppsa With verlicoJ-rajl dssjgn 7. 6^ 3uch r/nf fhe ro/Z/ng moments equal O; 6f^s°; c5e-o°- <^-^''; /e/f propel /or wind/7v/lJr)gpudoJer free- NACA ARR No. L5A13 Fig. 19a, b.c Hinge /jnc-, ^lOil 4- \Ta\\ 5 TOJI 6 Tail 7 pie c/ two) laJhmoiaUe) 3eco- I ^ 600 2400 I6CD i q,0 f ^ 1 I- 1 " t I- Inbcyard /'dof/on U7X Out^tyoard rokj/ton _ — £5ti mated by extrofiolaf/on of c/aJa (a) MCA specifications (dr^^ZO"). 1 ■& ^ 24K0 l(CO aoD o I t -J 1 ^75^ i tl n ,- I ^ 3200 (b)Army specif icaf/ons (d'r = ZO°) . I I 2400 l(£0 &X) Direction- rr* oJiy un5lat)/e A I 4 4 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS t I It ' / 2 J 4 5 6^ f ^-jQJi (c)A/ACA :5pecificaf/ons ■ rudder free (JO'^^-^/i'^- Figure /9. 'Fffed of mode of prcpeJIer rotaf/on upon fhe permiss/b/e asymmetric dra/re ttorsepoiver., oaJcu/ated from /est cloJa for the North Americon B-2& curptane cperatinq on one engine under N/\CA ond Artny flighf specifications. NACA ARR No. L5A13 Fig. 20a, b H/nge line ICOOv I ■^ I2CO 830 Tojl I To// a ToJl 3 leoo ^ 4CC> % a) o 1 .7&// ,? c5. [ U^^^ Taj /2 ^ ^- -< > / ^ -- ^ . — . ^0 Till / - ^ X- - / ^^ ^ ^ ^-t- J 7x /^ '-y^ ^ — ? c ' - — — ' y J <- ^ " " j \- il ^ ^ ::d ^ - ^- — /A — Lf :>J 7-H= >- 6 0- ^43q (c^;/l^/)(::>1 speci/'/cahons f^=o°)- aooc I6CQ ^ i2aX % -SO) NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS lOJl I ^ -01 Y7Z. %a Twl Z-' '-7>^: ^ Toy/ J ~30 ^:^ lo O .Q2 .CH. .06 .08 ./O ./^ ./'f- .16 -18 Vert/col- toil area, >St , fract/on /[/jng Q/ea. (b)Arm\/ 5p6c/f/ca//on5 0S=/O°J. model Test Oai^cLTor the Norf/? Afnencan B'2Q airptone op&xmfig on one engine under t\/AC/l and /\rmy ft/ght ^pecificol/on5 Operat/ng p/tpe/ter nofat/ng ou/txord . NACA ARR No. L5A13 Fig. 21a, b 2^00 4 8 IZ J^ ZO Z4 2^ Right rudder cfef/eci/on , di-, dep f/gumll -Companion of t/ye asymmefnc-poouQr c/xfrac/er/^- f/C6 ofana/rp/ane egalpped uj///? fu/^n one?' ^//^ig/e \/ert/cQ/1a//-5./VAC^.^oec/f/co/-/oo5j /S-O". NACA ARR No. L5A13 Fig. 22a, b /nboa/Tii rofoHon I I I I I Qu^doard. /v/a^/on Ouidoarct rofQf/on_ 'Inboa/n rotation lONn AEVISOHY COIIMIT EE (OR /ERONAUTICS I i I I Right rudder d.efJection, c^ , cLeg (a) /^r/ny specifications Qs ~/0°). I200 QOO 400 a /z Ang/G of sicLes/ip, ji3 , cLeg (t) /VACA specifications ; rudder free. Figure 22. - Comparison of the asynimetr/c -potver characteristics of an airplane equipped with twin and sing/e vertical tails and at various angles of sideslip. NACA ARR No. L5A13 Fig. 23a, b Q Q V O i MOO 2000 I6C0 1200 WO 4O0 O Tail Aspect ra+/o Z /■S4 I 4 Q 12 16 20 24 23 32 P/ghf- rudder def/ecHon j cfr > deg (a)A/ACA 5pGC/f/ca//or)5 f^^O% Ang/e of sides l/p^ jff j deo (b) A/AC A 6pec/f/ca-f/on5jru-o- -G- H-X t- -o- -o O-G .Calculated NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .08 .16 .24 .3£ (K> Asymmetric - thrust coefficient^ Tc .4S Figure Z6~ Aileron deflections reqaired to trim rolling moment creoled by asymmelric po\A/er (5=0°, operating pnopeller tvtatinq outt>oord. UNIVERSITY OF FLORIDA IIHIII'IIIIIIIIII 3 1262 08106 463 5 UNIVERSITY OF FLORIDA DOCUMENTS DEPARTMENT 120 MARSTON SCIENCE LIBRARY P.O. BOX 117011 .^,,n<^x GAINESVILLE. FL 32611-^011 USA