NACA

RESEARCH MEMORANDUM I

FULL-SCALE WIND-TUNNEL TESTS OF THE LONGITUDINAL

XV-1 CONVERTIPLANE IN THE AUTOROTATING

FLIGHT RANGE

By David H. Hickey

Ames Aeronautical Laboratory Moffett Field, Calif.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON May 17, 1956

Declassified December 13, 1957

J NACA RM A55K2la

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

FULL-SCALE WIND-TUNNEL TESTS OF THE LONGITUDINAL

STABILITY AND CONTROL CHARACTERISTICS OF THE

XV-1 CONVERTIPLANE I N THE AUTOROTATING

FLIGHT RANGE

By David H. Hickey

SUMMARY

An invest igat ion was conducted t o determine ce r t a in longi tudinal s t a b i l i t y , control , and performance charac te r i s t i cs of t h e X V - 1 converti- plane i n t he autogiro- and airplane-type phases of f l i g h t . The fixed-wing and three-bladed ro tor so share t he t o t a l weight t h a t i n autogiro-type f l i g h t , the wing l i f t i s small, whereas fo r airplane-type f l i g h t t h e ro to r l i f t i s small and constant.

I n the autogiro-type f l i g h t configuration (speeds from 75 t o 115 knots) the convertiplane was s tab le with speed, but was unstable with angle of a t t ack . I n t h e airplane-type f l i g h t range from 115 knots t o maximum speed, s t a b i l i t y both with speed and with angle of a t t a ck was noted. Adequate longi tudinal control was indi.cated throughout t h e t o t a l airspeed range.

Interference between the f ixed wing and ro tor caused t he l i f t - d r a g r a t i o s of t he complete t o be appreciably l e s s than the l i f t - drag r a t i o s calcula ted from the individual component data .

Drag of t he complete a i r c r a f t was reduced i n a "clean-up" program. This reduction indicated an approximate 10-knot increase i n maximum speed.

It i s shown t h a t r e l i a b l e predictions can be made of t h e s t a t i c s t a - b i l i t y contributions of t h e V-tab horizontal t a i l . Calculated downwash a t t he t a i l compared wel l with t he measured downwash.

NACA .RM A55K21a

INTRODUCTION

The XV-1 convertiplane i s designed t o combine t he fea tures of a hel icopter and an airplane. A helicopter-type ro tor , a small f ixed wing, and a pusher propel ler a r e combined i n an e f f o r t t o obtain t he be s t char- a c t e r i s t i c s of both a i r c r a f t through t h e f l i g h t range from hovering t o maximum forward speed.

The projected f l i g h t plan f o r t he XV-1 convertiplane c a l l s f o r : (1) hel icopter f l i g h t from 0 t o 80 knots forward speed, (2 ) t r ans f t i on (autogiro type) f l i g h t from 80 knots t o 115 knots forward speed, and (3) airplane-type f l i g h t f o r forward speeds greater than 115 knots. I n t h e t r a n s i t i o n speed range, autorota t ion of t h e ro tor provides t he major por t ion of t h e l i f t . A t speeds above those f o r t r ans i t i on f l i g h t , the ro to r l i f t i s reduced t o a small f rac t ion of t h e t o t a l , although s t i l l i n autorota t ion, so t h a t the convertiplane charac te r i s t i cs a r e comparable t o those of an a i rplane.

To enable accomplishment of f l i g h t i n t h e t r ans i t i on range and a i r - plane range, t h e XV-1 incorporates two novel fea tu res . The f i r s t of these i s a f l oa t i ng hor izontal s t a b i l i z e r t o minimize t h e adverse e f f ec t on longi tudinal s t a b i l i t y of t he large downwash changes associated with t he unusually broad speed range. The second i s a control mechanism t o allow s tab le ro to r operation a t high tip-speed r a t i o s . Although smail-scale model t e s t s had been car r ied out t o develop t h e various components of t h e a i r c r a f t and t o show t h a t t h e i r combination would enable accomplishment of t h e f l i g h t plan, a question exis ted as t o t h e app l i cab i l i t y of small- model data t o t h e fu l l - sca le convertiplane i n t he t r ans i t i on and a i rplane phases of f l i g h t . Therefore, a fu l l - s ca l e wind-tunnel invest igat ion was conducted i n t h e Ames 40- by 80-foot wind tunnel i n order t o subs tan t ia te t h e i n t e g r i t y of t h e XV-1 design.

The pr inc ipa l purpose of t h i s repor t i s the determination of longi- tud ina l s t a b i l i t y , control , and performance charac te r i s t i cs of the con- ver t ip lane i n the two autorota t ive conditigns. I n addit ion, c e r t a in relevant data a f fec t ing these charac te r i s t i cs a re compared with o ther data obtained i n the development t e s t s . These data a re presented i n r e f - erences 1 and 2. Furthermore, an ana ly t ica l method i s derived t o est imate the V-tab ho r i zon t a l - t a i l contribution t o these charac te r i s t i cs and i s compared t o t he fu l l - s ca l e and small model t e s t r e s u l t s . Par t of t h i s analysis i s presented i n appendix B of reference 1. Some f u l l scale , dynamic, and loads data f o r t h i s convertiplane a re presented i n reference 3.

SYMBOLS

a l i f t - cu rve slope, per deg

c chord, f t

NACA RM A55K2la

l b I 2 c2dy

E mean aerodynamic chord (M.A.C .) , pb/B , ft

drag c D drag coefficient, -

qsR

lift CL lift coefficient, - ~ S R

c m pitching-moment coefficient about the rearmost center-of-gravity position,pitching moment

qsRR

C m ~ horizontal-tail pitching-moment coefficient about the hinge line,

pitching moment

qSirC~

hinge moment Ch hinge-moment coefficient,

qStct

Ch6 change in tab hinge-moment coefficient with respect to tab deflec- tion, per deg

Ch, change in tab hinge-moment coefficient with respect to the tail angle of attack, per deg

rotor thrust coefficient, T c T

P(~R) 2sR

K trim-tab spring constant, ft-lb/deg

IT distance from moment center to floating-tail hinge line, ft

9 free-stream dynamic pressure, lb/sq ft

R rotor radius, ft

S area, sq ft

T resultant thrust of rotor, lb

4 NACA RM A5TK2la

w v e r t i c a l component of ve loc i ty at any point i n t he ro to r flow f i e l d , f t / s ec

v free-stream a i r veloci ty , f t / s ec

V free-stream air veloci ty , knots

a angle of a t t ack , deg

a f angle of a t t a ck of t h e propel ler t h ru s t ax i s with respect t o t he f r e e stream, deg

P con t ro l s t i c k def lect ion from center , deg

Y - r a t i o of trim-tab def lect ion a t some airspeed t o tr im-tab ET def lect ion at qT = 0

Oaf change i n angle of a t tack, deg

n6ST change i n s t a b i l i z e r incidence angle, deg

OCm pitching-moment coef f ic ien t increment

ACD drag coef f ic ien t increment

&s servo-tab angle with respect t o t he s t a b i l i z e r chord l i n e , deg

&so servo-tab s e t t i n g ( t ab def lect ion at o0 s t a b i l i z e r angle), deg

6t trim-tab angle with respect t o t h e s t a b i l i z e r chord l i n e , deg

6~ tr im-tab angle with respect t o t h e s t a b i l i z e r chord l i n e at zero 9 ~ 7 deg

&TO tr im-tab s e t t i n g a t zero qT and zero s t a b i l i z e r angle, deg

6sT s t a b i l i z e r angle with respect t o t h e t h ru s t ax i s , deg

E downwash angle, deg

v C1 t i p speed r a t i o , - wR

P densi ty of free-stream air, slugs/cu f t

T l inkage r a t i o , change i n t ab def lect ion with change i n s t a b i l i z e r angle

NACA RM A55K21a

w angular veloci ty of r o to r , radians/sec

Subscripts

R r o t o r

S servo t ab

T hor izonta l t a i l

t t r i m t ab

w wing

AIRCRAFT AND APPARATUS

Photographs of t he XV-1 convertiplane a s mounted i n t he t e s t sect ion of t he wind tunnel a r e shown i n f i gu re 1. The geometric cha r ac t e r i s t i c s of t h e a i r c r a f t a r e given i n f igure 2 and t a b l e I.

Power Plant

The a i r c r a f t was powered by a 550-horsepower, nine-cylinder, r a d i a l engine. The power from the engine was used e i t he r f o r dr iv ing t h e r o t o r i n hel icopter f l i g h t or f o r dr iv ing t he f ixed-pitch, pusher-type p rope l le r i n autogiro-type and airplane-type f l i g h t . The function of t h e engine f o r he l i cop te r f l i g h t w a s t o dr ive two cen t r i fuga l a i r compressors t h a t supplied a i r t o t h e pressure j e t engines on t he ro to r t i p s . Since h e l i - copter f l i g h t was not studied i n these t e s t s , t h e burners were not oper- a ted . A brake was provided t o lock t h e propel ler during hel icopter f l i g h t .

Rotor

The mechanics and cha rac t e r i s t i c s of the convertiplane ro to r hub and controls were e s sen t i a l l y as described and discussed i n references 2 and 4. A l inkage was incorporated between the ro to r blades and the swash p l a t e which caused the ac tua l blade pi tch angle t o depend. on the coning angle (component of the blade fla.pping angle t h a t i s independent of ro to r blade azimuth). One degree of posi t ive coning angle reduced the blade p i t ch angle by 2.2'.

NACA RM A55K21a

I n t he autogiro-type f l i g h t range, the incidence of the swash p l a t e was f ixed at +3O, and the blade pi tch angle a t 0.75R was f i xed at 6 O

(measured at zero coning angle) with reference t o the swash p la te . Changes i n l i f t coef f i c ien t of the ro to r were brought about by va r i a t i on of angle of a t t ack of the convertiplane.

I n the airplane-type f l i g h t configuration, the hub was mechanically locked t o the swash p la te , and the blade pi tch angle a t 0.75R with respect t o the swash p l a t e was f ixed a t 0'. A governor was employed t o tilt the swash p l a t e i n such a way a s t o keep the ro to r rpm constant . I n t h i s configuration, the rotor-blade-swash-plate linkage minimized blade f l a p - ping motions, thus delaying blade f lapping i n s t a k i l i t y a t the high advance r a t i o s t yp i ca l of the airplane-type f l i g h t configuration. Character is t ics of the ro to r were then a s follows: (1) small ro to r l i f t coef f i c ien t , (2) ro to r l i f t coef f i c ien t constant with convertiplane angle of a t t ack , and (3) ro to r l i f t constant with airspeed.

Horizontal T a i l

The XV-1 was equipped with a f ree - f loa t ing hor izonta l t a i l (designated V-tab hor izonta l t a i l ) i n order t o a l l e v i a t e t he s t a b i l i t y problems a r i s i n g from the large var ia t ion of downwash encountered throughout t he speed range. The t a i l and t h e linkage of t h e t abs t o the controls a r e i l l u s - t r a t e d i n f igure 3. Location of t he hinge l i n e of the hor izonta l t a i l was at 22-1/2-~ercent chord. Two 15-percent-chord t ab s , each hal f of t h e t a i l span, served as a t r i m t ab and a servo t ab . The t r i m t a b on t he XV-1 did not br ing about zero s t i c k force , but r a t he r allowed posi t ioning of t h e control column. A n adjus table centering spring on t he control column was used t o obtain zero s t i c k force . Another spring was placed i n t h e tr im-tab linkage t o compensate p a r t i a l l y f o r t he change i n downwash angle with forward speed. Downwash e f f e c t s on t he hor izonta l t a i l a t forward speeds from 0 t o 40 knots were minimized by a t h i r d spring which held t h e s t a b i l i z e r at 60' def lec t ion. The moment provided by t h e s t a b i l i z e r spr ing i s shown i n f igure 4 ( a ) . Both t ab s were equipped with a linkage which provided an approximate s ine curve re la t ionship between t h e s t ab i - l i z e r angle and t ab angles ( f i g . 4 ( b ) ) . I n the p r a c t i c a l range of s t ab i - l i z e r angles, t he linkage r a t i o was very near ly uni ty .

For t he t e s t s t he hor izonta l t a i l was equipped with selsyn u n i t s t o measure angles on t he t abs and s t a b i l i z e r . The t a i l l i f t was indicated by ca l ib ra ted s t r a i n gages on t he surfaces of the twin t a i l booms.

NACA RM A55K2la 7

The airspeed range of t h e t e s t s conducted i n t h e Arnes 40- by 80-foot wind tunnel was from 75 t o 150 knots.

The invest igat ion consisted of f i ve phases:

1. Tests t o determine t h e convertiplane longi tudinal character is- t i c s and t h e V-tab hor izon ta l - t a i l charac te r i s t i cs with t h e ro tor blades and propeller removed.

2. Tests t o determine t h e e f fec t of t he powered propel ler on t h e convertiplane longi tudinal charac te r i s t i cs and V-tab hor izon ta l - t a i l charac te r i s t i cs with t h e ro to r blades removed.

3. Tests t o determine t h e con -er t ip lane longi tudinal character is- t i c s i n both f l i g h t configurat ons with t he propel ler removed.

4. Tests t o determine t h e e f fec t of t he powered propel ler on t he con- ver t ip lane longitudinal charac te r i s t i cs i n both f l i g h t configurations.

5. Tests t o determine t h e drag of the conver-tiplane components and means of reducing t he drag. Figure 5 shows some of t he pylon modi- f i c a t i ons t e s t ed during t h i s phase of t he invest igat ion.

During t he powered-propeller t e s t s , it was necessary i n some cases t o allow p i t ch freedom above t e s t airspeeds of 75 knots because otherwise t h e convertiplane developed undesirable vibrat ions . P i t ch freedom was r e s t r i c t e d t o a t o t a l of 112' i n t h e autogiro-type f l i g h t configuration and 6' i n t he airplane-type f l i g h t configuration.

CORRECTIONS TO DATA

S t ru t drag t a r e s were applied only t o the data obtained i n t h e s tud ies of drag reduction. The drag t a r e of the p i t ch s t r u t was obtained by removing the p i tch s t r u t and i t s f a i r i ngs from t h e wind-tunnel t e s t sect ion, allowing t he a i r c r a f t t o f l o a t f r e e . The change i n drag due t o t h e removal of t he p i tch s t r u t was within t he l e a s t count of t he drag- indicat ing system. Drag of t h e main support s t r u t s was obtained a f t e r t h e convertiplane was removed from the t e s t section and f a i r i ngs were added t o t he s t r u t b a l l sockets. A value of ACD of 0.0006 was obtained f o r t h e main s t r u t s alone throughout the speed range invest igated.

8 NACA RM A55K2la

The l e a s t counts of t he force coef f ic ien t s when t h e convertiplane was res t ra ined i n p i t ch a r e shown below:

Quantity Least count

L i f t coef f ic ien t 10.001 T a i l l i f t coeff ic ient 1 . O O ~ Drag coef f ic ien t 1 .0001 Pitching-moment coeff ic ient 1.0003

The l e a s t counts of a l l angular data, excepting ?jt, were 10.25° with t he a i r c r a f t f ixed i n p i tch . The l e a s t count i n ?jt da ta was about k 0 . 5 ~ .

The l e a s t counts of force and angular data obtained during the f ree- to -p i tch port ion of t h e t e s t i n g were increased, a s sho~m below:

Quantity Least count

L i f t coeff ic ient t o . 005 T a i l l i f t coeff ic ient t .01 Drag coeff ic ient t .0002

The l e a s t counts of a l l t h e angular data with p i t ch freedom were about +lo.

Boundary correct ions were not applied f o r t he convertiplane because they were unknown; however, wind-tunnel wall corrections t o the data with ro to r blades removed were calculated and found t o be negl igible .

PRESrnTATION OF FIGURES

Full-Scale Data

These fu l l - s ca l e t e s t s of the XV-1 convertiplane were di rected toward an examination of only t h e autogiro-type and airplane-type f l i g h t configu- ra t ions because t he hel icopter charac te r i s t i cs could be obtained r e l a t i v e l y sa fe ly and ea s i l y i n f l i g h t t e s t s .

The three-component force data presented i n f igures 6 and 7 apply t o t he a i r c r a f t with ro tor blades and propel ler removed. These t e s t s had the following purposes: (1) t o determine t he cha rac t e r i s t i c s of t h e hori- zontal t a i l , (2 ) t o serve a s a ba s i s f o r examining t he e f f ec t on s t a b i l i t y and control of adding a ro tor t o an a i rplane configuration, and (3) t o serve a s a bas i s f o r examining the e f fec t on s t a b i l i t y and control of adding t he propel ler sl ipstream.

NACA RM A55K2la 9

The three-component force data presented i n f igures 8, 9, and 10 show t h e a i r c r a f t cha rac t e r i s t i c s with the propel ler removed i n t h e autogiro-type and airplane-type f l i g h t regimes. These data a re t h e b a s i s f o r the s t a b i l i t y and control analys is and performance evaluation pre- sented herein.

The force data presented i n f igure 11 and t a b l e I1 a r e t he r e s u l t s with power applied t o t he propel ler with t he ro to r blades removed and a re a l so u t i l i z e d i n t he s t a b i l i t y and control analys is .

The strain-gage and selsyn data presented. i n f igures 12, 13, and 1 4 show the cha rac t e r i s t i c s of t he V-tab hor izontal t a i l .

Figure 15 presents t he measured downwash angles a t t he t a i l with and without the ro tor . Calculated do~inwash angles a t t he t a i l i n t h e autogiro- type f l i g h t configuration a r e a l so included f o r comparison.

The pitching-moment data presented i n f igure 16 a r e cross p l o t s of the data presented i n f igure 7. These data a r e presented i n t h i s form t o show the speed s t a b i l i t y charac te r i s t i cs of t h e convertiplane without ro to r blades.

Figure 17 presents force data showing t h e e f f ec t of r o to r operation on t he pitching-moment var ia t ion with airspeed. These data were obtained from the data presented i n f igures 8 and 10 f o r t h e propeller-removed configuration . The corresponding powered-propeller data a t a tr im-tab s e t t i n g of -20' were those presented i n t ab l e 111. Only t he data from the t e s t s with t h e convertiplane f ixed i n p i t ch a r e presented i n f igure 17. I n addit ion, the calcula ted var ia t ion of p i tching moment with a i rspeed f o r powered-propeller conditions a t a trim-tab s e t t i n g of -60° i s shown. Figure 17 i s concluded with calculated curves, corresponding t o t h e pitching-moment curves, showing t he var ia t ion of s t i c k pos i t ion f o r t r i m with airspeed. These data were used t o appraise t h e stick-motion s t a b i l i t y of the convertiplane.

Figure 18 presents calculated values showing t h e control ef fect iveness of t he V-tab horizontal t a i l . These r e s u l t s were used i n t he ca lcu la t ions performed t o obtain t he curves i n f igure 17.

Figures 19 and 20 present force data showing t h e e f f ec t on l i f t and drag of t he combination of a fixed-wing and ro to r f o r both convertiplane configurations considered. The data presented i n f igure 21 show the apparent decrease i n t he wing angle of a t t a ck when t he wing i s submerged i n t he ro tor flow f i e l d . These data provide an est imate of t he l i f t in terference between a f ixed wing and a ro to r i n combination.

Figure 22 presents data showing sea-level maximum speed and t he increase i n maximum speed rea l ized by ce r ta in drag reductions. These data

10 NACA BM AfSfSK21a

were derived from t h e f ixed-pitch force data i n t ab le I11 and the data shown i n f i gu re s 8 and 10,

Small-Scale Tests

The r e s u l t s of c e r t a in t e s t s with the s m a l l models used t o develop t he convertiplane a r e presented herein i n order t o assess t h e i r value i n predic t ing the behavior of the fu l l - s ca l e machine. Although small-model t e s t s a r e not s t r i c t l y comparable t o the f i l l - s c a l e t e s t s , t h e i r use i n respect t o convertiplane design and development i s general, and an ind i - ca t ion of t h e i r r e l i a b i l i t y i s therefore believed t o be of value. Because an evaluation of t h e small-model data i s not relevant t o the p r inc ipa l purpose of t h i s repor t , the d-ata a r e included without discussion. A l l of the small model data were obtained from reference 1.

Test r e s u l t s presented herein were obtained a s follows:

1. Rotor-blades, t a i l , ro to r hub, and propeller-removed data were obtained from an 8/31-scale model of t he XV-1 t e s ted i n t he Universi ty of Washington 8- by 12-foot wind tunnel. Engine i n l e t s , landing skids , wing-tip skids, and t a i l fans were not reproduced on t h e model. These data a re presented i n f igure 6.

2. V-tab hor izon ta l - t a i l data were obtained f r o m t e s t r e s u l t s of a geometrically and dynamically s imilar 215-scale model of t he t a i l . The wing and fuselage t e s t e d i n conjunction with t h e t a i l provided a downwash d i s t r i bu t i on s imi la r t o t h a t expected f o r t h e fu l l - s ca l e convertiplane i n airplane-type f l i g h t . The t e s t s were conducted i n t he University of Washington 8- by 12-foot wind tunnel. These data a r e presented i n i ' i [p re 14.

3. Downwash angle at t he t a i l w a s measured with a complete 8/31-scale model. The model was mounted i n t h e re tu rn passage of t he Wright A i r Development Center 20-foot wind tunnel. The diameter of t h e passage was 45 f e e t and t h e model ro to r w a s mounted 12 f e e t above t he lowest por t ion of t he d-uct. These data a r e presented i n f igure lfS(a).

DISCUSSION

Limitat ions due t o propeller-induced vibrat ions on t h e range of var iables t h a t could be covered with t he a i r c r a f t f ixed i n p i t ch made it impossible t o obtain a d i r ec t evaluation of t he longi tudinal s t a b i l i t y and control throughout t he autogiro-type and airplane-type f l i g h t ranges. The data obtained i n the invest igat ion were therefore extended by ana- l y t i c a l means i n order t o determine more completely t he charac te r i s t i cs of t he convert iplane, I n order t o evaluate t he worth of t he analyses,

NACA RM A55K2la

comparisons of ana ly t ica l r e s u l t s with applicable fu l l - s ca l e data a r e included. The d e t a i l s of t he analyses a r e given i n Appendix A f o r hor i - zontal t a i l charac te r i s t i cs , Appendix B f o r downwash, and Appendix C f o r s t a b i l i t y and control . A representa t ive gross weight of 5,000 pounds was used throughout since the weight var ies i n f l i g h t from about 4800 t o 5300 pounds.

A l l pitching-moment r e s u l t s apply t o t he rearmost possible center- of-gravity posi t ion. Therefore, t he convertiplane s t a b i l i t y character- i s t i c s reported herein per ta in t o t he l e a s t s t ab le extreme.

Due t o t h e differences i n ro tor behavior, t he autogiro-type f l i g h t configuration w i l l be considered separately.

S t a t i c Angle-of-Attack S t a b i l i t y

Knowledge of t he var ia t ion of downwash angle and center of pressure with angle of a t t ack i s i n su f f i c i en t , i n t he case of t he XV-1, t o def ine the angle-of-attack s t a b i l i t y about a given center of gravi ty loca t ion . This i s t r ue because t h e V-tab hor izontal t a i l and t he ro tor can a l s o be eLxpected t o influence angle-of-attack s t a b i l i t y cha rac t e r i s t i c s .

The e f fec t of t he V-tab hor izontal t a i l on angle-of-attack s t a b i l i t y can be determined by examination of t he following equation (derived i n Appendix A ) , which defines t h e t a i l f l oa t i ng angle.

It can be deduced from t h i s equation t h a t t h e s t ick-f ixed s t a b i l i t y con- t r i bu t i on of t he V-tab horizontal t a i l shculd be t he same a s i f t h e incidence were constant only insofar a s exp l i c i t dependence on angle of a t t ack i s concerned. However, t he ve loc i ty a t t he t a i l influences (1) t h e dynamic pressure, qT, and (2) t he trim-tab r a t i o , y , and it i s apparent t h a t changes i n wake-velocity magnitude with angle of a t t ack can combine t o change t he s t a b i l i z e r angle, and thus t he s t a b i l i t y index.

The e f f ec t of t he ro tor l i f t on t he angle-of-attack s t a b i l i t y i s a r e s u l t of t he v e r t i c a l displacement of t h e ro tor hub from t h e center of gravi ty . This means t h a t t he ro to r t h ru s t vector need not necessar i ly a c t through t h e center of gravi ty and thus w i l l introduce moments which can a f f e c t angle-of-aktack s t a b i l i t y .

Autogiro-t~ype f l i gh t . - The data of f igure 8 show the XV-1 t o be unstable with angle of a t t ack a t t h e th ree t e s t airspeeds. Possible causes a r e , of course,

1. The adverse var ia t ion of downwash angle and wake ve loc i ty a t t h e t a i l with angle of a t t ack .

2. The adverse var ia t ion of t h e magnitude and or ien ta t ion of t h e r o t o r t h r u s t vector with angle of a t t ack .

The e f f ec t of downwash and wake-velocity var ia t ion on angle-of-attack s t a b i l i t y i s appraised by means of t h e equation f o r t h e var ia t ion of t a i l angle of a t t a ck with fuselage angle of a t t ack a s a function of downwash- angle and s tabi l izer-angle var ia t ion with angle of a t tack. This i s

Examination of t h e data shown i n f igure 15 indicates t h a t a c T / h f i s near ly un i t y , Therefore, t he f i r s t two terms on t h e right-hand s ide of t he foregotng equation combine i n such a way t h a t the angle of a t t a ck of t h e t a i l i s unaffected by changes of downwash angle with angle of a t t a ck of t h e convertiplane. Furthermore t h e data shown i n f igure 12 indica,te that a8,,/aaf i s negl igible . Thus aaT/aaf i s approximately zero,

which implies t h a t t h e wake ve loc i ty i s nearly constant with angle of a t t ack . One can therefore conclude t h a t the t a i l contribution t o angle- of-a t tack s t a b i l i t y i s negl igible .

.!An indicat ion of t h e ro to r e f f ec t on s t a b i l i t y can be deduced from ,-.,

comparison of t he C ~ v = 0 pitching-moment curve of f igure 6 and t he I

pitching-moment curves of f igure 8. From the data i n f igure 6, it i s seen t h a t t he contribution of t he t a i l t o angle-of-attack s t a b i l i t y was zero; i n f i gu re 8 t he t a i l moment was constant (contr ibut ion t o s t a b i l i t y was zero). Therefore, the di f ference i n t he slope of t h e pitching-moment curve i n f igures 6 and 8 i s d i r e c t l y a t t r i bu t ab l e t o a d i f ference i n t h e magnitude and l i n e of act ion with respect t o t he center of gravi ty of t he r o t o r t h r u s t vector. Thus, adding t h e ro to r t o t he convertiplane brought about an unstable change i n angle-of-attack s t a b i l i t y , a C , / b f , of approxi- mately 0,0009 per degree a t 100 knots airspeed.

NACA RM A55K2la 1.3

The data i n f igure 9 show the angle-of-attack s t a b i l i t y with reduced ro tor l i f t and t he ro tor lift allowed t o vary with angle of a t t ack , This brings about neu t ra l angle-of-attack s t a b i l i t y . Even i n t h i s small r o to r l i f t condition, f igure 10 shows t h a t t he ro tor causes an unstable change of acm/hf of 0.0005 per degree.

Approximate calculations indicate t h a t t he neu t ra l point of the con- ver t ip lane was a t about 18-percent M.A.C., or 2-percent M.A.C. ahead of t h e t e s t center of gravity. Therefore, with a forward s h i f t i n center- of-gravity locat ion, t h e convertiplane could be made s tab le with angle of a t t ack i n autogiro-type f l i g h t .

The s c a t t e r of t he data presented i n f igure 8 i s too la rge t o show a change i n angle-of-attack s t a b i l i t y with speed, However, examination of the downwash data ( f i g . l 5 ( a ) ) indicates t h a t , although i s about un i ty throughout the autogiro-type f l i g h t range, a s l i g h t decrease i n beT /h f with increasing airspeed occurs a t t he angles of a t t a ck f o r

l e v e l f l i g h t . This tends t o increase s t a b i l i t y . I n addit ion, t h e l i f t - curve slope of t h e convertiplane i s shown i n f igure 8 t o decrease with increasing airspeed. This change i s a t t r i bu t ab l e t o t he ro tor character- i s t i c s , and indicates t h a t t he var ia t ion i n ro tor t h r u s t with angle of a t t ack decreases with increasing airspeed. These e f f ec t s probably combine t o produce a small increase i n angle-of-attack s t a b i l i t y with speed.

Angle-of-attack s t a b i l i t y f o r autogiro-type f l i g h t could not be determined with t he propeller operating because of excessive vibrat ion. However, a comparison of data obtained with the ro tor blades removed both with and without powered propeller ( f i g . 11) shows t h a t t he e f f ec t of pro- p e l l e r power on s t a b i l i t y was negligible. It i s believed t h a t t h e addi t ion of t he ro to r blades would not s ign i f ican t ly a l t e r t h i s r e su l t . Therefore, it i s believed t h a t t he angle-of-attack s t a b i l i t y of autogiro-type f l i g h t w i l l be unaffected by propel ler operation.

I n summary, it can be s t a t ed t h a t i n autogiro-type f l i g h t , angle-of- a t t ack s t a b i l i t y did not e x i s t f o r the most rearward center-of-gravity posit ion. This r e s u l t was a consequence of the var ia t ions with angle of a t t ack of r o to r downwash, ro to r t h ru s t , and l i n e of act ion of t h e t h r u s t vector under circumstances where the ro tor provided a l a rge port ion of t he e n t i r e l i f t of the a i r c r a f t . Calculations indicate t h a t a 2-percent M.A.C, forward s h i f t of the center of gravi ty would give neutra l angle- of-a t tack s t a b i l i t y . These r e s u l t s assume no l ag between convertiplane angle-of-attack changes and ro to r rpm changes. The e f f e c t of l a g w i l l be discussed i n the aynamic s t a b i l i t y section.

.- The convertiplane angle-of-attack s t a . b i l i t y charac te r i s t i cs i n the airplane-type f l i g h t configuration a re shown by the data 5.11 f igure 10. These were due primarily t o the following ro to r charac te r i s t i cs :

NACA RM A55K21a

1; Rotor l i f t coef f ic ien t constant with angle of at tack.

2. Rotor l i f t small i n comparison with t he t o t a l convertiplane lift.

The implications of these ro to r cha rac t e r i s t i c s with respect t o angle-of-attack s t a b i l i t y fac tors a re :

1. The ro to r contribution t o downwash a t t h e t a i l was constant with angle of a t tack.

2. The change with angle of a t t a ck of t he ro to r l i f t contribution t o pi tching moment was small because of t he maintenance of small r o to r l i f t .

Item 1 ind ica tes t h a t t h e V-tab hor izon ta l - t a i l contribution t o s t a b i l i t y should be t he same with t he ro tor blades removed a s with the convertiplane i n t h e airplane-type f l i g h t configuration. Item 2 indicates t h a t t h e large des tab i l i z ing e f f ec t of t h e ro to r which i s present i n t h e autogiro- type f l i g h t configuration would be g rea t ly reduced i n t he airplane-type f l i g h t configuration. Therefore, t h e angle-of-attack s t a b i l i t y with t h e ro to r blades removed should be about t he same a s t he s t a b i l i t y of t he convertiplane i n t he airplane-type f l i g h t configuration. Comparison of t h e appropriate curves i n f igure 10 confirms t h i s . The adverse e f f ec t on s t a b i l i t y of allowing t h e ro to r l i f t t o vary with angle of a t t ack even though t h e ro to r l i f t i s kept small i s shown by t h e governor-inoperative data rep lo t ted from f igure 9.

The e f f ec t of speed on angle-of-attack s t a b i l i t y i s shown by t he 65, = lo0, 6~~ = 20' curves of f i gu re 10. Comparison of t h e 100- and

125-knot airspeed curves indicates t h a t l i t t l e , i f any, change i n angle- of-a t tack s t a b i l i t y with airspeed occurred f o r t h e airplane-type f l i g h t configuration. This can be a t t r i bu t ed d i r e c t l y t o t he ro to r rpm governor which kept t he ro to r l i f t constant.

The e f f ec t of powered propel ler on angle-of-attack s t a b i l i t y with t h e ro to r blades removed i s shown by t h e data of f igure 11. Again, t he data ind ica te t h a t very l i t t l e change i n angle-of-attack s t a b i l i t y occurred with a constant propel ler rpm. These curves a r e believed ind ica t ive of t he e f f ec t of power on t h e airplane-type f l i g h t configuration a s wel l a s on t h e autogiro-type f l i g h t configuration.

It can be concluded t h a t t h e XV-1 should have angle-of-attack s ta - b i l i t y i n t he airplane-type f l i g h t configuration. This i s primarily a r e s u l t of keeping the ro to r l i f t coef f ic ien t constant with angle of a t tack.

NACA RM A55K2la

S t a t i c Speed S t a b i l i t y

Speed s t a b i l i t y i s defined by t h e slope of t h e pi tching moment versus ve loc i ty curve a t t r im f o r constant angle of a t t ack and control se t t ings . A pos i t ive slope i s s t ab le . Speed s t a b i l i t y i s normally an inherent qua l i ty of powered ro to r s whereas autorota t ing ro to r s and a i rplanes gen- e r a l l y have neu t ra l speed s t a b i l i t y .

The neu t ra l speed s t a b i l i t y associated with t h e conventional auto- g i ro and a i rplane can be changed by t h e introduction of anything which causes an aerodynamic coef f ic ien t t o depend upon airspeed. Propel ler operation, a spring i n t h e control s y ~ t e m , and Mach number e f f e c t s a r e some of t he poss ible causes of d e v i a t i m from neu t ra l speed s t a b i l i t y by an autogiro o r a i rplane.

The pi tching moment of t h e autorota t ing ro to r i n general and t he XV-1 r o to r i n t h e autogiro-type f l i g h t configuration i n pa r t i cu l a r i s constant with airspeed because t h e ro tor rpm i s f r e e t o change with a i r - speed. This means t h a t f o r a given blade p i t ch angle and a i r c r a f t angle of a t t ack , t h e tip-speed r a t i o , l i f t coef f ic ien t , drag coef f ic ien t , and downwash of t h e ro to r tend t o remain constant with changing airspeed. The longi tudinal data ( c ~ , C D ) i n f igure 8 and t h e downwash-angle data of f igure l 5 ( a ) show t h a t t h i s i s e s sen t i a l l y t h e case f o r t he XV-1 i n t h e autogiro-type f l i g h t configuration. I n t he airplane-type f l i g h t configu- ra t ion , on t h e other hand, CL, CD, p, and ET vary with airspeed at a constant angle of a t t ack because t h e ro to r rpm was kept constant. How- ever, t h e ro to r was so l i g h t l y loaded (cL < 0.03) t h a t these var iables

R = had small e f f ec t on t h e speed s t a b i l i t y compared t o t h e e f f e c t of t h e V-tab hor izontal t a i l . From these considerations, it i s apparent t h a t rotor-removed data w i l l su f f ice t o determine t h e speed s t a b i l i t y charac- t e r i s t i c s of t h e convertiplane.

The important f ac to r s a f fec t ing speed s t a b i l i t y are :

1. Variation of t he t a i l incidence angle with airspeed, including propel ler e f f ec t s , which can be expressed as a6ST/a~Te This brought about :

(a) A s t ab i l i z i ng t a i l l i f t contribution t o pi tching moment with changing f l i g h t airspeed, Examination of f igure 1 4 shows a decrease i n t a i l incidence with increasing airspeed, which pro- duces a nose-up, o r s tab le , pitching-moment var ia t ion with airspeed.

(b) A des tab i l i z ing e f f ec t from t h e propel ler slipstream. Con- .

s iderat ion of t he momentum equation f o r propel lers shows t h a t , a t a constant engine rpm, t he increment of airspeed at t he t a i l would decrease with increasing f l i g h t airspeed. This would tend t o counteract t he e f f ec t of item ( a ) .

NACA RM A55K2la

2. Variation of item 1 with trim-tab se t t ing , which can be expressed a s d(a6ST/a~T)a6Toe Comparison of f igures 14(a) and 14(b) shows t h a t t he slope of the curve of s t a b i l i z e r incidence angle versus airspeed increased with increasingly negative trim-tab s e t t i ng , This ind ica tes an increasing t a i l contribution t o speed s t a b i l i t y a s t he trim-tab def lect ion becomes more negative.

Because t h e ro to r i n both t he autogiro-type and airplane-type f l i g h t configurations contributes very l i t t l e t o speed s t a b i l i t y ; t h e ro tor - removed data can be used t o determine t he speed s t a b i l i t y of t he XV-1 under both of these conditions. Figure 16 shows t he pi tching moment versus veloci ty curves f o r t he rotor-removed configuration. The curves of f i g - ure 16 a r e cross p l o t s of t h e data i n f igure 7. The powered-propeller curves i n f igure 16(c) were calculated by adding a pitching-moment incre- ment due t o t h e powered propel ler , obtained from t ab l e s I1 and 111, t o t he ro tor - and propeller-removed pi tching moments.

Analysis of f igure 16 indicates t h a t t he pi tching moment versus veloci ty curve f o r a given angle of a t t ack and veloci ty i s a va l id ind i - cation of speed s t a b i l i t y a t each tr im-tab se t t ing . The servo-tab s e t t i n g could be changed t o bring about t r i m , but the speed s t a b i l i t y a t a given speed did not change. It can therefore be concluded t h a t t he XV-1 w i l l have speed s t a b i l i t y throughout the f l i g h t range investigated with t h e propel ler removed (power o f f ) . The magnitude of the power-off s t a b i l i t y increases with l a rger negative trim-tab def lect ions and decreases wi th airspeed. The powered propel ler decreases speed s t a b i l i t y and i s very l i k e l y t o produce speed i n s t a b i l i t y a t t he higher engine rpmts and smaller trim-tab se t t ings . Large negative trim-tab s e t t i ngs can be expected t o improve t he powered-propeller i n s t a b i l i t y and may completely el iminate it. The analysis i n Appendix A discusses these charac te r i s t i cs more completely.

Longitudinal St ick Motion and Control

A more forward s t i c k pos i t ion f o r t r i m with increasing airspeed i s defined a s s t ab l e s t i c k motion. Possible s t i ck motion i n s t a b i l i t y was evidenced by t he XV-1, both with power off ( ~ r o ~ e l l e r removed) and a t ra ted power, f o r a trim-tab s e t t i ng of -20'.

The important fac to rs with respect t o s t i c k motion f o r t he given configuration a r e a combination of t he fac tors discussed i n t h e s t a t i c angle-of-attack and speed-s tabi l i ty sections. They are :

1. The change i n speed s t a b i l i t y rtith af (change i n angle-of -a t tack s t a b i l i t y with speed), which can be expressed a s

NACA RM A53K2la

2. Variation of downwash a t the t a i l with af, which can be expressed a s

3 Control linkage charac te r i s t i cs , which can be expressed a s avc Irl =o/asso.

4. Horizontal - ta i l charac te r i s t i cs with airspeed, which can be expressed a s a6ST/a~T and a(a6ST/a~T)/a6Toe of these var iab les , t h e trim-tab s e t t i n g was t he most e a s i l y changed. Therefore, an analysis was conducted t o determine t h e s t i c k motion cha rac t e r i s t i c s a t a trim-tab s e t t i n g of -60'. The d e t a i l s of t h e analysis a r e given i n Appendix C .

Figures 17(a) and (b) present data showing t he var ia t ion of p i t ch ing moment t o be trimmed out a s a function of airspeed f o r c e r t a in t e s t con- d i t ions . The pi tching moments t o be trimmed out were obtained from f i g - ure 8 o r 10, depending on t h e convertiplane configuration i n question. These data were chosen a t t h e l i f t coeff ic ient required f o r l e v e l f l i g h t a t each airspeed. The powered-propeller data a t a trim-tab s e t t i n g of -20' were obtained by adding an increment i n pi tching moment due t o t h e powered propeller from t a b l e I11 t o the power-off p i tching moment obtained from f igure 8 or 10. The powered-propeller data a t a trim-tab s e t t i n g of -60' were calculated by t he method described i n Appendix C . The changes i n servo-tab def lect ion required t o balance the moments discussed above were obtained from t h e calculated V-tab ho r i zon t a l - t a i l cha rac t e r i s t i c s shown i n f igure 18 and t he corresponding curves of t h e s t i c k pos i t ion required fo r t r i m versus airspeed. These data a r e shown i n f igures 17 (c ) and (d ) . These calcula t ions a r e believed t o give only t h e gross e f f e c t s of power and control se t t ing .

Autogiro-type f l i g h t . - Analysis of f igure 17( c ) ind ica tes :

1. Power-off s t i c k motion with changing airspeed a t 6 ~ o

= -20' i s

unstable f o r t he low-speed end and neu t ra l ly s t ab l e a t t h e high-speed end of t he autogiro-type f l i g h t range.

2. The e f fec t of powered propel ler on s t i c k motion with changing airspeed i s t o produce i n s t a b i l i t y a t

&TO = -20°. The s t i c k t r a v e l

required f o r t r i m exceeds t he forward l i m i t at airspeeds l e s s than about 80 knots, so t h a t t he convertiplane then becomes uncontrollable.

3. The e f f ec t of changing t he trim-tab def lect ion t o -60' under t h e conditions i n item 2 would br ing about s tab le s t i c k motion wi th changing airspeed. This change i n trim-tab s e t t i n g places t h e s t i c k posi t ion f o r t r i m wel l within t h e control l im i t s .

NACA RM A55K2la

It can be concluded t h a t t h e XV-1 i n t he autogiro-type f l i g h t con- f igura t ion can have s t i c k motion sta 'oi l i ty with powered propel ler a t a trim-tab s e t t i n g of -60°. Moreover, it i s believed t h a t t h e XV-1 can have s t i c k motion s t a b i l i t y and be control lable a t a trim-tab s e t t i n g a s low a s -40'. Free-to-float t e s t s qua l i t a t i ve ly support t h i s conclusion.

Airplane-type f l i g h t . - Because of t he di f ference i n ro to r character- i s t i c s , the re i s considerable change i n t h e s t i c k posi t ion var ia t ion with airspeed between t he autogiro- and airplane-type f l i g h t configurations. Analysis of f igure 17(d) indicates :

1. St ick motion with power off i s s t ab le a t 6 = -20' i n t h e airplane-type f l i g h t range investigated. To

2. The e f f ec t of powered propel ler on s t i c k motion i s t o produce almost neu t r a l speed s t a b i l i t y a t ZTo = -20'. The s t i c k posi t ion f o r t r i m i s wel l within t he control l im i t s .

3 . The e f f ec t of changing t he trim-tab def lect ion t o -60° under t he conditions i n item 2 would br ing about s tab le s t i c k motion. This tr im-tab s e t t i n g places t he s t i c k pos i t ion f o r t r i m wel l within t he control l i m i t s .

It can be concluded t h a t t h e XV-1 i n t he airplane-type f l i g h t con- f igura t ion can have s t i c k posi t ion s t a b i l i t y with powered propel ler a t l a rge negative trim-tab se t t ings . It i s believed t ha t the XV-1 i n t h e airplane-type f l i g h t configuraticn can have s t ab l e s t i c k motion and be control lable at t h e same tr im-tab s e t t i ngs (ZTo= -40' t o -60') appropriate

t o autogiro-type f l i g h t . Free-to-float t e s t s qua l i t a t i ve ly support t h i s conclusion.

An i n t e r e s t i ng property of the V-tab hor izontal t a i l i s i t s ease of' modification. Increased s t i f f n e s s of t he trim-tab spring would provide greater control ef fect iveness , i f needed, without l a rge a i r c r a f t modifi- cations. An increase i n t he s t i f f n e s s of t he spring would, however, reduce t h e speed s t a b i l i t y contribution of t he horizontal t a i l a s ind i - cated by t h e analysis i n Appendix A.

Longitudinal Dynamic S t a b i l i t y

Within l i m i t s , indicat ions of t he dynamic s t a b i l i t y cha rac t e r i s t i c s of t h e convertiplane could be obtained from t h e f ree- to-pi tch t e s t r e s u l t s . The r e s u l t s a r e only an indicat ion because of t he constant tunnel a i r - speed, f r i c t i o n damping i n t he support system, and r e s t r a i n t i n v e r t i c a l t r ans l a t i on ,

NACA RM A55K2la 1-9

I n the autogiro-type f l i g h t configuration the small amount of p i t ch freedom (1/2O) did not allow determination of the s t a b i l i t y character - i s t i c s . However, it i s probable t h a t the convertiplane would not be a s susceptible t o i n s t a b i l i t y during gusts and pull-up maneuvers a s t h e s t a t i c angle-of-attack s t a b i l i t y discussion would indicate . This i s possible because of t he l a g between changes i n convertiplane angle of a t t a ck and corresponding changes i n ro to r rpm, I n conditions such a s these, the convertiplane would probably be a s s tab le i n the autogiro-type configura- t i o n a s i n t he airplane-type configuration.

I n the airplane-type f l i g h t configuration, the t e s t s indicated t h e a i r c r a f t would be dynamically s tab le with angle of a t tack. Dynamic i n s t a - b i l i t y did occur i n t he airplane-type f l i g h t configuration when t h e r o t o r rpm governor had too muchlag. Several s l i g h t modifications t o t h e governor corrected t h e excessive l ag and made t he convertiplane dynamically s t ab l e .

Interference Between t he Fixed Wing and Rotor

Because of the general i n t e r e s t i n the e f f e c t on an a i r c r a f t of com- bining a ro ta ry wing and a f ixed wing, the data i n f igures 19 and 20 were replot ted from f igures 7, 8, and 10 i n order t o compare d i r ec t l y the l i f t and drag contributions of the a i r c r a f t components i n both t he autogiro-type and airplane-type f l i g h t configurations. Included a r e fu l l - s ca l e ro tor - alone data obtained from reference 2.

Autogiro-type f l igh t . - The data i n f igure 19 ind ica te t h a t t h e meas- ured l i f t i n the autogiro-type f l i g h t configuration was l e s s than t he sum of t he rotor-removed and t he i-&or-alone l i f t - fo r corresponding values of wing angle of a t t ack and ro to r tip-speed r a t i o . This r e s u l t i s not surpr i s ing f o r a biplane l z f t i n g system. Since t h e most tangible fac tor causing t h e difference i s the reduction i n ac tua l wing angle of a t t a ck caused by t h e ro tor downwash, t h e calculated reduction i n wing angle of a t t ack (based on t he theory of r e f . 5) from t h i s source i s shown i n f igure 21. For compari- son, t h e change i n wing angle of a t t ack which would be needed t o account fo r t he e n t i r e l i f t difference i n question i s a l so plot ted. The comparison indicates t h a t r o to r downwash can account f o r most of t he di f ference under some conditions and about half of it f o r t he remaining conditions.

The e f f ec t on t he l i f t - d r a g r a t i o s brought about by adding t h e l i f t i n g ro tor t o t h i s a l r c r a f t i s shown below. The data presented a r e a t t he angle of a t t ack of the complete convertiplane required fo r l eve l f l i g h t a t t h r ee airspeeds. The i so la ted ro tor data i n the two columns belo$r were chosen t o match t he corresponding rotor-tip-speed r a t i o and nominal rotor-blade p i t ch s e t t i n g of t h e complete a i r c r a f t .

20 NACA RM A35K2la

lThe addition of the hub caused air-flow separation on the pylon. These data include both the hub profile drag and the air-flow separation drag.

The l i f t - d r a g r a t i o s of the complete a i r c r a f t were appreciably l e s s than those derived on the bas i s of the sum of the components. This may be a t t r i b u t e d t o t he interference inherent i n a biplane l i f t i n g system.

Theoret ical r o to r l i f t - d r a g r a t i o s calculated from reference 6 a re included above. Considerable discrepancy between t h e experimental r o to r data and t heo re t i c a l predictions i s evident. Two possible explanations f o r t he lower experimental values a r e (1) ro tor blade p r o f i l e drag higher than estimated i n reference 6, and (2 ) r o to r hub and pylon air-flow- separation drag neglected i n t h e t heo re t i c a l calculat ions. It i s believed t h a t t he discrepancy a t 75 knots was due pr inc ipa l ly t o item 1. This was caused by t i p burners, s t r a i n gages, and so on. It can be noted t h a t t h e t heo re t i c a l l y predicted t rend toward higher l i f t - d r a g r a t i o s with increas- ing tip-speed r a t i o was not shown by t h e t e s t data. It i s believed t h a t t h i s was p a r t i a l l y due t o t he increasing importance of the hub- and a i r - flow-separation drag with forward speed of t h e a i r c r a f t . I n an attempt t o account f o r t h i s type of ro to r drag, t he measured hub and air-flow- separation drag (AC~ = 0.0039) was used t o modify t he t heo re t i c a l l i f t - drag r a t i o s . The r e s u l t of t h i s calcula t ion i s compared above with t h e ro-tor-blades and hub data. This comparison indicates t h a t t h e theory of reference 6, when concerned with p r a c t i c a l r o to r s , may be ser iously l imi ted a t t h e higher airspeeds because of t he d i f f i c u l t y i n accurately est imating t h e drag of r o to r components such a s the hub, exposed l inkages, and mounting s t ruc ture .

Airplane-type f l i g h t . - Figure 20 shows t h a t i n t he airplane-type f l i g h t configuration, in terference between t h e f ixed wing and ro tor s t i l l e x i s t s , t h e complete a i r c r a f t l i f t being l e s s than t h e sum of the com- ponent l i f t s .

A summary of t he e f f ec t of t he ro to r i n the airplane-type configu- r a t i on on t h e a i rplane l i f t - d r ag r a t i o s i s l i s t e d below f o r t he angle of a t t a ck of t he convertiplane required f o r l e v e l f l i g h t a t 125 knots.

NACA RM A55K2la

Configuration &.& ( a ) Aircraf t with ro to r blades removed

(includes hub- and a i r - f low- separation drag) 7.2

(b ) Rotor-blades alone 4.0 ( c ) Convertiplane 5 -6 (d ) Sum of t h e a i r c r a f t components

( (4 + ( b ) ) 5 08

It can be seen from the foregoing data t h a t t h e addit ion of even a l i g h t l y loaded ro to r t o an a i rplane w i l l reduce t he a i rplane e f f ic iency s ign i f ican t ly . I n comparison, it can be pointed out t h a t t h e l i f t - d r a g r a t i o s of conventional a i rplanes having t he same gross weight and maximum speed a r e i n the neighborhood of 10.

Ai rc ra f t Drag

A study was made t o determine t he drag contributions of t he various a i r c r a f t components i n order t o f i nd out i f t h e sea-level performance could be improved.

The drag t e s t s consisted of two pa r t s : (1) Attempts t o reduce t he air-f low separation due t o hub-pylon interference, and (2) a general a i r c r a f t clean-up.

Hub-pylon interference.- Flow separation on t h e pylon due t o t he influence of the ro tor hub had been noted i n small-model t e s t s , and t u f t s tudies on t h e fu l l - s ca l e a i r c r a f t confirmed t h e small-model t e s t s i n t h i s respect . The t u f t s tudies showed t h a t t h e separated area extended from t h e top of t h e pylon a t t he ro tor center l i n e down diagonally across t he pylon a s shown a t t he top of f igure 5. Even with unseparated flow, obtained with t he high f a i r i ng and sealed jo in t s , t he t u f t s indicated a downward flow on t he pylon, It i s believed t h a t t h e flow separation originated a t about t he maximum-thickness point on t h e rotor-shaf t housing, between the hub and t he pylon, Modifications made t o the pylon i n an attempt t o stop t he air-flow separation a r e shown i n f igure 5, A l i s t of t he drag increments rea l ized by some of t he modifications i s presented below. In a l l t h e high f a i r i ng configurations, the ro tor blades were removed.

Configuration ~ C D

High f a i r i n g 0.0002 High f a i r i ng , a l l j o in t s

sealed - ,0022 High f a i r i ng , low fence . O O O ~ High f a i r i ng , high fence .0003

22 NACA RM A55K2la

A s ign i f ican t drag decrease was brought about by use of a high f a i r i n g only with a l l the jo ints sealed. That the reduction i n drag resu l ted from elimination of the air-f low separation on the pylon was shown by t u f t s tudies; the high f a i r i n g would not, however, allow the ro tor t o ro t a t e . Thus, it i s not known i f the ro ta t ing ro tor would have caused separation with the high f a i r i n g . I n a.11 cases, fences increased the drag.

An attempt was made t o control t he air-f low separation on the pylon by removing the separated a i r through perforated areas and holes . The 8utorotat ing ro tor was used a s a pump. No decrease i n drag was evidenced by the avai lable suction through the perforated areas . Suction through the holes dezreased the drag coef f ic ien t by 0.0003, but the t u f t s indicated the same separated region. Fences used with the holes brought about a ne t increase i n drag. Although suct ion-ai r quanti ty was not measured, it i s probable t h a t i n su f f i c i en t a i r was drawn through the perforated areas and suction holes. I n addit ion, the perforated areas and suction holes may not have been i n advantageous locat ions .

.- Aircra f t clean-up was the other attempt at drag reduction. I n the clean-up programs, external protuberances were f a i r e d and gaps were sealed. Various components of the a i r c r a f t were removed and t h e i r drag contribution determined. Final ly , the drag contribution of the support system was evaluated. The r e su l t s of the clean-up a r e l i s t e d below f o r the angle of a t t ack corresponding t o l e v e l f l i g h t i n the airplane-type configuration a t 125 knots. Drag coef f ic ien t of t he a i r c r a f t before the clean-up was 0.0230 a t the angle of a t t ack required f o r l e v e l f l i g h t a t 125 knots.

Configuration and components

Rotor blades and hub sealed and f a i r ed Normal pylon, gaps sealed A l l pract icable f a i r i ngs Rotor blades removed Hub removed Tai l - fan assembly, p i t o t s t a t i c mast, wing-tip

skids removed Engine ducting unsealed Skid landing gear removed P i t ch s t r u t and p i tch-s t ru t f a i r i ngs removed L i f t s t r u t s

Of t he component drags l i s t e d above, the ro tor blades and hub account f o r 0.0102 o r 45 percent of t he t o t a l . I f t he hub-pylon interference drag were eliminated, the ro tor blades and hub drag would account f o r 0.0080 out of 0.0208, o r 38 percent of the t o t a l . This i s i n the configuration where the ro tor i s merely an appendage t o the convertiplane. Therefore, i n comparisons of the XV-1 type convertiplane with conventional a i r c r a f t ,

NACA RM A55K2la. 2 3

the inherent high drag of the autorota t ing ro tor system, even with low ro tor l i f t , introduces a severe performance penalty during high-speed operation.

When the landing skids, t a i l - f a n assembly, p i t o t s t a t i c mast, and wing-tip skids were removed, and a l l practicable f a i r i n g s were used, t he reduction i n drag coef f ic ien t was 0.0027. I n addit ion, when the pylon air- flow separation was completely eliminated the t o t a l drag reduction was 0.0047 o r about one-f i f th of the drag exis t ing with the airplane-type configuration a t 125 knots.

Figure 22 shows the improvement i n performance due t o the a i r c r a f t clean-up program i n the horsepower versus airspeed form. The reduction i n drag th8t would be rea l ized i f t he air-f low separation on the pylon were completely eliminated i s not included. An increase i n maximum speed of the a i r c r a f t from 120 t o about 130 knc ;s a t r a ted power i s indicated if the drag coef f ic ien t i s reduced by 0.0 25.

CONCLUSIONS

Results of t he fu l l - s ca l e wind-tunnel invest igat ion ind ica te t h a t t h e XV-1 convertiplane should successfully accomplish t h e autogiro-type and airplane-ty-pe port ions of i t s prescribed f l i g h t plan. Unexpectedly high drag, however, could ser iously l i m i t performance.

I n regard t o longitudinal. s t a b i l i t y , control , and performance charac te r i s t i cs :

1. By proper choice of trim-tab s e t t i ng t he convertiplane can be made s t a t i c a l l y s tab le with respect t o speed and s t i c k motion with t h e center of gravi ty a t t he most rearward posi t ion throughout t he airspeed range invest igated, Longitudinal control should be adequate, provided t h a t l a rge negative trim-tab s e t t i ngs a r e used.

2. The convertiplane was s t a t i c a l l y unstable with angle of a t t ack i n t he autogiro-type f l i g h t configuration. This was with t he center of g rav i ty a t t he most rearward posi t ion.

3. The l o s s i n l i f t due t o interference between the f ixed wing and the ro tor could not be calculated accurately using simple premises. This in terference caused the l i f t - d r ag r a t i o s of the complete a i r c r a f t t o be appreciably l e s s than those calculated from the individual component data.

4. Although the ro tor contributes l i t t l e l i f t during airplane-type f l i g h t , the ro tor blades, ro to r hub, and hub-pylon interference a r e respon- s i b l e f o r 45 percent of the t o t a l drag. I f the hub-pylon interference drag were eliminated, the ro to r and i t s hub trould contribute about 38 per- cent (based on the revised t o t a l drag) of the t o t a l convertiplane drag.

24 NACA RM A55K2la

5 Use of the methods developed in the appendices, along with the downwash theory of NACA TN29l2, permits a reasonably accurate estimate of the tail-contribution to longitudinal stability and control parameters.

Ames Aeronautical Laboratory National Advisory Cormittee for Aeronautics

Moffett Field, Calif., Nov. 21, 1955

NACA RM A55K2la

APPENDIX A

ANALYSIS OF A V-TAB HORIZONTAL TAIL

Consideration of t he longi tudinal s t a b i l i t y and control problems of a convertiplane indicates t h a t a f ree - f loa t ing hor izontal t a i l can be made t o give desi rable contributions t o t he a i r c r a f t ' s longi tudinal charac te r i s t i cs .

A survey of f ree - f loa t ing t a i l configurations ind ica tes t h a t a sa t i s fac tory design from the standpoint of small control fo rces , r ap id response, and freedom from dynamic i n s t a b i l i t y i s one where t he hinge l i n e i s located a t t he aerodynamic center. Although hinging a surface behind t he aerodynamic center contributes more t o longi tudinal s t a b i l i t y , t he analysis herein deals only with t he former arrangement t o avoid t h e complications due t o t h e dynamic problems associated with t h e l a t t e r .

Modifications necessary t o t he hor izontal t a i l hinged at t he aero- dynamic center a r e (1) a servo tab f o r control and (2) a l inkage between t he s t a b i l i z e r and the t abs t o provide a r t i f i c i a l s t a b i l i t y of t h e t a i l with respect t o i t s e l f , It would be convenient, f o r t he sake of con t ro l , t o have a t r i m t ab on t h e t a i l so t h a t control posi t ion could be var ied. I f t he linkage were a l s o connected t o t he t r i m t ab , more pos i t ive s t ab i - l i z a t i o n of t he t a i l would r e s u l t ; consequently, it w i l l be assumed t h a t t he linkage would be used on both t abs . Proper t ies of t h e f r ee - f l oa t i ng t a i l mounted a t t h e aerodynamic center with t he aforementioned controls and linkages a re (1) t h e s t a b i l i z e r incidence constant with angle of a t tack, (2) s t a b i l i z e r incidence dependent only on t he re la t ionsh ip between t he servo and t r i m tabs . Since t he s tab i l i ze r - tab linkage makes t h e t a b angles dependent on s t a b i l i z e r angle, t he s t a b i l i z e r must adjus t i t s i n c i - dence u n t i l an equilibrium posi t ion i s reached. These t a i l cha rac t e r i s t i c s make t h e t a i l s imilar t o an all-movable hor izontal t a i l . For t h i s con- f igura t ion t he linkage between the servo t ab and s t i c k can be adjusted t o give t he proper s t i c k motion. The t a i l , a s described thus f a r , can pro- vide stick-motion s t a b i l i t y .

Speed s t a b i l i t y proper t ies can be introduced i n t o t h e ho r i zon t a l - t a i l system by spring loading t he t r i m t ab so t h a t t he ac tua l tr im-tab def lec- t i o n i s a function of airspeed a s wel l a s the s t a b i l i z e r angle and t ab se t t ing . The r e s u l t of the spring-loaded t r i m t ab i n t h e system i s t h a t t he s t a b i l i z e r angle becomes increasingly negative with increasing a i r - speed, t he magnitude depending on t h e trim-tab spring r a t e .

I n t he usual configuration, t h e tr im-tab spring may not be required since neu t ra l speed s t a b i l i t y i s generally acceptable; however, i n t he case of the XV-1, the locat ion of t h e horizontal t a i l i n t he propel ler

26 NACA RM A35K2la

s l ips t ream might introduce a large des tab i l i z ing e f f ec t on speed s t a - b i l i t y . Therefore, t he trim-tab spring ac'ted i n a d i rec t ion t o improve speed s t a b i l i t y .

On the XV-1, an addi t ional spring was incorporated t o hold t he s t a - b i l i z e r a t an incidence angle of 60' t o 40 knots f l i g h t speed i n order t o avoid l a rge down loads from the ro tor downwash i n hel icopter f l i g h t . A t speeds g rea te r than 40 knots t he t abs became e f fec t ive and t he t a i l had su f f i c i en t aerodynamic moment t o overcome the s t a b i l i z e r downspring.

I n order t o understand b e t t e r t he f ly ing t a i l proper t ies , the govern- ing fac tors a r e presented i n equation form. A s imi la r analysis i s pre- sented i n reference 1. The s t ab i l i z e r - t ab linkage r a t i o i s defined a s

Then ac tua l t ab def lect ion i s , f o r t h e l i nea r range,

The e f f ec t of spring loading t he t r i m t ab can be expressed a s

6 t = 7 6 ~

f o r a given configuration i n incompressible flow

NACA RM A55K2la 27

The hinge moment introduced by t he trim-tab spring can be expressed a s

Furthermore, a t some equilibrium posi t ion, t he hinge moment of t h e t r i m tab must equal t he moment of the trim-tab spring

subs t i tu t ing equations ( ~ 3 ) and (A4) i n t o equations (A?) y ie lds

I n comparison with C h g ~ C! ha i s small. This i s ve r i f i ed by comparison

of the computed curves i n f igure 13 i n which t he simpler equa,tion f o r y

i s compared with equation ( ~ 6 ) . The value of -0.0064 f o r C used f o r t h e computations was obtained from reference 7. h6

An equation f o r the s t a b i l i z e r angle may be obtained by wr i t ing t he equa,tion f o r the moments about t he s t a b i l i z e r hinge l i n e .

= balance moment + ss I- ac,, (::) + gt(=) + downspring moment = 0

( A 8

28 NACA RM A55K2la.

Calculations showed t h a t t he moment due t o t a i l and tab unbalance was neg l ig ib le f o r t he XV-1. I n t he l i n e a r range t he dotmspring moment equals 6.2 + 0.2@jST foot-pound per degree ( f i g . k(b)) . The subs t i tu t ion of equations ( ~ 2 ) ~ ( ~ 3 ) ~ and (Ah) i n t o equation ( ~ 8 ) leads t o t h e following expression f o r 6 s ~ :

Equation ( ~ 9 ) i s evaluated by t he comparison of t he t heo re t i c a l , eqe r imen ta l f u l l - s ca l e , and in terpola ted small-model data showing t h e change i n s t a b i l i z e r angle with airspeed f o r several t r i m and servo-tab combinations i n f igure 14. The value of -0.0044 fo r ( a ~ ~ / a 6 ~ ) or

(acq/ast) was obtained from reference 8 . As can be seen, equation ( ~ 9 ) provides a f a i r estimate of the absolute values of s t a b i l i z e r angle and a good est imate of t h e slope of t h e curve of s t a b i l i z e r angle versus a i r - speed. The lack of corre la t ion was t raced t o the d i f f i c u l t y i n accurate ly determining t h e trim-tab r a t i o , 7 . Subst i tu t ion of experimental values of 7 i n equation ( ~ 9 ) gave calculated agreement within 1/2O.

In view of t h i s , it i s concluded t h a t t he simplif ied analysis pre- sented herein can be expected t o provide r e l i a b l e estimates of t he V-tab f ree - f loa t ing hor izon ta l - t a i l contribution t o longitudinal s t a b i l i t y and control parameters.

The change i n pi tching moment of t h e complete a i r c r a f t with servo- t a b or tr im-tab s e t t i n g may then be calculated by t he usual equation fo r t h e ho r i zon t a l - t a i l contribution t o pi tching moment.

This equation was used t o calcula te t he V-tab f l oa t i ng - t a i l ef fect iveness shown i n f igure 18.

APPENDIX B

ESTIMATION OF ROTOR DOWNWASH

I n order t o estimate t he s t a b i l i t y charac te r i s t i cs of an a i r c r a f t , it i s e s sen t i a l t o lcnow the downwash-field proper t ies with some accuracy, Therefore, an evaluation of t he ex i s t ing downwash theory a s applied t o t he XV-1 type of a i r c r a f t was considered t o be of i n t e r e s t .

Figure l 5 ( a ) compares t he downwash angles calculated using t h e theo- r i e s of referenceB 5 and 9 with t he dolmwash measured from the V-tab f l oa t i ng t a i l i n t he invest igat ion reported herein. Reference 9 allows calcula t ion of downwash through t he ro tor d isk only. Reference 5 extends t h e theory of reference 9 t o allow calcula t ion of t h e downwash angle a t any point i n t he flow f i e l d ,

I n order t o calcula te the resu l tan t ve loc i ty a t any point i n t h e ro tor wake and thus t he angle between t he ax i s of t h e wake and t h e f ree- stream direct ion, t he following must be known: CT, p, and angle of the t i p path plane with respect t o t h e free-stream direct ion. When these a r e known, the v e r t i c a l component of t h e veloci ty a t a desired point i n t he flow f i e l d can be calculated or read from the char ts provided. Then t h e downwash angle can be calculated from

For t h e XV-1, f u l l - s ca l e ro tor data from previously conducted t e s t s were avai lable so t h a t t he rotor-alone l i f t could be accurately obtained fo r a s e t of given conditions. With these experimental values, t h e v e r t i c a l component of t h e veloci ty , w, at t h e t a i l could be calculated and t h e downwash angle a t t h e t a i l due t o t he ro to r determined.

For t h e XV-1, both the ro to r and t h e wing contribute t o downwash a t t h e t a i l , I n order t o allow f o r t h e unloading of t he wing by t h e ro tor , t h e average change i n angle of a t t ack of t he wing due t o t he ro to r down- wash was calcula ted a s outlined above f o r t he case of the t a i l . These calcula t ions a r e shown i n f igure 21. The corresponding downwash due t o t h e wing f o r t h e revised angle of a t t ack of t he wing (af - ew) was found from f igure l5 (b) and added t o t he calculated rotor-alone downwash t o give t h e curves presented i n f igure l ? ( a ) .

A s can be seen i n f igure l ? ( a ) , the fu l l - s ca l e and small-model meas- ured downwash a t the t a i l were i n good agreement, bu t the calcula ted abso- l u t e values show poor agreement with experimental absolute values and only f a i r agreement with respect t o slope. However, s tabi l i tywise , slope i s the most important. I n view of t h i s , it can be concluded t h a t t he theory presented i n references 5 and 9 w i l l provide estimations of the downwash a t the t a i l with su f f i c i en t accuracy t o permit f i r s t - o rde r approximations of the longi tudinal s t a b i l i t y paraxneters of an XV-1 type a i r c r a f t .

NACA RM A55K2la

ESTIMATION OF STICK-MOTION STABILITY

In order t o obtain an indicat ion of t he e f f ec t of trim-tab s e t t i n g on stick-motion s t a b i l i t y f o r both a i r c r a f t configurations considered herein, it was necessary t o extend the l imi ted data obtained i n the wind tunnel. This was necessary since very l i t t l e data were obtained i n t h i s invest igat ion with large negative tr im-tab se t t ings and the propel ler and ro to r operating simultaneously.

As mentioned i n the discussion of stick-motion s t a b i l i t y , t h e curve of p i tching moment a t the l i f t coef f ic ien t required f o r l eve l f l i g h t a t a given airspeed versus airspeed f o r constant control s e t t i ngs can be used t o calcula te t he s t i c k posi t ion f o r t r i m va r ia t ion with f l i g h t a i r - speed. The procedure followed and assumptions employed i n ca lcu la t ing a curve of t h i s type a r e presented below.

The analysis i n Appendix A provides a means of determining t he s ta - b i l i t y contribution of the hor izontal t a i l f o r a range of t r i m - and servo- t ab def lect ions over a range of ve loc i t i e s . The calculated cha rac t e r i s t i c s a r e shown i n f igures 1 4 and 18. The degree of r e l i a b i l i t y t h a t can be expected from the analysis i n Appendix A i s shows by t he comparisons made i n f igures 13 and 1 4 with experimental values. Since the t a i l character- i s t i c s a r e a function of airspeed a t the t a i l , t he addi t ional increment of ve loc i ty a t the t a i l due t o t he propel ler sl ipstream must be considered before applying the r e s u l t s of t h e analysis i n Appendix A t o est imate s t ick-posi t ion s t a b i l i t y a t another trim-tab se t t ing .

Evaluation of Method

I n order t o es tab l i sh a va l id method of calcula t ing the increase i n airspeed a t t he t a i l due t o the powered propeller and i t s e f f ec t on t he pi tching moment, ca lcula t ions were made f o r conditions which corresponded t o avai lable data points. The method of the calcula t ion was a s follows:

1. The veloci ty increment a t the t a i l due t o t he propel ler sl ipstream was calculated by t he momentum method, assuming 25-percent propel ler d isk blockage.

2. A ACm due t o the change i n airspeed a t the t a i l was found from f igures 18(a) and ( b ) .

3. The AC, of item 2 was corrected t o free-stream q .

MACA RM A55K2la 3 1

4. The ACm due t o the propel ler t h ru s t ac t ing on a point away from the moment center was calculated.

5 . The ACmls of items 3 and 4 were added t o the propeller-removed data t o give the f i n a l estimation.

A comparison between the measured data and calculated data i s pre- sented below.

The data at 75 and 100 knots a r e fo r a constant angle of a t t ack f o r each airspeed with a trim-tab s e t t i n g of -20' and a servo-tab s e t t i n g of 10'. The a i r c r a f t was i n t he autogiro-type f l i g h t configuration. The data at 125 knots a r e f o r the same tab s e t t l lgs, but with t h e ro to r blades removed.

Although the agreement between computed and experimental values of p i tching moment was acceptable a t 100 knots, t he agreement was poor a t 75 and 125 knots. This lack of agreement i s believed t o be mainly due t o t h e inadequacy of t he momentum theory used t o p red ic t the average change i n veloci ty a t t he hor izontal t a i l . Based on t h e foregoing com- parison, correction fac tors t o t h e calculated veloci ty a t t he t a i l were obtained f o r each f r e e - s t r e m veloci ty and power s e t t i n g shotm above. The correction fac tors were extrapolated t o t he desired engine rpm.

Estimation of Pitching-Moment Increment Due t o Trim-Tab Deflection

The e f f ec t of changing t h e trim-tab def lect ion from -20° t o -60° with powered propeller was calculated by the following method:

1. Velocity a t t he t a i l was calculated by t he momentum theory and corrected by t h e required fac tor .

2. The ACm caused by a change i n trim-tab s e t t i n g from -20° t o -60' at t h e appropriate t a i l airspeed was found from f igure 18.

3 . The ACm of item 2 was corrected t o free-stream airspeed.

4. The ACm of item 3 was added t o t he powered-propeller p i tching moment f o r 6so = lo0, 6~~ = -20' (shown i n f i g s . 17(a) and ( b ) ) t o give t h e 6s0 = lo0, m0 = -60' p i tching moments.

Estimation of Stick Posi t ion fo r Trim

The s t i c k posi t ion f o r t r i m was calculated using t he above pitching- moment increments. A point by point construction of the .curves i s shown i n f igures 17(c) and (d) . The method was a s follows:

1. The pi tching moments f o r 6s0 = lo0, 6To = -60' of item 4 above

were changed i n order t o reference them t o t he ac tua l t a i l airspeed. The

ACm t o be trimmed out by t he servo tab was then obtained from f igure 18.

2. The servo-tab s e t t i n g f o r t r i m was then found from f igure 18(b) . This was converted t o equivalent s t i c k posi t ion by use of t he equation

32 NACA RM A35K21a

Accuracy of Calculated T a i l Character is t ics

Because of t he influence t h e calcula t ions shown i n f igure 18 can have on t he stick-motion s t a b i l i t y analysis , an indicat ion of t he accuracy t h a t can be expected from f igure 18 i s given below. The data show t h e experi- mental AC,/AV from f igure 16 (b ) and t h e corresponding calculated values.

These comparisons indicate t ha t t he accuracy of equations ( ~ 9 ) and ( ~ 1 0 ) and t he curves i n f igure 18 were suf f ic ien t f o r use i n estimating t h e stick-motion s t a b i l i t y shown i n f igure 17.

NACA RM A55K2la

REFERENCES

1. Hohenemser, K. H.: The Development of a V-Tab Controlled Floating Horizontal Tail for Rotary-Fixed Wing Aircraft. McDonnell Aircraft Corp. Report No. 3371, Jan., 1954.

2. Hohenemser, K. H.: Full Scale Rotor Tests of the Air Force Converti- plane Model XV-1 in the NACA 40- by 80-~oot Wind Tunnel at Moffett Field, Calif. McDonnell Aircraft Corp. Report No. 3379, Feb. 1954.

3. Hohenemser, K. H.: The Characteristics of the Model XV-1 Convertiplane in Airplane and in Autogiro Phase Flight Conditions as ~easured' in the NACA 40- by 80-~oot Wind Tunnel at Moffett Field, Calif. McDonnell Aircraft Corp. Report No. 3599, Nov., 1954.

4. Hohenemser, Kurt: A Type of Lifting Rctor With Inherent Stability. Jour. Aero. Sci., vol. 17, no. 9, Sept. 1950, pp. 555-364.

5. Castles, Walter, Jr., and DeLeeuw, Jacob Henri: The Normal Component of the Induced Velocity in the Vicinity of a Lifting Rotor and Some Examples of Its Application. NACA TN 2912, 1953.

6. Bailey, I?. J., Jr., and Gustafson, F. B.: Charts for Estimation of the Characteristics of a Helicopter Rotor in Forward Flight. I - Profile Drag-Lift Ratio for Untwisted Rectangular Blades. NACA WR L-110, 1944.

7. Sears, Richard I.: Wind-Tunnel Data on the Aerodynamic Characteristics of Airplane Control Surfaces. NACA WR L-663, 1943.

8. Ames, Milton B., Jr., and Sears, Richard I.: Determination of Control Surface characteristics from NACA Plain-Flap and Tab Data. NACA Rep. 721, 1941. (~ormex-ly NACA TN 796)

9. Coleman, Robert P., Feingold, Arnold M., and Stempin, Carl W.: Evaluation of the Induced Velocity Field of an Idealized Helicopter Rotor. NACA WR L-126, 1945.

NACA RM A55K2la

TABLE I.- PHYSICAL C'IW3ACTERISTICS OF THE XV-1 CONVERTIPLANE

. . . . . . . . . . . . . . . . . . . . D i s k a r e a , s q f t So l i d i t y r a t i o ( o r f a c to r ) . . . . . . . . . . . . . . . Blade t w i s t , center t o t i p , deg . . . . . . . . . . . . B l a d e c h o r d , f t . . . . . . . . . . . . . . . . . . . . Mass c o n s t a n t , p e r r a d i a n . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . Area, sq f t Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . Span, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . M.A.C., f t

Incidence of inner panel, deg . . . . . . . . . . . . . Incidence of outer panel at boom, deg . . . . . . . . . Incidence of outer panel a t t i p , deg . . . . . . . . . . Dihedral, deg . . . . . . . . . . . . . . . . . . . . . Sweepback of t h e quarter-chord l i n e , inner panel,

deg . . . . . . . . . . . . . . . . . . . . . . . . . Sweepback of t h e quarter-chord l i n e , outer panel,

deg . . . . . . . . . . . . . . . . . . . . . . . . . Horizontal t a i l

. . . . . . . . . . . . . . . . . . . . . . Aspect r a t i o

Distance from ro to r center l i n e t o t a i l hinge l i n e , f t . . . . . . . . . . . . . . . . . . . . . . . Area, sq f t

Span, f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chord, f t . . . . . . . . Spring constant ( t r i m t ab only), per deg -0.05 f t - l b Overhang . . . . . . . . . . . . . . . . . . . . . . . . Gap . . . . . . . . . . . . . . . . . . . . . . . . . .

NACA RJ!4 A55K2la 3 5

TABU 11.- AEZODYNAMIC DATA WITH POWERED PROPELLER AND ROTOR BWIl)ES REMOVED; v = 125 KNOTS

'-Aircraft free to float

TM

LE

111.- A

ERO

DY

NA

MIC

D

ATA

WITH

POW

ERED

PR

OPE

LL

ER

W

o\

Ro

tor

rpm

178

178

178

178

345

345

345

325

325

1-75

1-75

175

175

350

350

350

175

1.75

1.75

1-75

175

175

175

1-75

1-75

175

1-75

175

1.75

1-75

1.75

175

1.75

1.75

Eng

ine

rpm

2250

2250

2250

2250

1500

2100

2250

1500

2300

2350

2350

2350

2350

1500

2200

2350 0

1500

1550

1600

1650

1700

1.750

1800

1850

1900

1950

2000

2050

2100

2150

2200

2250

2300

ko

nfi

gu

rati

on

1 -

au

tog

iro

-ty

pe

fl

igh

t C

on

fig

ura

tio

n 2 -

air

pla

ne

-ty

pe

fl

igh

t C

on

fig

ura

tio

n 3

- sam

e a

s 2,

bu

t ro

tor

rpm

g

ov

ern

or

ino

pe

rati

ve

2

~ir

cr

af

t fre

e t

o f

loa

t

NACA RM A55K2la 37

(a) Front view.

Figure 1.- Photograph of the XV-1 convertiplane mounted in the Ames 40- by 80-foot wind tunnel.

3 8 NACA RM A55K2la

(b) Three-quarter rear view.

Figure 1.- Concluded.

NACA RM A55K2la

r Moment center

Al l dimensions in f e e t

Figure 2.- General arrangement of t h e XV-1 conver t ip lane .

NACA RM A55K2la

(a) Stabilizer downspring hinge moments.

-80 I L ~ t i c k full forward I / / , I

-60 -40 -20 0 20 40 60 80 100 a,, , degrees

(b) Linkage character ist ics. - Figure 4.- Varia t ion o f downspring hinge moment and t a b angle wi th

s t a b i l i z e r angle.

NACA RM A55K2la

Hiah fai ' pper 6" fence Lower 6" fence

All dimensions in inches

Perforated areas .05"diameter holes 60 to the square inch

Perforated area number I I

Figure 5.- Pylon fence and suct ion configurat ions.

a( ,

degr

ees

Figure 6. - Longitudinal characteristics with the rotor blades and propeller removed and zero tail

load.

NACA RM A55K2la

MACA RM

NACA FiM A55K2la

NACA RM A55K2la

0 .02 .04 .06

.08

.I0

-4

-2

0 2

4

6 8 .012 .008 .

004

0

-.00

CD

a

f ,

de

gre

es

6

,

Figure 8.- Longitudinal characteristics of the autogiro-type flight configuration; propeller

removed.

af ,

degr

ees

Figure 9.- Longitudinal characteristics of the airplane-type flight configuration with the rotor

rpm governor inoperative; propeller removed, V =

113 knots.

.I8

.I 6

.I4

.I2

.I0

CL .0

0

.06

o

10

-20

125

Ro

tor

bla

de

I0

-40

12

5 .7

3 -

.76

.04

0

15

-40

12

5 .7

3-

.76

A

I0

-20

10

0 .5

8-

.61

.02

h

10

-20

12

5 .7

3 -

.76

0

10

-20

10

0 .3

8-

.68

* +

( Gov

erno

r in

op

erat

ive)

0

0

.Q

I .0

2 .0

3

.04

-2

0 2

4

6

8

.004

.002

0 m

.002

-.0

04 -

.Om

-.O

M -.

010

C~

a',

deg

rees

C

m

Figure 10.- Longitudinal characteristics of the airplane-type flight configuration; propeller

removed.

NACA RM A 5 5 K 2 l a

aST , degrees

( a ) V = 75 knots

BST , degrees

( b ) V = 100 knots

. - Aerodynamic c h a r a c t e r i s t i c s of t h e V-tab horizonta

NACA RM A 5 5 K 2 1 a

SST, degrees

( c ) V = 125 knots

-4 -3 - 2 - I 0 I 2 3 4 5 6 8,, , degrees

(d) V = 150 knots

Figu re 12 ,- Concluded,

NACA RM A55K21a

NACA RM A55K21a

60 70 80 90 100 110 120 130 140 150 160 Q, knots

(a) = -20°

Figure 14.- Va.riation of s t a b i l i z e r incidence a.ngle with a i r speed f o r various servo-tab and tr im-tab s e t t i n g combinations.

NACA RM A55K21a

7

6

V)

5 ?2 m (U u

11 4 Y

3

2

Line Data source 0 Ful l scale

I Smal l model (Ref . Ca lcu la ted (Ref . -- Calculated (Ref.

0 -3 - 2 - I 0 I 2 3 4 5

a f J degrees

(a) Autogiro-type flight configuration.

af degrees

(b) Rotor blades removed.

Figure 15.- Comparison of full-scale, small-model, and theoretical down- wash angles at the tail.

NACA RM A55K2la

V. knots (a) Ef fect of angle of attack, propeller removed,

6 s o = ~ 0 0 6 % ~ -20'.

V, knots (b) Effect of control setting, propeller removed ,a f =0:

V, knots (c) Effect of propeller operation, a = 2O, 6glOq 8% = -20°.

Figure 16.- Pitching moment versus alrspeed cha rac t e r i s t i c s ; r o to r blades removed.

NACA RM A55K2la

V, knots

(a) Autogiro-type flight.

V, knots

(b) Airplane-type flight.

Figure 17.- Pitching moments at the lift coefficient for level flight and corresponding stick position for trim-

NACA RM A55K2la

0 25 50 75 100 125 150 V, knots

(c) Autogiro-type flight.

--- 10 -60 Rated

\/, knots

(d) Airplane-type flight.

Figure 17.- Concluded.

60 NACA RM A55K2la

8 ~ o 8 degrees

(a) Trim-tab effectiveness , Bso = 0.

(8) Servo-tab effectiveness, STD = 0.

Figure 18.- Calculated longitudinal-control effectiveness; propeller removed.

NACA RM A55K21a

NACA RM A55K2la

NACA RM A55K2la

NACA RM A55K2la

Figure 22.- Thrust horsepower avai lable and required a t sea l eve l ; gross weight = 5,000 pounds.

NACA - Langley Field, Va.