/JACfr u SO \S NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED February 19^5 as Advance Confidential Report L5AI3 DETERMINATION OF TEE STABILITY AND CONTROL CHARACTERISTICS OF A TAILLESS ALL-WING AIRPLANE MODEL WITH SWEEPBACK IN THE LANGLEY FREE-FLIGHT TDNNEL By John P. Campbell and Charles L. Seacord, Jr. Langley Memorial Aeronautical Laboratory Langley Field, Va. NACA WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- • nically edited. All have been reproduced without change in order to expedite general distribution. Document Digitized by the Internet Archive in 2011 with funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/determinationofsOOIang 7, 2 9)1 lO NACA ACR No. L5A13 CONFIDENTIAL NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE CONFIDENTIAL REPORT DETERMINATION OF THE STABILITY AND CONTROL CHARACTERISTICS OF A TAILLESS ALL-WING AIRPLANE MODEL WITH SWEEPBACK IN THE LANGLEY FREE-FLIGHT TUNNEL By John P. Campbell and Charles L. Seacord, Jr. SUMMARY An investigation to determine the power-off stability and control characteristics of a tailless all-wing air- plane model with sweepback has been made in the Langley free-flight tunnel. The results of the free-flight- tunnel tests were correlated with results from force tests made at high Reynolds numbers in order to estimate the flying characteristics of the full-scale airplane. The investigation consisted of force and flight tests of a 1^.3 -foot -span dynamic model. The effects of flap deflection, center-of-gravity location, and addition of vertical-tail area were determined. The following conclusions were drawn from the results of the investigation: The full-scale airplane will undergo a serious reduction in stick-fixed longitudinal stability at high lift coefficients unless early wing-tip stalling Is eliminated. The directional stability of an all-wing airplane without vertical tail surfaces will be undesirably low. The effective dihedral of an airplane of this type should be kept low. An elevon and rudder control system similar to that used on this design should provide sufficient control. INTRODUCTION The desire to obtain improved performance for mili- tary airplanes has recently increased the interest- in tailless-airplane designs. One of the most promising tailless designs, from the considerations of performance, CONFIDENTIAL CONFIDENTIAL NACA ACR Mo. L5A13 is the Darge all-wing airplane or "flying wing." Inherent in the all-wing airplane, however, are certain undesirable stability and control characteristics that must be elimi- nated before this design can be considered satisfactory. In order to study these stability and control character- istics and to find means of improving them, an investi- gation is being conducted in the Langley free-flight tunnel (designated FFT) of a free-flying dynamic model of a tailless all-wing airplane with sweepback. The present report gives the results of force and flight tests of the model with windmilling propellers. Tests were made with the lift flaps retracted and deflected. For some tests, auxiliary vertical tail sur- faces were installed on the model. The effects of changes in the center-of-gravity location and trim lift coeffi- cient on the flight characteristics of the model were determined. In order to estimate the flying characteristics of the full-scale airplane, the test results were correlated with results of force tests of a similar design run at high Reynolds numbers in the Langley 19-foot pressure tunnel (designated 19-ft PT ) . SYMBOLS The following symbols are ixsed herein: C L lift coefficient (Lift/qS) Cp drag coefficient (Drag/qS) C m pitching-mement coefficient (Pitching moment /qcS) Cj rolling -moment coefficient (Rolling moment/qbS) C n yawing-moment coefficient (Yawing moment /qbS) Cy lateral-force coefficient (Lateral force/qS) c chord, feet c" mean aerodynamic chord, feet CONFIDENTIAL NACA ACR No. L5A13 CONFIDENTIAL S b q V P P i|/ a h pb 2V P n Cl M wing area, square feet wing span, feet dynamic pressure, pounds per square foot airspeed, feet per second mass density of air, slugs per cubic foot angle of sideslip, degrees angle of vaw, degrees (for force-test data, i< = -p) angle of attack, degrees static margin, distance in chords from center of gravity to neutral point helix angle generated by wing tip in roll, radians rolling angular velocity, radians per second rate of change of rolling-moment coefficient with ' 6C 7 helix angle 6« 6\ \2V '1 rate of change of yawing-moment coefficient with P angle of sideslip, per degree (dC n /dp) rate of change of rolling-moment coefficient with angle of sideslip, per degree (dC^/d,:) flap deflection, degrees elevon deflection, positive down, degrees (with subscripts r and I to indicate right and left elevon, respectively) rudder deflection, positive down, degrees (with subscripts r and I to indicate right and left rudder, respectively; If both right and left top rudder surfaces are deflected simultaneously as longitudinal trim flaps, no subscript is used) CONFIDENTIAL k CONFIDENTIAL NACA ACR No. L5A13 R Reynolds number APPARATUS The investigation was made in the Langley free- f light tunnel, which is described in reference 1. A photograph of the test section of the tunnel showing the model in flight is presented in figure 1. Force tests to determine the static stability characteristics were made in the Langley free-flight tunnel with the model mounted on the six- component balance, which is described in reference 2. The mass and dimensional characteristics of the model are as follows : Weight, pounds 2.55 Wing area, square feet 2.51 Sean, feet L|..3 Aspect ratio 7o6 Wing loading, pounds per square foot 1.02 Radius of gyration in roll, k^, foot O.7S Radius of gyration in pitch, ky, foot 0.35 Radius of gyration in yaw, k z , foot 0.82 Mean aerodynamic chord, foot 0.655 Sweepback of 0.25-chord line, degrees 22.00 Dihedral, degrees Taper ratio (ratio of tip chord to root chord) . . 0.25 Root chord, foot 0.937 Tip chord, foot O.23I4. Eleven: T:/pe Plain Area, percent wing area 5*^-0 Span, percent wing span 33.00 Rudder: Type Snlit, drag Area, percent wing area 2.86 Span, percent wing span 20.00 Vertical tails: Type Twin center fins Area, percent wing area 4.. 00 Aspect ratio 2.00 CONFIDENTIAL NACA ACR No. L5A15 CONFIDENTIAL 5 Airfoil section Modified NACA 103 Root, percent thickness 21 Tip, percent thickness 15 Geometric twist, degrees o Aerodynamic twist, degrees (approx.) 1| The component parts of the model are identified in the tables and figures as follows: Wing W Propeller shaft housings H Propellers P Vertical tails; two tails mounted on nacelles, each tail having 2 percent of wing area V Split flap (center-section lift flap, 5f = 60°) . . . F Combinations of these letters represent the combination used in the tests. The standard configuration is desig- nated WHP. A three-view drawing of the model is presented in figure 2. Photographs are given in figures 3 and I|_. In plan form the wing has both sweepback and taper and has a split flap that extends from the center line of the airplane to the inboard ends of the elevons. For all flap-down tests, the flaps were deflected 60°. The control surfaces consist of elevons that extend from 0.33- to 0.71- and split rudders (fig. 5) that extend from O.71H to 0.91^-. The split rudder is so linked with the elevon that in flight tests the lower surface of the rudder moves down with the downgoing elevon and the upper surface moves up with the upgoing elevon. This linkage arrangement provides additional effective aileron- and elevator-control-surface area as shown in figure 6 . The upper surfaces of the split rudders can be *im flaps •im when the lift flap is deflected. The lower surfaces of the split rudders remain at zero v/hen the top surfaces are deflected as trim flaps. The upper surfaces of the split rudders ca deflected upward simultaneously to serve as tri: to provide pitching moment for longitudinal tri: The controls of the model were operated in flight by electromagnets in the same manner as described in reference 1. CONFIDENTIAL 6 CONFIDENTIAL II AC A ACR No. L5A13 For some tests vertical tail surfaces having a combined area of l. percent of the wing area were mounted on the propeller-shaft housings to provide additional directional stability. (See figs. 2 and I). . ) For propeller-on tests the model was equipped with two freely windmilling two-blade pusher propellers. A modified NACA 103 airfoil with a thickness of 21 percent at the root and 15 percent at the tip was used on the model. The trailing edge was reflexed enough to give a slightly positive pitching moment at zero lift. This airfoil was used to obtain a maximum lift coefficient in the free-flight (low Reynolds number) tests more nearly equal to that of a full-scale airplane than is possible to obtain with other airfoils (especially low-drag airfoils) at low Reynolds numbers. The free -flight-tunnel model was almost identical in plan form to the model used in the tests at higher Reynolds numbers in the Lang ley 19-foot pressure tunnel. The models differed in airfoil section, geometric dihedral, and geometric twist. The airfoil sections of the model tested in the Langley 19-foot pressure tunnel were NACA 65(316) -019 at the root and 65(3l8)-015 at the tip; the geometric dihedral of this model was 2° compared with 0° for the free-flight-tunnel model. The model used in the Langley 19-foot pressure tunnel had I|. geometric twist, whereas the free-flight-tunnel model had a geometric twist of 6°. The aerodynamic twist for both models, however, was approximately I4. . TESTS Force tests were made to determine the stability and control characteristics of the model with flaps retracted and deflected. The moments were computed with the center of gravity at O.25 mean aerodynamic chord and are referred to the stability axes. The stability axes are defined as an orthogonal system of axes in which the Z-axis is in the plane of symmetry and perpendicular to the relative wind, the X-axis is in the plane of symmetry and perpen- dicular to the Z-axis, and the Y-axis is perpendicular to the plane of symmetry. The conditions in which force tests were made are given in table I. CONFIDENTIAL NACA AGR No. L5A13 CONFIDENTIAL 7 Flight tests were made at lift coefficients varying from 0.3 to 0.8 with flaps retracted and from 0.6 to 1.1 with flaps deflected. The center-of-gravity position was varied from 20 to 25 percent of the mean aerodynamic chord for flight tests in both the flap-retracted and flap- deflected condition. Table II gives the conditions for which flight tests were made. RESULTS AND DISCUSSION In interpreting the results of the free-flight-tunnel tests of the tailless all-wing model the following points were considered: (1) The tests were made at very low Reynolds numbers (150,000 to 350,000); hence, the results of the tests of a similar design made at high Reynolds numbers (about 6,600,000) were used in estimating the flight characteristics of the full-scale airplane from the free- flight-tunnel test results. (2) The controls of the model were fixed except during control applications; hence, no indications of the control-free stability of the design were obtained. (3) In determining the control effectiveness of the design, no consideration has been given to control forces. (k-) No power was applied to the propellers during the tests. The results, therefore, cannot be used to predict power-on stability. Longitudinal Stability Force tests .- The results of force tests made to determine the longitudinal stability and control charac- teristics of the model are shown in figures 7 an( i 8- On these figures, data from tests of the model of similar plan form tested at high Reynolds numbers are also plotted. The slope of the pitching-moment curve for the flap- retracted condition of the free-flight-tunnel model changes from negative to positive with increasing lift coefficient. This change in slope indicates a change to CONFIDENTIAL 8 CONFIDENTIAL NACA ACR No. L5A13 static longitudinal instability at high angles of attack. This change in stability is characteristic of swept-back wings because of the tendency of the wing tips to stall first. The instability appears to be much greater for the free-flight-tunnel model than for the similar model tested at high Reynolds numbers. This difference is probably explained by the fact that the difference in the Reynolds numbers at the root and tip sections of this design causes a much greater difference in stalling char- acteristics on the small-scale model than on the model tested at high Reynolds numbers. For the flap-deflected condition (fig. 8), the pitching-moment curves for the free-flight-tunnel model were very similar in shape to those obtained with flaps up but did not turn up at high lift coefficients as much as the curves for the flap-retracted condition. The data of figure 8 indicate that most of the change in shape of the pitching-moment curve from flap up to flap down was caused by the upward deflection of the trim flap. The flap-deflected pitching-moment curve from high-scale tests (fig, 8) indicates practically no change in longitudinal stability with increasing angle of attack. The difference in the angles of zero lift indicated in figures 7 and 8 for the two models is probably caused by the difference in the location of the chord line from which the angle of attack is measured. The difference in the slopes of the lift curve is probably a result of the difference in the Reynolds numbers of the tests. It is unlikely that these differences in lift characteristics would cause appreciable differences in longitudinal flight characteristics . Flight tests .- The longitudinal stability as noted in the free-flight-tunnel tests was satisfactory up to a lift coefficient of 0.7 with flaps retracted and 1.1 with flaps deflected with the normal center-of-gravity location (25 percent M.A.C.). Above these values of lift coeffi- cient, however, difficulty was experienced in flying the model because of a tendency to nose up and stall after disturbances in pitch. This behavior was believed to be a direct result of the change in longitudinal stability at high angles of attack, which was indicated in the force-test results by the change in slope of the pitching- moment curve. Although at times the pilot could prevent the nosing-up motion by applying down-elevator control, CONFIDENTIAL NACA ACR No. L5A13 CONFIDENTIAL 9 the nosing-up tendency was considered a very objectionable characteristic that would probably prove dangerous for a full-scale airplane. This nosing-up tendency should be expected on any airplane having pitching-moment character- istics similar to those of the model. (See fig. 7«) The longitudinal stability of the free -flight -tunnel model was satisfactory at those lift coefficients at which the static margin h was O.0I4. or greater ( C T = 0.7, flaps retracted; Ct = 1.1, flaps deflected) and flights were possible at conditions at which the static margin was as low as 0.02. On the basis of the force-test results it appears that the static longitudinal stability of the corresponding airplane at high angles of attack would be greater than that of the free -flight -tunnel model. The data of figures 7 and 8 indicate that the airplane with the normal center-of -gravity location would have a static margin of O.0I4. up to a lift coefficient of 1.0 with flaps retracted and up to the stall with flaps deflected. The stick-fixed longitudinal stability of this particular airplane design, therefore, would probably be satisfactory for all power-off conditions except at high lift coeffi- cients with flaps retracted. Longitudinal Control The force-test results presented in figures 7 ant ^ 8 indicate that the longitudinal control provided by the elevons was sufficient to trim the model over the flight range for flap-retracted or -deflected condition with a total elevon deflection of about 20°. Inasmuch as the force-test results of the model tested at high Reynolds numbers indicate much more powerful elevon control than was obtained with the model at low Reynolds numbers, it is probable that the elevator control for the full-scale airplane will be satisfactory in flight. In the flight tests, the model could be trimmed over the speed range with a total elevon deflection of about 20°. For the flap-retracted condition, the upper surfaces of the split rudders were deflected with the elevons for longitudinal trim. Abrupt elevon deflections of ±5° from the trim setting provided adequate longitudinal control for keeping the model flying for all stable conditions. CONFIDENTIAL 10 CONFIDENTIAL NACA ACR No. L5A13 On this design it is possible that the most critical condition for elevator control will be at take-off. Unless careful attention is given to the location of the landing gear, the elevons alone may not be powerful enough to meet the Army requirements for getting the nose wheel off the ground at 80 percent of take-off speed. Use of the trim flaps in conjunction with the elevons will help provide enough longitudinal control to meet this requirement . Lateral Stability Force tests .- The lateral stability characteristics of the model as determined by force tests are shown in figures 9 to 11. The values of the effective-dihedral parameter C^ and the directional-stability parameter C n obtained for the different test conditions from these figures are plotted in figure 12 in the form of a stability diagram. The values of C n and C^„ for corresponding r P conditions for the model tested at high Reynolds numbers are also presented in figure 12. The values of C„ for the flap-retracted condition at angles of attack of 0° and 6° are relatively low (about O.OOOJO). increasing the angle of attack to 12° with flaps retracted caused an increase in C n to 0.00055. This increase in C n „ with increase in lift coefficient is r characteristic of a swept-back wing. The lower values of C n shown in figure 12 for the p model tested at high Reynolds numbers are attributed to the lower drag of this model. For an all-wing tailless design with low dihedral, the drag of the wing contributes a major part of the static directional stability. The values of C^ shown for the free-flight model in figure 12 correspond to an effective dihedral angle between 2° and 1±°. The value of C 7 increased with increasing lift coefficient as expected for the swept- back wing. The higher values of C 7 for the model tested at large Reynolds numbers is caused by the fact CONFIDENTIAL NAOA ACR Wo. L5A13 CONFIDENTIAL 1] - that this model had 2° geometric dihedral whereas the free-flight-tunnel model had 0° geometric dihedral. Flight tests . - The lateral stability characteristics of the model noted in flight were fairly satisfactory except for low directional stability in the flap-retracted condition. This low directional stability was shown principally by slow lightly damped yawing oscillations that were started by gust or control disturbances. The directional stability was not dangerously low, however, inasmuch as neither divergences nor unstable oscillations were noted. The adverse yawing noted in flights in which aileron control alone was used was quite small because the elevons were deflected upward tor-ether for longitu- dinal trim and therefore operated as "trimmed-up" ailerons, which usually produce only small yawing moments. Deflection of the flaps or addition of the vertical tails caused noticeable improvement in the damping of the yawing motion of the model, and the lateral stability characteristics at these conditions were considered generally satisfactory. The effective dihedral of the model appeared to be satisfactory, inasmuch as no excessive rolling during sideslip was noted and the lightly damped yawing oscil- lations v/ere accompanied by very little rolling. Previous free-flight-tunnel investigations have shown that, for an airplane with low directional stability, low effective dihedral is necessary to avoid a poorly damped rolling (Dutch roll) oscillation. It is probable that the lateral stability character- istics of a full-scale airplane of the design tested would not be so good as those of the free-flight model because the values of C ne of a full-scale airplane will probably be lower than those for the free-flight model. At the higher lift coefficients, which could not be reached in the free -flight-tunnel tests because of longitudinal instability, the requirements of the airplane would be more severe for directional stability and the airplane would probably be considered unsatisfactory in this respect. In order to secure satisfactory flying characteristics with a tailless all-wing airplane of this type, it appears desirable to maintain a low valixe of effective dihedral and to supplement the directional stability of the wing by means of vertical tails or an automatic stabilizing device. CONFIDENTIAL 12 CONFIDENTIAL NACA ACR No. L5A13 Lateral Control Aile ron cont rol.- The aileron control provided by the elevons appeared to be weak in the flight tests. Abrupt eleven deflections of ±15° did not provide satis- factory aileron control in flight. Previous free-flight- tunnel tests have shown that, if aileron deflections greater than *I5° are required for satisfactory control on a model, the ailerons on the corresponding airplane are likely to be weak. A better quantitative indication of the weakness of the aileron control was obtained in the force tests, the results of which are presented in figures 13 and 1I4. and which are summarized and compared in figure 15 with results of tests at high Reynolds numbers. Computed values of the helix angle pb/2v produced at different lift coefficients by various elevon deflections are shown in figure 15. The values of pb/2V were obtained by multiplying the force-test values of rolling-moment coef- ficient by 0.8/Ct, . (See reference 3«) The high P Reynolds number data of figure 15 indicate that the flying-qualities requirement for a minimum value of 0.07 for pb/2V is not met by this design at lift coefficients above about O.I4. with ±15° elevon deflection. The free- f light-tunnel force tests indicate even weaker aileron control but this result is partly attributed to the low Reynolds number of the tests, to the wing section used, and to the initial reflex of the trailing edge of the wing. The free-flight-tunnel test results do indicate, however, that linking the rudder surfaces to move as ailerons with the elevons provides a substantial improve- ment in aileron control. In order to obtain satisfactory aileron control with elevon surfaces located well inboard of the tip as on this design, larger-chord surfaces than those on the free- flight-tunnel model should be used or the rudder surfaces should be linked with the elevons in order to provide greater effective elevon area. Rudder control .- The split rudders on the model provided sufficient yawing moments to balance out the adverse yawing moments encountered in the flight tests during aileron rolls. Inasmuch as the yawing moments caused by aileron deflection were small (fig. li|) because of the initial upward deflection of the elevons for CONFIDENTIAL NACA ACH No. L5A13 CONFIDENTIAL 1? longitudinal trim, the rudder yawing moments only had to oppose the adverse yawing moments caused by rolling. The adverse yawing moments caused by rolling were apparently small for the model, as indicated by the small amount of adverse yawing in flights with rudders fixed and elevons alone used for control. These results indicate that the rudder control of this all-wing airplane should be adequate during normal flight. Usually the most severe requirement for rudder con- trol of multiengine airplanes is that the rudder control balance the asymmetric yawing moments introduced by the failure of one engine during a full-power climb. Calcu- lations based on the force-test data presented in figure l6 indicate that, with rudders of the size and type used on this design, an airplane of this type having a 150-foot span and two 3000-horsepower engines would meet the Army requirements for maintaining steady flight with 10° or less sideslip at 120 percent of the stalling speed with one engine inoperative and the other engine operating at full power. CONCLUSIONS The following conclusions concerning the power-off stability and control characteristics of large all-wing tailless airplanes with sweepback were drawn from the Langley free-flight-tunnel test results and from a corre- lation of these results with results obtained from force tests made at high Reynolds numbers: 1. Stick-fixed longitudinal instability at high lift coefficients, or at least a serious reduction in longitu- dinal stability, should be expected for airplanes of this type unless the premature stalling of the wing tips is eliminated. The upward deflection of a trim flap at the wing tip will reduce the tendency of the tips to stall first and will thereby improve the longitudinal stability at high lift coefficients. 2. The directional stability of this type of airplane without vertical tail surfaces will be extremely low. Although the airplane v/ill be flyable, it will probably not be considered entirely satisfactory because of the tendency to sideslip to large angles following slight gust or control disturbances. CONFIDENTIAL lil- CONFIDENTIAL NACA ACR No. L5A13 3. The effective dihedral of an airplane of this type should be kept low in order to minimize the amount of rolling accompanying the lightly damped yawing oscil- lations that are likely to be encountered. [).. An eleven and rudder control system similar to that used on the model in these tests should provide sufficient longitudinal and lateral control for an airolane of this tvpe . Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. REFERENCES 1. Shortal, Josenh A., and Osterhout, Clayton J.: Preliminary Stability and Control Tests in the NACA Free-Flight Wind 'Tunnel and Correlation with Full-Scale Flight Tests. NACA TN Nc. SlO, I9I+I. 2. Shortal, Joseph A., and Draper, John W. : Free-Plight- Tunnel Investigation of the Effect of the fuselage Length and the Aspect Ratio and Size of the Vertical Tail on Lateral Stability and Control. NA.CA ARR No. 3D1?, i°l+3. 3. Kayten, Gerald G. : Analysis of Viand- Tunnel Stability and Control Tests in Terms of Flying Qualities of Full-Scale Airolanes. NACA ARR No. 3J22, 19l;3. CONFIDENTIAL NACA ACR No. L5A13 15 CONFIDENTIAL TABLE I FORCE-TEST CONDITIONS FOR TAILLESS ALL-WING AIRPLANE MODEL IN THE LANGLEY FREE-FLIGHT TUNNEL Test a (deg) (deg) Configu- ration (a) 6 e (deg) *1 (deg) 5 Pp (deg) Figure 1 -k to 20 WHP 7 2 -I4. to 20 WHPF 8 3 -k to 16 WHP -10 -10 -10 7 k -1+ to 16 WHPF -ko -1,0 8 5 -I4. to 16 WHPF -10 -ko -ko 8 6 -30 to 30 WHP 9 7 6 -30 to 30 WHP -10 9 8 12 -30 to 30 WHP -20 9 9 8 -30 to 30 WHPF -10 -ko -ko 10 10 6 -30 to 30 WHPV -10 11 n 6 -30 to 30 W -10 11 12 to 12 WHP -10 (Right only] 13 13 to 12 WHP 10 (Right only) 13 H+ to 12 WHP -20 (Right only) 13 15 to 12 WHP 20 (Right only) 13 16 to 12 WEP 10 lit 17 to 12 WHP -10 H+ 18 to 12 WHP -20 11+ 19 to 12 WHP 20 11+ 20 to 12 WHP ±20 16 21 to 12 WHP ±lj.0 16 22 to 12 TOP ±60 16 a Explanation of configurations is given in section on "Apparatus." CONFIDENTIAL NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ACR No. LSA1} CONFIDENTIAL l6 TABLE II PLIGHT-TEST CONDITIONS OF TAILLESS ALL-WING AIRPLANE MODEL IN LANGLEY FREE-FLIGHT TUNNEL Lift coefficient Configuration ?a) Center-of-gravity location (percent M.A. C. ) 0.3 to 0.8 WHP 0.25 .6 WHPV .25 .6 WHPV .22 • 5 WHPV .20 .6 to 1.1 WHPVF .25 .6 to l.i WHPF .25 .7 WHPVF .22 • 7 WHPVF .20 a-r. Explanation of configurations given in section on "Apparatias ." NATIONAL ADVISORY COlffilTTEE FOR AERONAUTICS CONFIDENTIAL NACA ACR No, L5A13 Fig. 1 0) T3 O e bo .H < > I— I O E- -G ! 55 10 ' Ed Q 55 o c c 3 be l 0) >-H M <-• bo C O c o o to a; E-* bo NACA ACR No. L5A13 Fig. 2 is- it ^v ^ 1 1 i' \ 1 A \ / \\ / \i C i LA ■t. <* v S _ I |5*. :i I i i ! 1 I 1 NACA ACR No. L5A13 Fig. < i — i O O O NACA ACR No. L5A13 Fig. 4 O o DO 0) rH C M •H 0) O T3 nj CD c +j CO C 0) o <-3 13 rH oj ■^ >> t-i -J <►-• ctJ < O -H i— i rH Eh S -H Z a; x w •H 3 Q > < hH Be. ■tj . Z C rH o O 0) o ^ C > . a) "* rH bo > ■H ^H •H ■ CO rH ■^> r-H c (H V 1 e d) -i a' T3 T3 3 tJ 1 +3 ■H -H A 03 1 lO CD S-i 3 UO NACA ACR No. L5A13 CONFIDENTIAL Fig, Left lower rudder surface adds to elevon area Right aileron control (longitudinal trim flaps ) Right upper rudder surface adds to elevon area Right aileron and rudder control (longitudinal trim flaps 0°) If rudder deflection is greater than elevon de flection, de fleeted rudder surfaces are not moved by elevons Pight aileron control „ (longitudinal trim flaps -40°) Lpp f r judder surfaces deflect with elevons only for elevon deflections greater than -4C° ^-^ Initial deflection-'!^ / Tor trim, -40° o° +5 / V Pight aileron and rudder control (longitudinal trim flaps -40° I For rudder deflection, upper rudder surface deflects from -40° trim flap position When elevons are used as elevators, the upper rudder surfaces deflect with the elevons NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 6.- Elevon and rudder arrangement used to obtain additional aileron effectiveness. CONFIDENTIAL Fig. 7 NACA ACR No. L5A13 4- 8 Ang/e of attach } oCjdeg 20 ./ O -/ Pi/cfing -moment ooeftictent , C m figure 7.- Lift) drag, and p/tcf)/ng-moment c/io/xxc/er- /st/cs of /node/5 of /Uittess at/- wing curptone with sweeptxtct; fi/aps retracted (1/VtiP). NACA ACR No. L5A15 Fig. 8 IS 6 e d r Source of (deg)(deg) data o & o rrr a -40 FfT --0-/0 -40 rrr a o -JO /9-ff PT ■/o -JO ^ 4 a /a /6 Ang/e of o/tacH } oc ) aeg dO ./ P/fcfi/ng-rnonent coeff/c/enfj C^ Fig< Fig. 9 NACA ACR No. L5A13 8^ .2 O -2 ^ (tea) (deg) o o O □ e> -to O 12 -20 -30 -20 -/O O 10 20 Angle of yaiVj t/Jjdsg JO NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 9. -Effect- of a note of attack on the lateral ■stability characteristics of the lang/ey free-f/zaht-funne/ mode/ of the tailless all-iv/ng a/rp/ane iv/th sweep- Dock. Flaps retracfed (WHP); o}=0 6 . NACA ACR No. L5A13 Fig. 10 -20 -fO O /O 20 Angle of yaw , i// , deg NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Fiqure to.- Effect of flaps on lateral stability charac- ter/sf/cs of Langtey free - f tight -tunne/ model of tailless a//-w/nq airplane with sweepback. Fig. 11 NACA ACR No. L5A13 -.02 -30 -20 -to o io eo Angle of yaw > W ; deg 30 CONFIDENTIAL Figure It- - Effect of propeller housings and verf/caj tails on lateral stability characteristics of longley frse- f I laht - tunnel model of fail less oJI-winq cvr/jtane with sweeploacrx. CK = 6° ; C L --070. NACA ACR No. L5A13 Fig. 12 ,o 3 k V5 1 ^ n^ St $ $ i !S < I— I Z w a *3 ^ S S Fig. 13 NACA ACR No. L5A13 CONFIDENTIAL P2 .0/ |^ -.0/ £ -.03 .0/ IT 18 is ^ 5 -.0/ o 4 a Ang/e cf attack , ex. J ahg CONFIDENTIAL NATIONAL ADVISORY COMMITOE FOR AERONAUTICS figure. /3. - l/ar/of/on of e/et/on effectiveness with ang/e of oftacA for langtsy free. -f//ght '- tonne/ model of fas/tess oJI-wing curplane. with sweeptxuzk. NACA ACR No. L5A13 Fig. 14 CONFIDENTIAL 8 5 .0/ o -D/ .0/ O -Of o- -10 -zo 10 20 r™f *=- r^-f fes= „ \ 4 3 Ang/e. cf attack / ck. / cfeg CONFIDENTIAL ta NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS F/gure t4* — rlo/t/ng and yaw/ng moments produced by cteftect/ng as e/evons /tie spttt rudders on the uingfcy rree-fl/ght-tunnej model of- cl tcuttess cJt- \AJing curpJane with sweep&acfr. Fig. 15 NACA ACP No. L5A13 QOQ§ 6CCV/DCL/ ' /\?/qc/ ' s/buo X'l&H NACA ACR No. L5A13 Fig. 16 CONFIDENTIAL ,0/ — ^ c r* * I t: 1 > t * > ~ ybtv/ng moments proc/uced by r/g/ir sp//t -rudder def/ec+to* on f'he LangJey free -f//&hf- fat/me/ mode/ of a. tat f /ess a//-U'tr>g a/r/o/one wrrh sweepback . CONFIDENTIAL .yNIVERSITY OF FLORIDA 3 1262 08104 988 3 UNIVERSITY OF FLORIDA DOCUMENTS DEPARTMENT ^S^f ,ENCEUBRARY GAINESVILLE, FL 3261 1-7011 USA