Document
REPORT No. 796
AN INTERIM REPORT ON THE STABILITY AND CONTROL OF TAILLESS AIRPLANES
By LANGLEY STABILITY RBSEAFICH DIVISION COMPILED by CHAELES J. DONLAN . SUMMARY S F B O L S lift coefficient Problems relating to the stability and control o f tadless drag coefficient airplanes are di8cussed i n co&ation o f contemporary experi- r o b - m o m e n t coefEicient ence and pradice. In the pea& stde o f the design qf ia2less yawing-moment coefficient airplanes, it appears that: pitch-moment coefEicient (1) Sweepback a3ord.s a method o f supplying tail l e n g t h for airspeed diredwnal and 7ongitwEinal stability and control and d o m yawing angular velocity th utdizalion o f a higMi$fEap b u i i n f r o d m undesirable tip density of air stalling tendencies that must be overcome before the admntages mnss of airplane o f sweepback cun be r d i z e d .
dynamic pressure ;also, pitching angular velocity (2) The dumping in pitching appears 20 b e litUe e $ & on thu Iongitudina.l,behumkr o f the airplane provided the static margin is never p m W to become negatiw.
thrust coefficient ( s 3 ) (3) Ths dwedwnal stability m w t be a6 great a6 for conmn- h e - m o m e n t co efEcien t iwnal airplanes if the same requirements regarding satisjktory angle of attack stability and control chur&tics are to be &ed to.
angle of sideslip
(4) T h s in$uence o f i!.h,e lateral resistance and i % dumping
angle of sweep in yawing on the flyins pvalities is somewhat obscure; however taper ratio; ratio of tip chord to root chord it is believed that these paranwtms wiU be o f secondury im- wing area, except n s designated otherwise by subscript portance $ adeqvafe directional s W i c y is supplied.
wing chord, except ns designated otherwise by subscript (6) On accouni o f the di$ieulties encountered in obtaining mean aerodynamic chord adequate stability and control with taZess airplanes, it appears napect retio i!ha.t a thorough reetduation o f the relative perjormunce to be distance of aerodynamic center from center of gravity expected from taiuess and conventional designs should be mmi5 vertical displacement of thrust mis from center of bdore proceeding further wii!h stub%ty and control s t u d h .
gravity (positive when thrust mis is below center of IXTRODUCTION gravity) wing span, except ns designated otherwise by subscript Much interest has been shown in tailless airplanes during propeller diameter the pnst few years. A number of tailless-airplane designs stick force have appeared and prototypes of several of these designs trailing-edge angle (see f i g . 11) have been flown extensively. It appears desirable at this landing-gear angle (see fig. 13) t h o to nmplify and expand an earlier work (reference 1) control-surfm deflection relating to the stability and control of tailless airplanes in the light of the recent flight experience acquired and the related studies that have accompanied the development of new designs.
It is the purpose of this paper to assemble and record some expressions of fact and opinion pertaining to numerous problems that have assumed significance in tailless-airplane to supply specXc quantitative design design rather than data. Tho problems specifically discussed in this paper pertain to the requirements and attainment of longitudinal and lateral stability and control and to spinning, tumbling, and steadiness in flight ns regards gunnery and bombing platform. A discussion is also included of some of the relative merits of tnilless and convrntiond nirplnnes.
REPORT NO. 7 96-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS wings hnve indicated that triling-edge split flaps nro partic- ularly suitable for swept-back wings because of the reln- tively small pitching-moment increment accompanying the production of a given lift increment (reference 2). The rntio of the pitching-moment increment to the lift increment produced by a flap depends, of course, on the position of the centroid of the flap load relative to the aerodynamic centor of p relative density the wing. The centroid of the flap load has been observed
(%)
to move forward dong the wing chord as the hinge-line position of the flap is shifted forward, with the consoquenco that the ratio of the flap pitching-moment increment to the flap lift increment is reduced. The a-tent of the forward movement of the centroid of the incremental flap load nc- companying a forward shift of the flap hinge line that may be Subscripts : expected for full-span trailing-edge split flaps i s given in figure 1. It mas noted in reference 3 that the ratio of the flap i flap; also, flipper pitching-moment increment to the flap lift increment could a aileron be considerably reduced by moving the flap hinge line fonvn.rd e elevator with only slight losses in the magnitude of the flap lift nsso- t tab ciated with a given flnp deflection. It appeam, therofore, r rudder that shifting the hinge line of the flap affords a promking $4 about quarter point of mean aerodynamic chord meam of minimi7;ing the pitching moments musod by high- LONGITUDINAL STABILITY AND CONTROL lift flaps, but more data on t h i s effect me needed bofore specific recommendtrtions can be made.
It was noted in reference 1 that a straight wing with a slight reflex amber and dihedral has all the necessary aero- dynamic characteristics for both longitudinal and lateral stability. A straight wing employing a tmihg-edge flap as a trimming control, however, suffers an undesirable loss in maximum lift, particularly if the static margin is large.
In order to improve this condition, the installation of leading- edge slats has been considered. This solution has found little favor, however, because of the accompanying increase in profile drag and the unusually high attitude required for landing with leading-edge slats. At the present time the most practicable method of overcoming the deficiency in mozjmum lift appears to be to incorporate sweepback (or sometimes meepfomard) into the wing. The majority of the contempormy problems in longitudinal stab%@ of tail- less airplanes arise from the adoption of t h i s solution.
EFFECTS OF SWEEP Advantages of sweep.-Sweepback gives the wing an effective “tail length” and is therefore especially adaptable for tailless airplanes. This tail length is proportional to the product of onehalf the span of the portion of the wing with sweep and the tangent of the meep angle; consequently, (1) high-lift flaps can be located at the center of the wing where their lift increments produce only minor changes in the pitching moment about the center of gravi@ of the a i r - plane, (2) flaps for longitudinal control can be located near the wing tips whare only minor changes in lift are needed to produce the requisite pitching moments for trim, and (3) more leeway is permitted in locating the center of gravity inasmuch as the aerodynamic center of the wing can be controlled by the angle of sweepback.
If only high lift is considered, the results of an investiga- tion relating to the use of various types of flap on swept-back AN INTEEUM REPORT O N T H E STABILITP AND CONTROL OF TAILLESS AIRPLANES It is known that, for trailing-edge flaps, an increase in flap chord shifts the centroid of the incremental flap load forward and thus causes a reduction in the mtio of the flap pitching-moment increment to the flap lift increment. This effect can be observed in figure 1 by comparing the results for different flap chords. At the present time, the optimum combination of flap size and flap hinge-line position for specifk designs must be determined by experiment.
The lift increments produced by flaps are governed also by the plan form of the bnsic wing design. The imporbat‘ factors are (1) the aspect ratio, (2) the taper ratio, and (3) the angle of sweep. O f particular interest for tailless airplanes is the so-called self-trhmhg flap, which is a flap arranged to produce zero pitching-moment increment about the aerodynamic center of the’wing. The effect of aspect ratio on the lift-coefficient increment produced by a self- trimming trailing-edge split flap on a swept-back wing is shown in @ y r e 2. The effect of taper.mtio on the liftrco&cient increment produced by a flap is discussed in reference 4 and nn indication of the offect to be expected can beobtained from figure 3. In general, a moderate taper ratio of the order of 2:l is recommended. The effect of sweepback on the lift increment produced by a self-trimming trailing-edge - ~ ~~ split flap on a swept-back wing is shown in figure 4. The 0 .2 .-% . 6 . 8 LO Tbow rafib, X data iU figures 2 to 4 were taken from an analytical inveati- P r a m 3.-Eff& oI taper m i i o on the inamnent I n U t mdaknt proanced by traIltng- gation of s d - t r h m h g trailing-edge split flaps (reference 2).
edge split f3ap having m o pituhing-moment fnumnent abont t h e aerodynamfc cantar of the a. A-XIo; A-7.3; ~ 0 s o s ‘ ~ p ~ .
Although trailing-edge split flaps have been found to bo particularly beneficial on swept-back mingS in producing forward wings, provided the high-lift flaps are placed on the high lift, it is cautioned that there me considerations other outer portion of the wing span and t h e flap for longitudinal than high lift involved in the selection of a flap for a specXc control is placed at the center of the wing.
design. For example, consideration of the minimum drag Disadvantage of sweep.-A most disagreeable ohmacter- of flaps for take-off, ground clearance, and the operation of istic of a swept-back xing is the inherent tendency to s t d l a pusher propeller in the flap wake may lead to the adoption prematurely at the tips, a phenomenon primarily associated of some other flap even at some sacrzce in lift.
with the lateral flow of the boundary layer. This chmc- Increases in maximum lift can be expected with swept- teristic is particularly undesirable bemuse it occurs first 0 8 /6 z+ 32 40 Aspect rotio,A Sweepback anyle,A, deg mawE 2-Efrmt d asp& rntlo on the inaement i n Ut coefeotent prod& by traIllng- FIGWE L-Efflet of sweepbaot on the incnwnent In U t cdW@nt produced by m a - edge WUt Onp hnvhg zem pltohbg-moment inmment nbont the aerodynamSo center of edge split flsp h a - zero pltahlng-moment inoranent a b u t the aarodynamfo canter or the wlng. S d J ; A-200; c r ” 0 a ; d-Boo.
the wing. s 4 2 ; A-7.3; epO2&; 8f-W.
REPOBT NO. 796-NATIONAL ADVISORY COMMITTEE FOR ~ O N A U T I c 8 over the rem portion of the wing where the control surfaces (1) Wing t&.-It has been proposed to wnsh outltho me located. The tip stall is manifested as a pronounced wing tips, that is, to lower the angle of attack of the section pitchkg and r o l l i n g instability accompanied by a tendency near the tip. Reference 5 shows that the m o u n t of wash- of the elevators or ailerons to float upward. A n example out required to benefit the tip stcllling characteristics is of the effect produced by the tip stall on the pitching moment sufEcient to incrense the drag of the wing seriously at low of a swept-back wing is given in figure 5. The rapid increase angles of attack. One method of avoiding the high drag in positive pitching-moment coefficient accompanying the is to have a portion of the wing tip rotatable in flight.
tip stall is characteristic.
The rotatable wing tips should be so proportioned with Swept-forward wings tend to stall first at the central part respect to the elevator that the airplane cannot bo stnlled of the wing. Center-section stalling causes pitching insta- until the tip angle has been sufEciently reduced to eliminate bility but the rolling instability associated with the tip stall the tip stall.
(2) Chunge in aij.foil sedion.-The initial stalling of the wing sections on the outer span of the wing can bo con- trolled somewhat by increasing the thickness or changing the camber of the airfoil sections used. The results of rofer- ence 5 indicate that this method can appreciably increnae the angle of stall of a wing without flaps or meepbnck, .I particularly if a change in camber is used in conjunction with w h g twist. The analysis in reference 6 does not con- sider the &ecb of sweep or flaps. Changing the wing sec- tions, however, generally has the disadvantnge of increasing the drag of the wing a t low angles of attack.
(3) Fld-plutt? separators.-It has been suggested that tho tip stall might be delayed by meam of vertical flat plates or h alined with the wing chord at about one-half tho clis- tance to the wing tip and extending around tho trailing edge of the wing and forward almost to the leading edge. Tho function of the plate is to prevent cross flow of the boundary layer by "separating" the fields of flow along the wing span.
Expariments on sweptrback wings with flatrplato separators installed have indicated that some increme in the anglo of stall can be obtained by this method alone but that gonernlly a new stall is induced just inboard of the plate itself. Bottor results might be obtained if the flat-plate separators are iisod in conjunction mith changes in wing plan form, particularly in the vicinity of fahe wing tip.
.I (4) Changes in plan form ut tip.-According to tests niado in the Langley free-flight tunnel, a -go in wing plan form at the tip alone has little effect on the tip stall (fig. 5), 119 evidenced by the instability manifested by the pitching- 0 2 .4 B .B J. 0 1.2 L i f t coefficient, C ' moment curves for a l l tip arrangements. It appecus from FIOURE L-Effect d chang in Up shape on the pltchlng-moment CharsCterlstIa of a m p t - associated tuft studies that flow sepamtion always occura -*. .
at the junction between the tip and the inboard portion of the wing. In any event, the change in plan form sliculd of swept-back wings does not occur. This advantage of extend inboard of the original stalled regions.
sweepfonvard, however, is partly offset by the difliculty created in obtaining adequate static balance on accountof (5) Leading-edge slats.-The use of tip slats haa been found to be the most dective method of delaying the tip s t d .
the forward shift in the aerodynamic center of the wing mused by sweepfonvard. With swept-fomard wings, the Leading-edge slats may increase the angle of s t d as much fuselage or load-carrying element must be placed ahead of as IOo if judiciously located. Tests of models in the Langloy the wing in order that the center of gravity may be ahead free-flight tunnel have indicated the necessity of oxtend- of the aerodynamic center. ing the slat at least over the portion of the wing dected by Remedies for tip stalling.-Before satisfactory flight be- the stall. It has been found that slnt spans of the ordor of havior can be assured, provision must be made for delaying 30 to 50 percent of the wing span are necesscrry to abolish or eliminating the tip stall. Various schemes have been completely the effects of the tip stall. Typical stalled areas proposed for delaying or eliminating the tip stall and a behind a swept-back wing with various slat mangemrnts number of such schemes are summarized as follows: are shown in figure 6.
-4X INTERIM HEPORT O N THE STABILITY A N D CONTROL OF TAILLESS AIRPLANES 423 . Effects of power.-Tho analysis of the effects of power on the longitudinal stability is somewhat simpler for tailless airplanes than for the conventional airplane on account of the absence of the horizontal tail. For convenience, tho effects are divided . $ t o three pnrts: .I (1) Effects associated with normal force and direct thrust of propelleis ' (2) Effects associated with slipstrecrm velocity and
J " domwmh behind propellem
-L- (3) Effects rtssociatcd with dynamic action of jets C The effect of the propeller normal force is small for the convontional arrangements of propellers nnd is u s u d y a firred factor for a given design. Methods of estimating the -c C effect are avaihble in reference 8.
u .I E As with conventional airplanes, the effect of the thrust
8 on stability is directly proportional to the product of tho
h thrust and the perpendicular distance from the center of
; *
gravity of tho airplane to the thrust line. T h i s effect is con-
P
trolled, of come, by the vertical location of the pivpeller and the inclination of the thrust line. The farther nbove the center of gravity tho thrust line pwes, the greater is the stabilizing effect produced by n given thrust; and the farther below the center of g ~ ~ ~ v i t g t h e thrust line passes, the . I grater is the instability produced by a given thrust. In any w e , the farther from the center of gravity the thrust line passes, the greater are the changes in trjm due to the thrust that accompany changes in power. This effect is illus- 2 .4 . 6 . 8 1.0 I. 2 trated in figure 7. The effects of power were small when the Lift coefticient, Cr.
thrust-line axis passed close to tlio center of gravitg of the FIQIJBE O.-EU& of slnt on tho pitching-moment chnrnrlmbllrn of a sncppt-bacg wing.
When the thrust line was 0.048C below tho center airplane.
If fixed slats me used, an undesirable increme in dmg may of gravitg, however, the stability decreased appreciably.
result at low mgles of attack. It may be possible, however, to build retractablo slats that have only minor effects on the ovor-all drag of the wing at low a n g l ~ of attack after more research and work on the development of retractable slats havo been done.
(0) Taper.-Pnrt of the stalling of swept-back wings can bo attributed to high taper. The use of highly tapered swopt-back wings should be avoided, therefore, inasmuch ns data on tapered wings indicate that the ben&cid effects of sweepback can be obtained with moderate taper ratios of tlic order of 2:1 (reference 5).
LONGITUDJNAL STABTLITY As with a conventional airplane, a tailless airplane is statically stablo if the center of gravity is ahead of the aerodynamic center. The pcsition of the aerodynamic cen- ter is appreciably affected by (1) the addition of a fuselnge or a streamlino nacelle, (2) sweepback, and (3) power. Thc extent of the fomard shift of the aerodynamic center pro- duced by a fuselage or nacelle has been discussed in reforcnco 0. The basic procedures for calculating the aero- dynamic center of wings of various plan f o i m are given i n reference 4. Applications of lifting-surface theory to iho doterminntion of the span loading of swept-back wings can bo found in reference 7. The effects of power on longitudinnl stability me discussed in the following pnmgraph.
FXWJBE 7.-FBect of power on the longitudinal stabillto d a puher-type tnllles nlrplana REPORT NO. 796-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS A t a lift coefficient of 0.8, the stcttic margin decreased from or slightly behind the latoral axiS through t h o center of 0.04 to 0.012 and the unbalanced pitching moment intro- ,gavity of the airplane.
For jet-propelled airplanes, the location and inclination of
duced by the thrust required about l o o of down elevator to
the jet axis exercises an effect on the stability characteristics trim the airplane.
of the airplane similm to the effectproduced by the thrust of The propeller slipstream is an important contributing item a propeller. At the present time, it appears that tho location to the longitudinal stability characeeristics of the a i r p l a n e of the jet exis should be governed by the same factors which particularly for tractor arrawgements. The controlling fac- were considered in the discussion conceding the location of tor is the location of the aerodynamic center of the portion the thrust line of conventionally powered airplanes.
of the immersed in the slipstream. I f the aerodynamic Damping in pitah.-As pointed out in reference 1, the low center of this portion o f the wing is behind the center of value of Cmg associated with tailless airplmes is no serious , g r a v i t y of the airplane, the slipstream produces a stabilizing disadvantage so f a r as the damping of the oscillations is effect; if the aerodynamic center of t h i s portion of the wing is ahcad of the center of gravity of the airplane, the slipstream concerned if the airplane has a positive static margin. It produces a destabilizing effect. D e s i g n parameters affecting appem that damping is introduced by the particular coupling the contribution of the propeller slipstrem are (1) the loca- of the modes of motion ns affected by the low vdue of Cm4 and by the reduced radius of gyration in pitch t 1 . 9 shown in reference 1 and figure 9. The results of tests in the Langley free-flight tunnel (reference 9) indicated that changes in tho rotational damping in pitch have little effect on the longi- tudinal steadiness for values of the static margin greater than 0.03.
tion of the section aerodynamic centers, (2) the s p a n ~ e location of tho propellers, and (3) the inclination of the pro- peller d s . The basic moment of the immersed whg sec- tions also has an effect. Figure 8 indicates the magnitude of some of the power effects to be expected.
For the tmctor-type tailless airplane shown in @ure 8, the thrust line passes near the center of gravity so that the effect of the thrust is negligible. The aerodynamic centers of the wing sections immersed in the slipstream me ahead of the center of gravity, however, and the slipstream therefore produces a destabilizing effect.
From consideration of changes in static margin nnd trim, it nppears desirable on t a i l l e s s airplanes of the pusher type to locate the thrust line close to thc center of gravity of the airplnne (:<O.Ol is recommended and, if feasible, to locate 0 .L
)
Radius of gyraf/on, ky, chord the propeller so that the aerodynamic centers of the wing FIQWE 9.-Effect of radios of gyration In pitch on the short--perIodl ongldtudlool c d l n t l o n sectinns affected bv the inflow to the propeller me either on of D taluess and a mnventlonnl nirplane. Btatlc margIn-O.07e ~ - 1 2 .
AN INTERIM REPORT O N THE STABILITY AND COhTTROL OF TAILLESS AIRPLANES It mils pointed out in reference 1 that the reduced dampkw in pitch of a tdess airplane might result in an uncontrollable motion of the airplane if the static margin is allowed to become negative. This contention has been supported by subsequent tests in the Langley free-flight tunnel (reference 9). The tests indicated that a serious form of instability may develop when the static margin of a t d e s s airplane becomes negative. As a result of this danger of uncontrollable motionsmithnegative static margins, it is recommended that the center of gravity of a tailless airplane never be permitted, under any conditions, to reach a position behind the aerodynamic center.
Tumbling.-A form of dynamic instability of tailless air- planes may be m d e s t e d as tumbling. Tumbling consists of a continuous pitching rotation about the lateral axis of the airplane. The maneuver is extremely violent and imposes severe accelerations on parts of the airplane.
So f a r a8 is h o r n , there me no authenticated instances of the occurrence of tumbling in flight. Models of tailless airplanes have been made to tumble in the Langley 20-foot free-spinning tunnel, however, by forcing the model to simulate a whip stall. At the present time, however, little is known about the mechanics of the tumbling motion.
Tests conducted in the Langley 20-foot free-spbning tunnel have shorn that the position of the center of gravity has a pronounced effect on the motion. It appears that provision of a large static margin prevents tumbling but that a stable tumbling condition may exist if the static margin is slight.
Tests have shown also that once the tumbling motion has started the normal flying c o n h l s me relatively ineffective for recovery from this stable tumbling condition.
In view of the severity of the tumbling maneuvers, it is recommended that tumbling tests be required of models of all fighter t d e s s airplanes.
LONGITUDINAL CONTROL One of the dBcult problems in the design of tailless nir- be deflected considerably more than that of a conventional planes is the provision of adequate longitudinal control. The airplane in order to produce the same changes in trim lift type of longitudinal conk01 usually employed consists of an codicient in flight. The elevator on tailless airplanes, being elevator (or flap) placed at the trailing edge of the wing.
an integral part of the wing, must also operate at d l angles With this type of control, the loss in lift caused by the flap of attack of the wing up to the stall. The elevator must deflection required to trim the airplane can be appreciable, therefore be balanced over large range of angle of attack particularly for a tailless airplane with a large static margin.
and deflection.
The computed loss in lift that results from t;rimming the In order that push forces may be required to increase the airplane at various values of static margin is shown in *lane speed ( h m trim speed) and that pull forces may figure 10. It is evident from figure 10 that the loss in lift be required to reduce the airplane speed, the inherent u p caused by the longitudinal control can be minimized by floating tendencies of the elevator with increasing angle of placing the control surfaces a t the tips of highly swept-back attack must be reduced. The critical case for stick-force wings of high napect ratio. When the longitudinal control is placed near the wing tips, the elevator can be combined reversal (called elevator snatch) is that for neutral longi- with the aileron in m mangementi to be discussed later in t u d i n a l stability (or zero static margin). If there is to be the section entitled “Aileron Control.” no stick-force reversal for this case, the variation of the Design requirement.-It is to be expected that the elevator elevator hinge moment with angle of attack must be zero stick-force requirements for tailless airplanes should be the or positive at all angles of attack throughout the Bight same as for conventional airplanes of the same class. The range. When t h i s condition is W e d , the elevator either balance requirements for tailless airplanes, however, me remains stationary or floats down (ts the angle of attack is more severe than for conventional airplanes. For the same increased. Further discussion of this point may be found static margin, the elevator of a tailless airplane usually must in reference 10.
HEPORT NO. 7 9 6-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS FIOUBE ll.-VnrfntIoo of d o n h h p m o m e n t mfedent wlth angle of attack for 0.1& elevators on a mdfied NACA 65,3418 airfoil. &.-e 9 4 9 . 4 pmm per s q m root.
Types of control.-A plain flap is unsuitable for use as an &Ted aerodynamic center of the wing moves forward with an clovator on a tailless airplane mainly because it floats upward increase in trailbg-edge angle, and the stick-free aerodynamic m tho angle of attack of the wing is increased. In figure 11, center of the wing moves backward with an increase in the upfloating tendency of the flap is manifested by the in- trailing-edge angle.
creasingly negative flap lunge moments that are developed (2) Special tlenting.-It hna been suggested that an internal aa tho angle of attack is incrensed. Various balancing balance be used which hns a vent near tho airfoil leading edge.
schomcs havo been proposed for reducing or eliminating Analysis of available data indicates that, a t large angles of the upfloating tendency of plain flaps but no aerodynamic attack, however, this aiinngement would have the same un- balances are yet known that completely satisfy the design favorable characteristics as the internal-balance arrange- requirements. Several balance arrangements, however, show menta vented at the hinge line.
promise of being satisfactory in two-dimensional testa but An analysis of pressure-distribution data indicates that m havo received no esperimentd verification in three- internal balance vented near the trailing edge of the airfoil dimonsionnl tests. A few of the proposals are discussed in mould give the desired hinge-moment variation w i t h angle of tho followi~g parngraphs: nttack. The fact that the pressure changes in this region (1) Bmels.-Figure 11 presents the variation of elevator of the airfoil are small, however, appears to demand an section hinge-moment coacient with q l e of attnck a t internal balance of such length n . q to be impracticable.
zero elovator deflection for straight-aide and beveled elevt+ (3) Slots M o f elemtors.-A.s the upfloating tendency tois with and without internal balnnco vented at the hinge inherent in a l l control surfaces a t large angles of attack is line. The curves indicate that the desired hingemoment caused by air-flow separation over the control surfaces, it variation with angle of attack cannot be obtained with these has been proposed that slots be placed in the wing ahead of nrrangoments of bevel and internal balance. Since the the elevator as one means of suppressing this effect. Vorg slopes for all elevators me nearly p a d e l at large angles of little research has been done on this particulm scheme how- attack, it is not to be expected that favorable curves can be ever and, a t the present time, all that can be said is that it obtained either by further incrensing the trailing-edge angle might be advantageous.
or by incrensing the length of tho internal balance vented (4) Azctomelicdy controlled tabs.-Several rather mechan- at tho hinge line.
ically complex types of balance have been proposed to pre- Beveled elovators also deet the location of the wing aero- vent elevator-force reversals. Because a tab is normally n dynamic center. The magnitude of the effect depends on powerful means of charging elovator hinge moments, it has tho chord and spnn of tho elevator. In general, the stick- JIN l 3 T E E U X FLIWORT O N THE STABILITY LND CONTROL OF TAILLESS AJXPLBNES Inasmuch as the spoiler may be located ahead of an aileron, been1proposed to place a tab on tho elevator and cause the an upward deflection of the spoiler mould cause the aileron tab to deflect upward in such a manner that the elevator to have an upfloating tendency and a t the same time muse floats down when the angle of attack is increased. The the ailerons to be less nearly balanced.
deflection of the tab would be controlled either by linking Control for take-off .-Under take-off conditions, the longi- it to an internal balance, suitably vented, or by linhing it tudinal control, besides supplying a pitching moment large to a free-floating spmnmise portion of the elevator called a enough to trim the wing at the lift codicient corresponding flipper. The ilipper should be located along the span in a to the ground &e of the airplane, may be required to supply region where the stdl is first manifested over the control of several such the “additional” pitdung moment necessary to counteract surface. Two-dimensional characteristics flipper-tab arrangements have been computed from section (1) the pitching moment of the weight of the airplane about data, and the results are presented in figure 12. Some of the point of c o n k t with the ground, (2) the pitching moment created by the friction force on the wheels, and (3) pitching the configurations result in hinge-moment slopes that are I f similar moments arising from interference caused by the proximity either zero or positive at dl angles of attack.
characteristics could be obtnined in three-dimensional flow, of the airplane to the ground (references 11 and 12). In no stick-force reversal would occur for these combinations. order to make certain that the airplane has adequate longi- tudinal control to compensate for these additional pitching The stick force could be controlled by adapting a spring moments arising during the take-off, the Army requirements either to the same tab or to an auxiliary tab.
for an airplane equipped with tb tricycle landing gem state ( 6 ) Spoilers.-The possibility of using a spoiler as an that the longitudinal control shall be powerful enough to elevator has been suggested as a means of avoiding stick- pull the nose wheel off the ground at 80 percent of the take- force reversals. The loss in lift accompanyingthe production off speed during operation off terrain where the co&cient of of a given pitching moment is greater with the spoiler con-
friction is 1 /lo (reference 13). An idea of the magnitude of
trol, however, than w i t h the elevator control. Unpublished the “additional” pitching moment that the longitudinal tests of reamardly located spoilers on b o different models control must supply to compensate for the extraneous effects confirm tho fact that spoiler projections of less than 0.01~ associated with take-off may be obtained from fiewe 13.
produce negligible changes in lift. Such a spoiler is undesir- Because of the short moment arm associated with the able for longitudind control because a small stick movement elevator of a tailless airplane, it is extremely difhult to produces no change in trim, whercas a larger movement of the stick may produce large changes in trim and normal ac- celoration. The charsrtcteristica of spoilers c(u1 be controlled somowhat by adjusting the spoiler span and by incorporating spccial von ting to t,he spoiler.
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----
Seciicm M Y / ~ of oftack tu,, deg P I ~ U R E 12.-VnrIatlon of CA, mlth PO at &=Oo for vnrious fltpper-tab arrangements on a bornled elevator. Gectlon data.
REPOHT NO. 7 9 6-NATIONAL ADVISORY C0MMI"EE FOR AERONAUTICS design an elevator that can done supply the pitching peller force. If, in addition, the pusher propeller is moiintcd moments necessary to meet the &my take-off requirements; behind a vertical tail surface, nn ndditional increment in for example, point A spotted on figure 13(c) was computed directiond stability is realized from the vertical tail surfnco for a typical t a i l l e s s airplane. For this case, a pitching- because of the effect8 produced by the inflow of air into tlie moment coe5cient of -0.335 is needed to raise the nose propeller.
wheel off the ground. The elevator effectiveness Cma, for The destabilizing effect of a fuselnge or streamlino nacelle on the directional stability hns been discussed in reforence 1.
this airplme is only -0.003 per degree, and thus the elevator The destabilizing effect of the fuselage and nacelle of tailless cannot rake the nose wheel for takeoff. In order to remedy airplanes is usually a t least ns great n s the stabilizing effects t h i s situation, it has been proposed to utilize the nose wheel contributed by the + e done. It is therefore necessary on ns a jack to adjust the ground angle during the take-off run.
t a i l l e s s airplanes to provide some method of supplying If some scheme of t h i s type is not provided, it appears directionnl stability.
likely that tailless airplanes may experience diiliculty in The provision of adequate directional &ability for tnilless mising the nose wheel off the ground at tske-off if the landing- airplanes is more difficult thnn for conventionnl airplnnes g a r angle e is l q e , particulmly with large static margins.
bemuse of the short longitudinal moment a m . A variety Center-of-gravity range.-On the basis of the l o n g i t u h d stability and cont-rol problems which have been discussed, it .m r ~ p p m that the permissible range of center-of-gravity position compatible with satisfactory flight behavior is more critical for tailless airplanes than for conventional airplanes.
.002 I f the static margin becomes negative, there is danger of b encountering longitudinal instability either tis a divergence % from straight flight or as tumbling. I f the static mmgin is fis factory flighf t .m/ 0 aracte ridlcs too great, the elevator control may not be powerful enough R to raise the nose wheel off the ground at takeoff. Further- more, ;f the static margin is large, the elevator deflection required to trim the airplane in level fright may seriously impair the e5ciency of the wing with a consequent loss in performmce of the airplane. At the present time, a range of ultimate static margin from 0.02 to 0.08 appears to be FICWBE 1 4 . - E f f & o I ~ ~ stability C . and dihedral oflmt CiP on flight ohnraclcrlstlar rensonnble for tailless airplanes.
BJ dotermlned by tab o f a model in Langley fm-fflght ttmncl. CL-1.0.
LATERAL STABILITY AND CONTROL of fin arrangements and end plates of the type discussed in DIRECTIONAL STABILITY reference 1 has been teated on models of tailless airplanes iu the L a n g l e y free-flight tunnel in an effort to improvo tho Since the publication of reference 1, several models of t a i l - directional stability of specific models. A rbum6 of somo of lesa airplnnes have been tested in the Ln.ngley free-flight the more pertinent considerations that have evolved froin tunnel. It has been v d e d from these teats that the these tests is given i n the subsequent discussions.
m o u n t of directional stability possessed by tailless CLirplmes (1) Fins.-It hns been found that, for a tailloas airplnne should be ns great n s required on conventional airplanes i f having a straight wing, adequate directional stability can be the same requirements regarding s&igWoryfE7bins quaiah provided by vertical tail surfaces located at the centor section are to be adhered to. The value of the directional-stability of the wing near the t r d m g edge (or on the fuselage if oiie parameter CnP, recommended for conventional airplanes, is is available). The size and number of vertical tails nocessnry usually greater than 0.001 per degree. As evidenced from for a specific design of course depends primarily on tlie figure 14, however, models have been flown in theLangley degree of directionnl stability required. Whon multiple freeflight tunnel and with a value of C? of only one-third tails are used, it appears to be preferable to use ns few tnils this amount dthough the best flying qualities of these models of ns high aspect ratio as possible because (n) h of high were obtained with values of C,,$ in oscess of 0.001.
aspect ratio are more effective than h s of low aspect rutio, The inherent aerodynnmic characteristics of the wing done (b) the interference effects between ndjacent vertical fins nre have sometimes been tried n s the source for directional minimized, and (c) much of the fin is outside tho relatively stability. The amount of stability contributed by the wing thick boundary layer on t.he uppor rear surface of tho wing.
depends on the wing plan form and the lift coef6cient. The If the tailless airplane has a swept-back wing, the usutil effect of the wing plan form does not appear large but more practice is to place the vertical tail surfaces at the tips rathcr are needed on this subject. The directional stability of data than a t tho center section in order to tnke advantage of thc the w i n g alone increases somewhat with lift coefficient. The longer moment arm available. 'Then verticnl fins n r o placed directional stability at lorn angles of attack for the wing alone a t the wing tip extremities, however, the moment rum nmo- hns g e n d y been found to be inadequate although adequate ciated with the drag of the tip fin is so large (one-half tlie stability mny sometimes exist at high anglcs of attack.
span) that the drag charnc.teristica as well na the lift chnrnc- usually contributes a mall degree of A pusher propeller teristica of the tip iins exert an influence on the diroctionnl directional stability because of the stabilizing normal pro- stability. The relative contribution of the lift and drag of AN INTERIM REPORT ON T H E STABILITP BND CONTROL OF TAILLESS AIRPLANES V e r k d toil5 tested M a tailless-airme e/ A
Tail & 'a
T'wo wt7g-tip fails Twu - 4 ~ toll.
foed-cuf 5 ' foedin ! Y Area 5 I O fp- 3) & = d m f r O 4 Q 5 0.000/0 t ; , a00012 the tip fins to the directional stability of the airplane of fins are set at the correct angle of t0e-h for zero lift in course depends on their inherent aerodynamic characteristics. straight flight. When the airplane sideslips, the angle of Some attention must accordingly be devoted t o setting the attack of the leading tip i k is made more negative and thus initial angle of the tip fins. causes a large increase in the proHe-drag coe5cient due to If directional stability is to be obtained with tip fins of flow separation; whereas, at the same time, the angle of low aspect ratio (less than about 2), the tip fin must be set attack of the t r - tip fin is increased positively and thus causes orJ$ a relatively 9Ill&u increase in ita profiledrag with some initial toe-in because of the large induced drag When coef6cient. Drag fins of this type have not been tested in aasociated with lifting surfaces of low nspect ratio.
flight. A lateral oscillation may possibly develop as a result the airplane is yawed, the stabilizing moments generated by tip h are produced by the large induced drag of the of drag hysteresis, although such an efTect has not been forward wing tip. If, on the other hand, directional stability observed in tests of small-scale models.
is to be obtained with tip fins of moderate or high aspect The most effeotive tip fins tested in the Langley fiee-flight ratio, the tip fins must be set with some initid toe-out.
tunnel have been based on the pr0He-cLm.g principle. Tip With toed-out tip fins, the stabilizing moments are gener- h based on induced-drag principles have been somewhat ated by the outwardly directed lift as explained in reference 1. less effective. The tip fins based on lift principles have been The stalling characteristics of the tip f i n s , moreover, are the least effective tested bemuse of the short moment arm nn important design consideration. When an airplane associated with the lift tip fins. The moment m, however, with toed-out tip fins is yawed to an angle sufEcient to stall is controlled by the angle of sweep so that, for wings with a the rear tip f b , a large destabilizing moment is generated by large amount of meepback, it may be feasible to design an the increased drag of the rem tip fin. On the other hand, effective lift tip f i n . Central fins have generally been satis- when an airplane with toed-in tip fins is yawed to an angle factory, particularly if mounted on the end of a fuselage.
sufficient to stall the forward tip f i n , a large stabjlizing (2) Turiwkbwn wing tips.-The amount of inherent moment is produced. The manner in which the stalling of directional stability possessed by a wing may be incrensed toed-in and toed-out tip Gns d e c t a the directional stability by turning down the wing tips; thus, i n efFect, the wing tips of the airplane is illustrated in figure 15. are made to function somewhat as lower-surface tip fins and the increased directional stability is manifested through the It has been suggested that the effectiveness of drag tip outward lift developed on the ming tips. The incorporation fins can be augmented by employing an airfoil section posses- of positive dihedral angle on the wing, however, results in a sing aerodynamic characteristics similar to those s h o r n in decrease in directional stability because the lift of the wing figure 16 for the NACA 4306 airfoil. In practice, the tip REPORT NO. 798-NATTONAL ADVISORY COMMITTEEI FOR AERONAUTICS -8 -4 0 4 8
w of a m “,&9
FIoWE l O . - A u o d ~ a m f ~ &~II@&XWCS d NAOA 4 W afrIon SeCtlOn.
itself is directed hward rather than outward (fig. 17). The destabilizing influence of a positive dihedral angle must be taken into account in computing the directiond stability required of the wing tips. An examination of figure 17 indicates that the effects of the dihedral can p r a c t i d y nullify the effects of the turned-dom ming tips. Turned- present time, the solution to the problem of creating ade- down wing tips are believed to be less satisfactory for securing directional stability than iins of the types previously quate directional control rests primarily in reducing the yawing moment that such a control must overcome; thus, discussed.
(3) Azctomalic control.-It has been suggested that a tnil- it is of particular advantage on tailless airplanes to locate lass airplane of very low directional stability with hsd con- the propellers as close as possible to the center line and to trols could be flown satisfactorily if an automatic pilot were provide ailerons that create favorable yawing moments geared t o the directional control in such a manner that when when deflected.
the airplane sideslipped the amount of directional control The provision of adequate directional control on a t a i l l e s s supplied would be sufficient to incrense the effective value of airplane with rudders based on lift principles is difEcult.
because of the small moment arm available for control.
URb. Referehce 1 includss the suggeation that the directional control could bc linked with the aileron control in order to Computations have indicated khat rudders based on lift minimize the effects of adverse aileron yaw. It is believed principles alone generally are not able to counteract the yaw- that satisfactory flight behavior could be obtained with such ing moments generated by severe asymmetric thrust condi- tions even if mounted a t the tip of a swept-back wing.
automatic-stabilizing schemes although, at the present time, Lift rudders must also develop an appreciable side force no flight investigations of such applications have been became of the short moment arm. In order t o compensate reported.
for this side force, the tailless airplane must be sideslipped or DIRECTIONAL CONTROL banked an appreciable amount because of its low lateral The requirements of rudder control for tailless airplanes resistance. Some of the fight ditliculties that may mise as a r o essentially the same as for conventional airplanes. a result of these circumstmces are discussed in reference 14.
Rudder control is necessary to counteract the adverse yaw Some use has been made, therefore, of directional control occurring during rolling maneuvw and to provide sufficient that is dependent upon drag characteristics because of the directional control to trim tho airplane directiondy at large moment arm which can be obtained by locating the operation under asymmetric power conditions. At the drag directional control at the wing tip.
AN INTERIX REPORT ON THE STABILITY AND CONTROL OF TAILLESS AIRPLANES 43 1 It appeals possible to design a rudder based on drng prin- on a plain wing with no sweepback. It is noted that for a ciples utilizing a double split flap (brake flap) that could trim wing with no sweepback C?,@ is practicdly independent of the yawing moments mused by asymmetric thrust condi- lift co&cient .
It is cautioned, however, that split-flap tions (fig. 18).
EfFect of sweepback.-Systematic investigations to de- rudders may generate undesirable rolling moments along t e r m i n e the effect of sweepback and taper on C, m e being vith the y&&g moments produced. This type of rudder conducted. The limited data available at the present time if the drng may also affect the performance of the airplane indicate that the effective dihedral of n swept-back wing increments necessary for control are very large. At the increases with angle of attack; it is thus advisable to use a prescnt time, specific designs of rudders of this type should geometric dihedral angle of about Oo in order that, at the be developed e.rperimentally.
higher lift co&cients, the effective dihedral does not exceed The use of propellers mounted in the wing tips has been 3 O or 4 O . The incrense in Crp with mgle of attack for a proposcd ns a method for supplying directional stability and swept-back wing is not so detrimental ns might h t bo control. Such a system could, of course, be used easily supposed, however, because of the accompanying increase in wmthercock stability. an empirical formula for e s t i - mating the effect of sweep on is discussed in reference 1s.
Effeot of sweepforward.--The effective dihedral of a swept-fornard k g decrensses as the angle of attack is at alOb from wng center - increased. Some idea of the magnitude of the effect to be /me;r)=0.80; S-550;$ ~ 4 0 ; expected is given in reference 18. There is an indication also that the weathercock stability of a swept-fomard wing may decrease with incrense in angle of attnck. This effect would make the attainment of lateral stability over a large range of angle of attack dii%cult. More information on swept-forward wings is needed, however, in order to evaluate these effects.
AILERON CONTBOL The aileron control of a tailless airplane presents no prob- lems greatly different from those for conventional airplanes.
An effort should be mnde, however, to avoid adverse aileron yawing momenta, particularly if the directional stability is low, in order to m i n i m i z e the sideslip developed during rolling maneuvers. Adverse d e r o n yawbg moments can be minimized by uprigging both ailerons or by utilizing rotn- table wing tips of t.he type previously described. In order to overcome the effects of dverse aileron yaw, it may be of advnntnge to employ a spring connection between the nileron and rudder control in a manner described in the section entitled "Tacticnl maneuvers." It is desirnble also that no pitching moments be produced by &e deflection of the PIOWFIE 18.-Ynwlng moments m t c d by double split perforated drag mdden mounted ailerons because the ailerom have n m l y the same moment m vlng Icadsdlng, pounds per aunre foot.
at wing Ups. c p O Z O c q, propalsIw el8doncy; arm ns the elevators. It is necessnry therefore to use ail- erons with an equnl up and down deflection.
with an automatic pilot. It is believed, however, that Spoiler control.-The use of spoilers for ailerons on t d e s s structural considerations may make such nn arrangement airplanes has been advocated from time to time. If only imprncticable at the present time.
upgoing spoiler projections me used, the pitching moments DIHEDEAL developed me prohibitive. A spoiler m ~ ~ e r p e n t employ- ing equal up and d o m projections wodd improve this con- Tho requirements of dihedral for stability are essentially dition but the data available me insufEcient for evaluating the same for a tailless airplane as for a conventional airplane.
conclusively the merits of such a system.
Computations of the type presented in references 16 and 16 and investigations conducted in the Langley free-fight tunnel Elevon control.--For some tnilless airplanes utilizing a (fig. 14 and reference 17) have indicated that, in the interest swept-back wing, ailerons placed nenr the wing tips have of lateral control and steadiness in gusty air, it is desirable been macle to act nlso as elevators because the most effective to keep tho effective dihedral angle s m d . The results of position for both controls is near the wing tips and bemuse these investigations have indicated that, for satisfactory larger-span lift flaps can be employed if the two controls nre lateral stability, the effective dihedral angle should not exceed combined. Such an arrangement, d e d elevons, combines a value corresponding to - C,=O.OOl per degree. T h i s value the design reauirements of both aileron and elevator in one control and int.roduces additional problems.
of G,B corresponds t o a geometric dihedral angle of about 5' R E P O R T NO. 796-NATIONAL ADVISORY COMMI!lTEX FOR AEFtONAUTICG Trimming-tab operation of devons difiers from that for The total effective deflection range for an elevon must be nilerons nlone in that the tab must trim the hinge momont of the s u m of the ranges required for the aileron and elevator.
each elevon to zero when it is desired to tzim the nirplano in The fact that the neutral position of the elevon may be at roll in order to prevent the development of elevator stick some upward deflection when it is functioning as an aileron cnn be utilized to a certain extent in reducing the aileron forces. For ailerons alone, it is essential only thnt tho tab stickforcea. With a large static margin, however, the f u l l cwse one aileron hinge moment to bdnnce that of tho other deron deflection used with the large upward elevator aileron.
Section data from unpublished tests of an internally bnl- deflection required at low speed may produce large pitching anced, beveled, 0.18~ elevon with 0 . 2 6 ~ ~ tab indicnte that for moments and small rolling moments because the upgoing angles of attack up t o the s t a l l a full-olevon-span tab elevon may stnll. In order to improve this condition in deflected f20" could him to zero the hinge moment of nn some designs, the use of an auxiliary longitudinal trimmhg The pitch flap elevon deflected f25". The snme data, however, indicate device called a pitch flap has been proposed.
is located o u t b o d of the aileron. With such a device, the that little if my additional rolling moment can be produced by deflecting the elevon upward beyond 26" at luge angles of lateral control could be obtnined at low speeds by supplying attack.
most of the trim with the pitch flaps and thereby minimizing the upward deflection of the elevons. The elevons then DYNAMIC STABILITY would be deflected ns ailerons over a grenter linear range of Damping in yawing.-For t d e s s airplanes, the rotatioDal the curve of rolling moment against deflection.
damping is invnriably low on account of the reduction of tho The conditions regulating the balance of an elevon for a t a i l length. A comparison of the measured dnmping- typical l q e tdess ctirplane are indicated in fgwe 19. The moment coacient due to yawing a t alift coefEcient of 0.00 for ranges of vnlues of Ckd and Cam that satisfy the stipulated various tailless airplanes and n conventional airplane is givcn in figure 20. The valueswere obtained by the free-oscillation elevator and aileron requirements independently were method described in reference 20. The portion of e,,, evduated by the methods given in references 10 and 19.
The crosshatched region includes d l values of C,, and Ch, contributed by tho wings can be estimated from tho data jn reference 21. It was pointed out in reference 1, however, that satisfy simultaneously the stipulated elevator and that within the usud limits of dihedral and directional st.n- d e r o n requirements. The elevon must be balanced over a bility the damping of the lateral oscillntions is genornlly much larger deflection range than either the elevator or greater than would be indicated from only tho dnmping duo nileron done nnd, because of the increased deflection range to yawing velocity. Subsequent experience in flying tailless required, greater physical limitations are imposed concern- models in the Langley free-flight tunnel hns substantintccl this ing the length of the internal bnlance that can be used. The statement, and it appears that t;he small values of tho dnmp- considerations that have alrendy been discussed in regard ing panmeter Cnr associated with tailless airplnnes will not to controlling the upfloating tendency of the elevator with angle of attack also apply to the elevons.
be excessively detrimental to the flying qualit.ics provided the directional stability of the airplnne is adequate. Tho damping of the lateral oscillations is likely to be criticnl in the high-speed conditions because both Csr and the coupling .mu2 between the yawing m d rolling motion tend to diminish at the low angles of attnck.
On account of the low values of C . , associated with tailloss .awl airplanes, some apprehension hns existed concerning tho large angles of sideslip that may be developed whon the airplane is subjected to a disturbnnce of the type produced by asymmetric loss of thrust. There appenr to be no d a h pertaining to the direct effect of Cxr on a sideslipping motion of this type. The experience acquired in flying tailloss- airplane models in the Langley free-flight tunnel has in- dicated that the effect of (2% is probably secondmy to other parameters. The results presented in reforence 16 indicate that the maximum amplitude of the sideslip oscillation is influenced markedly by the rolling moment due to tho sideslip CI, and particularly by the yarning moment due to 70003 the sideslip CxB. Incrensing either the directional stability or the dihedral reduces the magnitude of the sideslip goner- ated by a yarning moment but the greatest reduction in to result from incrensing the diroctionnl sideslip appears st ability .
AN INTERIM REPORT ON THE STABILITP AND CONTROL OF TAILLESS AIRPLANES recovery. The moments produced by trailing-edge drag rudders in the stalled range of angle of attack may be con- siderably different from those in the unstded range. Some types of drag rudder have been found to produce apprecin- ble pro-spin rolling moments when applied against the spin Typic41 crznwtima! oirplanc and therefore are not effective for recovery. It is recom- mended, therefore, that the aerodynamic characteristics in yaw for different rudder deflections of tailless-airplane designs that have drag rudders be obtained at angles of attack beyond the s t d l if the possibility of a spin appears likely. The results of these tests would facilitate the evaluation of the relative merits of alternative rudder designs. For a complete investigation of the recovery characteristics, spin tests of the model are usually required.
Tactical maneuvers.-The suitability of tailless airplanes Tail toe-in C , & anqle, deg for performing tactical maneuvm of the nature required for 5 a00031 -.ffl4 formation flying, bombing, and aerial combat hns been the IO BO033 -.Of8 Tailless all-win y airplane wifb subject of hequent discussion. From considerations pre- 15 .00045 -.025 atos tip tails viously discussed, it appears that adequate directionnl stability is a necessary requirement for steadiness nnd enso of control. The fact that the lateral resistance nssociated with tnillw airplanes is low mar preclude the possibility of m&- ing flat turns with the rudder alone. At the present time, however, little information is available concarning the influ- ence of side mea on the lateral flying qualities of tailless nir- T d e 6 s all-ning a i r p l m s kjfb planes. More research is needed on this subject, pmticulmly OIOS ceder tails in regard to the efFectsproduced by the different directional- control devices mentioned in this paper.
n
The argument has been advanced that a pilot flying a fighter tailless airplane d l experience difficulty in keeping his gunsight alined with the met. It is believed however if the tailless airplane possesses the same directional that, stability and dihedral charmteristics as are demanded for conventional airplanes, the controlled motions during the normal accelerated maneuvers should not m e r appreciably from those of the conventional airplane.
In view of the likelihood that the successful tailless- Spinning.-Tests conducted in the Langley 20-foot free- airplane design may yet have lower directional stability than spinning tunnel have indicated that the steady-spin charac- conventional airplanes, the &ect of adverse aileron yaw on teristics of tailless airplanes are essentially the same as for the pilot’s aim may be more pronounced and in such eases n conventional airplanes. T h e control manipulations re- spring connection between the aileron and a trimming tab on quired for recovery from a steady spin, however, have been the rudder may be necessary in order to satisfy the following found to depend on the type and location of the control criterion : surface employed.
For tnilless airplnnes that have a vertical t a i l mounted at Ga 0.8 o m , - > - a < - the rear of a fuselage, the application of rudder control G!, Qg Q , would probably affect the spin in a manner similar to that for a conventional airplane because the vertical tail is not Such an arrangement should improve the stmdiness of &=ht.
bladcoted by the wing. I f the vertical tail is located on the rear upper surfnce of the wing, however, the rudder control GENERAL CONSIDERATIONS OF TAILLESS A N D is likely to be ineffective because of the blanketing effect of CONVENTIONAL AIRPLANES the wing.
In recent years opinion hns been divided as regards the For tnilless airplanes that have vertical tails at the wing relative adaptability of tailless and conventional airplanea tips, the application of rudder control would probably be for both fighter and bomber airplanes as evidenced by the effective for spin recovery, particularly if the rudder extends For tailless-airplane desi,- without a variety of designs that have appeared. Some observations below tho wing.
concerning the reI~tive merits of tailless and conventional fmclnge, spin recovery hns been found to be expedited by deaigns are offered here h m consideration of the stability application of rolling moments against the spin. The ai- and control problems that have been discussed.
lcrons therefore should be moved against the spin for best REPORT NO. 7 9 6-NATIONAL ADVISORY C O M M I T I X E FOR AERONAUTICS 2. Pitkin, Marvin, and Maggin, Bernard: Analysis of Faotors Small airplanes.-On m i m t of the thin ning sections AfFecting Net Lift Increment Attainable with Trailing-Edgo required for high speed, the volume enclosed by the wings Split Flaps on Tailless Airplanes. NACA ARR No. IA118, of a small rhplme is not lmge enough to carry nll the load; 1944.
consequently, it i s necessaq on smnll airplanes of either the Dent, M. M., and Curtii, M. F.: A Method of &timating the 3.
tailless or conventional type to incorporate a fuselage or Effect of Flaps on Pitching Moment and Lift on Taillw Air- craft. Rep. No. Aero 1861, British R. A. E., Sept. 1043.
some other lond-carrying element. It appears also that n 4 . Pearson, Henry A., and Anderson, Raymond F.: Calculation of vertical t d is necessary for directional stability. The differ- the Aerodynamic Characteristics of Tapered Wings with Pnrtial- ence between a small tnillsss airplane nnd a small conven- Span Flaps. NACA Rep. No. 666, 1939.
tional airplane, therefore, is essentially due to the suppres- 5. Anderson, Raymond F.: A Comparison of Several Taperod Wings sion of a horizontal tail as a meam of obtaining longitudinal Designed to Avoid Tip Stalling. NACA T N No. 713, 1030.
6. Multhopp, H.: Aercdynamica of the Fuselage. NACA T M No.
stability and control. I f the conventional airplane were 1036, 1942.
permitted a reduction in maximum lift comparable with 7. &hen, Doris: Theoretical Distribution of Load over a Swopt- that tolerated on t d w airplmes, the tail size could be Back Wing. NACA ARR, Oct. 1942.
reduced considerably. With the small horizontal tail then 8. Ribner, Herbert S.: Formulas for Propellers in Yaw and Charta of allowable, the conventional airplane might have a perform- the Side-Force Derivative. NACA ARR No. 3E19, 1043.
9. Campbell, John P., and Paulaon, John TV.: The Effects of Static anc0 corhpmable with that usually claimed for tailless &- Margin and Rotational Damping in Pitoh on the Longitudlnai planes without the restrictions attached to the longitudinal Stability Characteristica of an Airplane as Determincd by control.
Tests of a Model in the NACA Free-Flight Tunnel. NACA Large airplanes.-For large airplanes having spans of 150 ARR No. L4F02, 1944.
to 500 feet, the volume of the wing done may be suf6cient 10. Greenberg, Harry, and S t s d e l d , Leonard: A Thooretical Invosti- of Airplanes with Frco Controls gation of Longitudinal Stability to enclose bulk or weight of an appreciable magnitude even Including Effect of Friction in Control System. NACA Rap.
with the thin wing sections required for high speed. There No. 791, 1944.
is little renson to suspect that conventional airplanes of 11. Recant, Isidore G.: Wind-Tunnel Investigation of Ground EffocL equal span mill have any less wing space available for cargo on W i n g s with Flaps. NACA T N No. 706, 1939.
purposes than t d e s s airplanes. It appem, therefore, that 12. Tad, Itiro, Taima, hfasuo, and Gimitu, Sodi: Tho Effoct of Ground on the Aerodynamic Characteristics of a Monoplnno the suppression of. the fuselage LLS a loadcctrrying element is 'Xing. Rep. No. 166 (vol. X U , 2), Aero. Res. Inst., Tokyo primarily a matter of airplane size rather than of type.
Imperial Univ., Sept.. 1937.
In spite of the suppression of the fuselage, however, a 13. Anon.: Handbook of Instructions for Airplane Dosignem. Vol. I, vertical tail may be necessary on any large airplane, pmtic- Materiel Div., Army Air Corps, 8th ed., July 1, 1930.
14. Hartman, Edwin P.: Wind-Tunnel Testa of a %Engine Airplano ulmly on bombers, if optimum directional stability and con- Model as a Preliminary Study of Flight Conditions Arising on trol m e to be obtained. Some method must also be provided the Failure of One Engine. NACA T N No. 646, 1038 for obtnining longitudinal control. Whether the longi- 15. Fehlner, Leo F.: A Study of the Effects of Vortical Tail Aroa and tudinnl control is obtained by elevons or by CL horizontal tail D i h e M on the Lateral Maneuverability of an Airplane. NACA located on a t r d boom would seem to have a secondmy ARR, Oct. 1941.
16. Jones, Robert T.: The Influence of Lateral Stability on Disturbed influence on the ultimate performmce to be expected. On Motions of an Airplane with Special Referonce to tho hfotions the basis of the present knowledge of the stability and con- Produced by Gusts. NACA Rep. No. 638,1038.
trol chnrncteristics of tniUess airplanes, it appears desirable 17. Campbell, John P., and Seacord, Charles L., Jr.: Tlio Effoct of to make a comprehensive study of the compmtive perform- Mass Distribution on the Lateral Stability and Control Char- mce to be espected from tailless and conventional nirplanes acteristics of an Airplane as Determined by Tcsts of a Modo1 in the Free-Flight Tunnel. NACA Rep. No. 760, 1943.
before proceeding further with stability and control studies.
18. Bamber, M. J., and House. R. 0.: Wind-Tunnel Investigation of Effect of Yaw on Lateral-Stability Charactoristh. I-Four N. A. C. A. 23012 Kings of Various Plan Forms with and without Dihedral. NACA T N No. 703,1939.
19. Cohen, Doris: A Theoretical Investigation of tho Rolling Oscilla- LANGLDY MEMORIAL AERONAUTICAL LABORATORY, tions of an Airplane w i t h Ailerons Free. NACA Rep. No. 787, NATIONAL ADVISORY COMNJTPED FOR AERONAUTICS, 1 9 M LANGLEY FIELD, VA., August 19,19&.
20. Campbell, John P., and Mathems, Ward 0.: Experimental Dc- termination of the Yarning Moment DUG to Yawing Contributed by the Wing, Fuselage, and Vertical Tail of a hlidwing Airplano Model. NACA ARR No. 3F28, 1943.
REFERENCES 21. Harmon, Sidney &I.: Determination of the Damping Momont in 1. Jones, Robert T.: Notes on the Stability and Control of Taillc+s Yawing for Tapered Wings with Partial-Span Flaps. NACA Airplanes. NACA T N No. 837, 1941.
ARR No. 3H26,1943