XH-51A COMPOUND HELICOPTER DESIGN AND DEVELOPMENT

b y

P. W. THERIAULT and D. R. SEGNER Lockheed- California Company Burbank, Cal i fornia

AIAA P a p e r No. 65-75?

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3 XH-51h COMPOUND HELICOPTER DESIGN AND DEYELOPmNT

P. W. Theriaul t a M D. R. Segner

LCCXmED-CAmOPNUI COWANY

"ABSTRACT"

Several import tnt design considerations are discussed r e l a t i v e to the application of t h e

lockheed r i g i d r o t o r concept to a conpound helicopter.

in the adaptat ion of the XH-51A hel icopter to the compound eonfigvration u t i l i z i n g B J-60

j e t engine fo r a u i l i a r ) . thrust are presented.

compound program* included wind tunnel and f l i g h t tests of both three and four-bladed

ro tor configurations. F l igh t tests of the compound configuration through Phase I and

PhaSe I1 are described and results both in term. of measured da ta and handling char8c-

t e l z s t i c s are presented.

w i l l be presented.

Special conditions encountered

Test- undertaken to mpport t h e m-5U

A movie summarizing t h e XS-5lA compound hel icopter pmgrem

INTRODUCTION

Over the p a s t seven years t h e Iockheed-Celifornia Company has been engaged i n the de-

velopment of a r i g i d ro to r system designed to provide a Signif icant improvement i n the

h a n d l i x qYalitie8, speed, and general u t i l i t y of the helicopter.

th is n r k were company wed and covered analytical studies, w i d tunnel tests, free

f l i g h t model tests, and culminated in the design and e o n s h c t i o n of B full scale research

'vehicle, the CL-L75, in d i e h various configurat ions of the r i g i d rotor wstem yere test-

ed i n f l ight .

The inltial phases Of

F0113wlrq the lockheed sponsored f l i g h t development work with t h e CL-L75, two M-5lA re-

Search he l i cop te r s were designed, ConStNcted Bnd f l a w under a contract b d e d j o i n t l y

*The XH-51A conpound he l i cop te r program was conducted under Contract NO. DA-Lh-177-AMC-

1W(T), spon9ored by the u. S. Amy Aviatmc. Wateriel laboratories.

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by t h e U. 5. Amy end the u. S. Navy. These two research machines f w t h e r explored the

p o t e n t i a l Of the r i g i d rotor concept as a i )uE hel icopter . This c o n t r a c t YBS completed

by a m i l i t a r y research evaluation a t the 1'. S. !i'aval A i r Test Center, P a t u e n t , Yaryland

in late 1963.

Subsequently, the i. s. hlmy hviat ion ' l a t e r i e l Laboratories a t Ft. %stis, Va., contract-

ed with t h e I o c b e e d - C d i f o r n i a Coxpny f o r fu r the r research and explorat ion of the high

speed f l i g h t cha rac t e r i z t i c s of the Lockbed r i g i d rotor system. I n p a r t i m l a r , t h e ob-

j e c t i v e of t hese following programs was to emlore t h e Charac te r i s t i c s of the r i g i d r o t o r

when operating u a empound h e l i c o p x r . It is the purpose of this paper to descr ibe t h e

features of t h e r i g i d m t a r pe r t inen t t o eoxpaund f l i g h t , t h e design features of the M-SlA

c0mpoY"d Vehicle, t h e mpport t e s t i n g t h a t was accomplished, and t h e r e s u l t a of t h e f l i g h t

t e s t PPOgram.

FIGm H)TOh rSATJBS

The Ioekheed r i g i d m t o r system is shorn schematically i n Fig. 1.

CantileYered from the hub with a s ingle degree of freedom about the b l a i e spaoitise axis

f o r changing blade angle of a t tack. In addition, a unique control system i s incorpor-

a ted featur ing a c o n t r o l gyro and a spring ca r t r idge arrangement Yhich ;irmides a high

degree Of s t a b i l i t y and c o n t r o l l a b i l i t y for the p i l o t . The d e t a i l design f ea tu res and

c h a r a c t e r i s t i c s of t h e Lockheed r i g i d rotor haYo been described i n previous papers.*

It fea tu res blades

r2n important f ea tu re o f the r izid rotor Bystem appears e spec iaUy a t t r a c t i v e , however,

den Considered relstive to the compwmd mode o f f l i g h t .

a b i l i t y of the r i g i d ro to r t o r a i n t a i n f U 1 con t ro l 7ouer independent of the m-tude

of t h e lift on t h e rotor .

systems i n which t h e con t ro l capab i l i t y Of t h e hinged system i s a direct f m e t i o n of

t h e l i f t on the ro to r .

This feat- relates t o the

This f a c t o r i B i n direct con t r a s t to conventional hinged mtor

t lgure 2 illLqtrates i n s inp le form the fo rce diagrarr, f o r a

*FOP example see paper e n t i t l e d "Results of the XH-51A Rigid Rotor Research Helicopter

?rogram", by W. 3. S t a t l e r , R . h. Heeppe, and E. S. Cruz of Iockheed-CaLiEornia Company.

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!>japed rotor vehicle i n vkich t h e p i l o t has caused a d i f f e ren t i a l i n lift from one side

of t h e rotor to the other by lwvement of t h e cyc l i c control S t i c k

d i f f e r e n t i a l in l i f t is to r o t a t e the r o t o r plane r e l a t i v e t o tire fuselage m d t h e fuse-

lage o i l y responds to the o f f s e t Of the e lift Vector from the Center of gravity.

'The effect of the

For a given r o t a t i o n of tne rotor p l a e , then the response of tts fuse lee is d i r e c t l y

p r o p r t i o l a l to the t o t a l lift on the =OM*.

li fundaremal d i f f e rence WCUI h a w e 7 in the ease of the i r g i d rotor, 'ig. 3. +ere the

d i f f e r e n t i a l in lift acm4s the rotor r e a c t s a5 a moment on the fuselage r e s d t i n g ir: im-

mediate iespanse of t h e vehicle.

%ne response making me contml of t h e vehicle a k c t i o n 0911 of the amount of c y c l i c

blade p i t ch intrcduced. Therefore, unloading t h e r i g i d ro tor i n campound f l i g h t modes

does not a f f ec t con t r c l characteristics and e l i d n a t e s the need f o r a u i l i q aerodma-

$ace , c m t r o l trar.sition 2nd corndination

The magnitude of the t o t a l mtor l i f t does i.at a f f ec t

con t ro l sufaees for high speed rliight.

problem are eli*nated and consis tent control respoilse cha rac t e r i s t i c s =e maintained re-

gardless of t h e r o t o r l if t .

The a b i l i t y t o reduce the Xft On the rOMr to e s s e n t i a l l y zero in high speed f l i r h t has

several e w d y w o r t a n t benefits. First, the co l l ec t ive and cyclic p i t ch angles re-

quired of t h e rotor blades a t high sped are considerably reduced r e su l t i ng in m c h lower

v ib ra t ion levels and blade stresses than would o t k m i s e be possible.

blade stall is eliminated the-bby rewdr% ore of the prislr., livitins facbra an eonven-

t i o n a l machines i n high w e d f l igh t .

t h e r i g i d r o t o r when applied to a compowvi hel icopter configuratirm.

To explore these benefits t h e U. S. Amy Aviation Hateriel laborator ies a w d e d a re-

search con t rac t to t h e Icckheed-CalifOrmia Company i n April 1964 to mdifg one of *&

Iaekheed X&5U he l i cop te r s to a compound configuration.

t h e Xd-51A campound c m f i w a t i o n derired frw the basic m-5U helicopter.

objective of tk f i r s t p n s e of the W-51A compawui prrograa was to ac&eve a k ~ e l n i g h t

speed of 200 h o t s and explore the general fiat c h a r a c t e r i s t i c s of the ve5iele mer

t h e entire speed range.

Seeolll, r e t r ea t ing ,

.U1 of t k s e features iadieated a high po ten t i a l for

Fig. L shows a photograph of

The primary

The semnd ?ham of the pmgrur. called f o r a marimvm lme l

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f l i d h t w e d of 230 b o t s and further explorat ion af t h e s l e d load f ac to r diairam par t i c -

u l a r l y i n the high speed range.

the >ro-rm with the only conf i ewa t ion changes reqvlred involved minor changes t o the

empennage and vindskield sui"ort.

A l l object ives of these tw phases yere exceeded dur ing

X E - S U SC?<.l'liG g?:SI.>N

;onzistent v i m t h e iesem-cn object ives , t i - r m i f i c a t i o n of one of t h e i i -S lA machines

t o the conpcw~d design was t o be accomplished atr!miaun Cost. Stcdies of warn to pro-

v i d e nuxiliiry Coward ?repulsion for a conpound mode of f l i g h t iadicated the simplest

w a n 5 woul? be t o i n s t a l a swnlus j e t meire wit1 its om inde:endent controls.

LPadeoff s t w y vas made of eldsting .let EwiMS per t inen t to t h e sub jec t i n s t a l l a t i o n .

?},e eYalYBtion included tne J69-T-25 witrl a 588 level m i l i t m y t h r u s t of 695 pounds a t

20i knots, the JES-5 with a sea level m i l i t a r y mmst of 1820 powds at 2W h o t s , and

t i e J-60-P-2 with a mil i t a ry thlust of a10 p u d s a t 2 ( 0 h o t s .

vas calcalated t o be 1200 pounds a t 2W b o t s , and 2200 pounds at 2 9 hots.

The s e l ec t ion of the JSO engine ib-5 m a d e on the tollowing basis.

A

The required jet thrust

1.

7.

It was better matched to meed .:oal 0: 230 b o t s t m n the otier j e t engines.

A single enshe would allow one d:ie of t i s i h e l a g e to be m~1UCtered f o r an

emergency ex i t .

h nacelle complete with s t r u c b n a l a t t a c m n t s wa5 avai la~le I i m a T39A anb

could be used wlth a minim. nodif icat icn.

360 engines were available through Amy sowces.

A s ing le JM, ill l i g h t e r t han eitrer tu0 389s OT tlirec S6Y9's.

3.

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5.

The Jbo j e t engine i 8 located on the left s i d e i n a fmuard i o s i t i u r t (Cip . 5) opposite

the p i l o t in order t o allow the p i l o t free egress i n an; emergen-, t o keep the lateral

and longi tudinal cen te r of g ~ a v i t ] l w a t i o n s *thin allowable l imi t s , 2nd t o balance )ax-

ing moment caused by t h e cambered t a i l r o t o r ? y l m a t niz!~ speetis. €m v e r % i c ~ l locat ion

of t h e j e t engine vas selected m e h tha t the j e t t h T U s t i e c t o r i : r ~ R d ~ ri;U.m l i t w i n g

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3 moment about t h e vehicle's cen te r of gravi ty w h i l e , a t the same time, providing reasonable

i s o l a t i o n of the horizontal tail and t a i l ro to r from t h e j e t exhaust.

CaLculatiOns indicated t h e Ying area should be establ ished a t 70 square f e e t t o e s s e n t i a l l y

unload t h e r o t o r completely a t a speed goal of 230 h o t s . (Fig. 6)

taining B l i gh tve igh t w i n g struct- and of mininiaing the ring down load i n hover i n as-

pec t r a t i o of L.05 and a tapeP r e t i o of .S were selected. The wing posi t ion on t h e fuse-

lage was selected t o minimize nacelle wing in t e r f e rence vhile at t h e s- time m a i n t a i n -

ing a relative simple carrythrough s t ruc tu re f o r the wing.

t i ons indicated the center of g rav i ty would be sh i f t ed approximately two to three inches

to the left due t o the a s s m e t r i c a l configuration r e su l t i ng born the i n s t a l l a t i o n Of the

je t engine and correspondingly the center l i n e Of the Wing ~8 located at the cg s l i g h t l y

t o the l e f t of the center li-e of the vehicle .

o f f se t the b a t t e r y and instrumentation :od was mounted on the t i g h t wing t i p . To enswe

autorotat ion capab i l i t y i i t h the campound configuration, wing spo i l e r s were i n s t a l l e d on

the upper surface of the wing a t ap?roaamately 30% Chord so tha t the p i l o t could d e s t m y

the l i f t on the Xing and increase the load on the roto? system f o r t h e purpose of provld-

ing autorotat ion capab i l i t y a t him speeds.

Since the r i g i d r o t o r provides a1 control c a p b i l i t y as explained previously.

I n t h e o r i g i n a l XI-51.4 eor.figuration, B cambered tail rotor pylon (NE. 7 ) was included

to help provide the required anti-torque yawing moment in forward f l i g h t and thus r e l i e v e

the loads on t h e t a i l ro to r i t s e l f .

m-51 compound configuration md the f i n ar'aa i s increased by a sir inch chord extension

aft of the t r a i l i n g edge.

the torque of the main ro to r is low because this rotor i s unloaded.

ing moment then uill be due to t h e j e t t m s t mult ipl ied by i t s l a t e r a l offset .

k m t a t h e yawing moment, due to *,he o f f s e t j e t engine, i s almOst exact ly balanced by t h e

cambered b i l ro to r pylon.

68 pounds, with the m j o r load being carr ied on the tal rotor pylon.

I n the i n t e r e s t of ob-

Xeight and balance calcula-

To further reduce the center Of g r av i ty

io other movable emfaces were incorporated

T h i s cambered t a i l ro to r pylon is r e t a h e d on the

In the conpound c o n f i w e t i o n when operating a t high speeds,

The major uobalanc-

A t 250

For t h i s condition, the tail ro to r thrust requirement is a d 7

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The general arrangement of t h e compound configuration is shown in .Fig. 6.

of the vehicle increased from the basic 361rO lbs. f o r t h e hel icopter configurat ion to L5W

lbs. f o r the compovnd configuration. The basic r o t o r system developed an the X I - 5 1 ~ he l i -

copter uaa employed with no changes &atsoever to t h e system.

recon3 t ha t muing the compound test program the rdor t i p speeds exceeded a Mach nvmCer

Of 0.92 without t h e development Of anY s ign i f i can t problems, in s p i t e of t h e f a c t the

t h e o r e t i c a l t i p c r i t i c a l Mach number is 0.8.

The gross weight

Thus it i s notewmthy t o

s u p m m TESTING

Tests conducted to suppon the campovnd program consisted pr imari ly of three separate pro-

grama.

Of t h e basic tllselage with empennage but without rotors.

program was to develop a snitable fillet betveen the j e t nacelle and fuselage to minimize

i n t e r f e rence &e& The drag d a t a obtained wi th t h e f i n a l f i l l e t c o n f i y r a t i o n indicated

an equivalent f l a t p l a t e area of less than th ree S q u a r e f e e t f o r t h e fuselage, wing, and

nacelle configuration.

t e r n of s t a b i l i t y and t h e required s e t t i n g f o r opthum trim.

Three incidence s e t t i n g s on the hor i zon ta l tail weye tested permit t ing a real is ti^ anal-

y s i s O f t r i m requirements and loads on the s t a b i l i z e r .

shift of the wing centerline of f ive inches t o the left of tne fuselage c e n t e r l i n e xae

optimum t o compensate f o r t h e unsynrmetrical planform of the wing with the j e t engine

mounted an the l e f t s ide.

The f i r s t of these involved conducting a wind b e l test of a U5 scale model

The primary 2urpose of this

The tes t a l so pmvided d a t a on hor i san ta l t a i l effect iveness i n

The d a t a also confinred a l a t e r a l

The Second progran undeitaken involved further e rp lo ra t ion of t h e speed load f ac to r en-

velope of the t h e e bladed ?J(-5lA hel icopter (Fig. 9 ) with p e t i c u l a r eaphas i s on the

e f f e c t s of large l a t e r a l Center of gravi ty o f f se t s on maneuvering ChiU-acteristics of t?e

r i g i d rntor.

f a c t o r s t o 2.05 5'9 nit:;out any unusual changes i n handling c h a r a c t e r i s t i c s .

shows a comparison of l ong i tud ina l s t a b i l i t y a t 65 b o t s f o r three l a t e r a l cen te r of g a -

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l a t e r a l Center af g rav i ty O f f s e t s uP,tO L.5 inches were tested at load

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v i t y positions.

l a te ra l centel of p a p i t y o f f se t s .

f i gu ra t ion due t o a s m t r i c a l engine mounting, therefore, could easi ly be aeconuaodated

my minor changes i n s t a b i l i t y ane evident as t r e s u l t of si , -nif icant

lateral center of gravi ty o f f se t s i n tne co;n?ound con-

based on t h i s data.

The third program concerned an erabation of the four blade rotor system on the XX-5 lA

hel icopter configuration. (Fig. 11) Ground and f l i g h t tests were conducted on this con-

f igu ra t ion t o e s t a b l i s h an ini t ia l basel ine of da t a l d t h the four blade rotor . These data

were used for comparison with the three blade mtw and for ver i f i ca t ion of the se l ec t ion

of the foVr blade system f o r the eom?oud helicopter.

One Of t h e advantages of t h e four blade mtor y d ~ expected to be an inmrw-nt i n maneu-

vering s t a b i l i t y .

f o u r blade m t O r S and ind ica t e s t h e ant ic ipated improvenent was realized.

3ent Was one Of t h e f a c t o r s leading t o the choioe Of t h e four blade rotor f o r the coin-

p m d configuration.

m t o r i n s t a l l e d i s sidlar to t h a t f o r t h e three blade potor.

of the two m m f i p a t i o n s then are sihllar except f o r tne improvement i n S t i ck forces per

?ip. 1 2 shows a campallson of maneuvering s t a b i l i t y f o r the three and

This improve-

C o n t m l sower both l a t e r a l l y and longi tudinal ly w i t h tk four blade

H d L i n g cha rac t e r i s t i c s

g v i t h f o r u a d speed with t h e four blade reto7 noted a k i e .

A number of autorotat ion e n t r i e s were made a t speeds from 60 t o 120 knots.

cha rac t e r i s t i c s were noted during these t e s t s .

a function of normal load factor wa$ invest igated to pmvide basel ine da t a for cozparisan

with t h e compound autorotat ion cha rac t e r i s t i c s wd to develo? g i lo t ing techniques for the

very high Speed autorotat ion e n t r i e s antici?ated i n t h e compound f l i g h t test program.

The reSUlts of t h i s test indioated t h a t control Of a i r s p e d and mtoi speed dvrlng t h e

initial Stages of au to ro ta t ion en t ry a t h i g k r s?eeds i s wre r ead i ly a t t a inab le by hold-

ing t h e co l l ec t ive i n a nearly fixed p s i t i o n while I’egulating mior speed dth load fac-

t o r appl icat ion and airspeed with descent rate.

cyclic con t ro l to keep the mtor speed within prescribed limits and reducing airspeed to

an appropriate range is less r e a d i l y e t ta inable .

NO adverse

hotor ‘pa con t ro l during autorotat ion as

Control by use of Wth co l l ec t ive and

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Other SuoDOrting programs N t h as s t a t i c load measurements, w h i r l tower teste, and labora-

tory ?merams were also conducted as p a r t of the development phase but vi11 not be report-

cd i n this p a ~ r .

LY-SlL C O P O W V K S C L E “ L I X T hCS!iLTs

Although a Specif ic object ive of t h e .W-51A cOnOO,3ound program was to meet or exceed certain

l e v e l f l i g h t speeds and load factors , q u a l i t a t i v e answers t o a number of quest ions from a

? i lo t ing viewpoint were equal ly important. Questions r e l a t i n g to t h e technique of t r a n s i -

t i on from hel icopter to Canpound flight .node, e f fec t s of wing stall on handling Chmacter-

i s t i c s , au tomta t ion entry a t high speed, and maneuvering c h a r a c t e r i s t i c s d high s p e d

canmared to a f ixed wing a i m l a n e canld best be --red by direct p i l o t cmments.

of the following eomen t s are therefore q u a l i t a t i v e i n nature.

Some

Fig. 13 shows the l e v e l f l i g h t perfor-ce of t h e compound configuration. T h i s figure

i nd ica t e s the t r ade off be tvem Collective blade s e t t i n g and e u r i l i v y t h r u s t over the

airspeed envelope.

Sett ings up t o 150 knots.

provided t h e lowest level of blade stresses ani vibrat ion.

held constant for t h e remainder Of the speed extension tests.

t h i 5 co l l ec t ive s e t t i n g e s s e n t i a l l y unlmads the m a i n rotor a t s p d s of 203 knots and

beyond.

a t high speed indicated t h i s s e t t i n g provided e n t i r e l y sa t i s f ac to ry chapac te t i s t i c s re-

quir ing no co l l ec t ive manipulation u n t i l t h e hel icopter had decelerated to conventional

hel icopter autorotat ion airspeeds.

“ig. lh i nd ioa te s t h e speed load f ac to r envelope of t h e compound Configuration.

dicated, a maximum level f l i g h t speed for +.he Compound Of 236 lolots was obtained which

is the f a s t e s t known speed achieved witn any r o t a r y w i n g vehicle. I n addi t ion a m a x i -

mum load f ac to r of 1.9h g’s and a minimun load f ac to r of 0.33 g’s was obtained a t a speed

of 200 knots.

Several tests were n m with Tarious co l l ec t ive and auxiliary thrust

A t this speed a co l l ec t ive p i t ch angle Of 3.h to 3.8 degrees

This s e t t i n g Y ~ S t he re fo re

The data ind ica t e s t h a t

One f i n a l f ac to r influenced this c o l l e c t i v e r e t t i ng . ’“torotation en t ry t e s t i n g

A s i n -

This capab i l i t y r e f l e c t s the advantage of unloading t h e ro tor i n high

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3

speed f l i g h t .

As pointed ou t earlier, the a b i l i t g of the cmpound eonfigvration t o Unload the rotor a t

high speeds has o the r Signif icant benefi ts .

compound hel icopter in tu0 impoptant parameters, nme ly , blade c r i t i c a l bending moment

and cabin vibrat ion.

reduction i n both of these parameters, cornoared to the hel icopter , as the rotor i s un-

loaded i n f o m a d f l i g h t .

x u n d i n the t e s t i n g QO f m is the c r i t i c a l nsch number on the advancing blade of t h e

rotor which t e f l e c t s an increased vibrat ion l e v e l and B lateral trim s h i f t a t the higher

speeds. T h i s e f f e c t may be a l l ev ia t ed by designing t h e blades fo r a thinner t i p a i r f o i l

seet ion or by slowing the rotor ro t a t iona l s p e d .

xaneuvering s t a b i l i t y , i n terns of control forces r e w i r e d to p r d c e normal load f ac to r s ,

was invest igated over a f l i g h t envelope of 100 to 2W b o t s .

Fig. 15 co.?.?aes the pure hel icopter and t h e

A s can be ~ e e n from the test d a t a t h e canpound ind ica t e s B decided

The only l i m i t i r y f ac to r t o forvard speed evident on the con-

These date are shorn i n

Fig. 16. Tuo c h a r a c t e r i s t i c s are significant. First, the cyc l i c control fo rce per g re-

main~ pos i t i ve throughout the range Of speeds invest igated although i t becomes progress-

i v e l y l i g h t e r as g 's are increased. Second. the increased con t ro l s e n s i t i v i t y with air-

speed is also evident. Phis sane cha rac t e r i s t i c h8s been found nore Or less vn ive r sa l ly

in both pure he l i cop te r s and f ixed wing aircraft. The prcblem i s n o t too s ign i f i can t ,

but it does i nd ica t e fu tu re compound he l i cop te r s w i l l r equ i r e some mems of adjusting

con t ro l Sens i t i v i ty t o match the f l i g h t enVMment .

Autorotation entries were invest igated progressively over the f l i g h t envelope up to and

including speeds of 195 knots ca l ib ra t ed airspeed.

d i f f i c u l t y i n maintaining prnper con t ro l of r o t o r rpm.

st rong and maneuvering was performed with ease.

Although t h e r i g i d mtor, i n operating a t near zero l i f t , has an advantage i n that its

drag is l o w md hence rate O f 'pm decay i s reduced, it still requires new a t o m t s t i o n

en t ry tecnniquea a t high speed.

the n i l o t t o deploy the s p o i l e r s a8 =on 85 t h e engine f a i l - yas sensed (or simulated).

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A t all speeds, t h e p i l o t reported no

Control effect iveness -mined

The simplest nethod found dvring these t e e t s was fo r

This permitted the ~ e l a t i v e l y high ro to r angles of a t t ack required f o r autorotat ion to be

developed without pmducing excessive loads on t h e a i r c r a f t . Once t h e s p o i l e r s were de-

ployed, a climbing t w r , was entered to PPoduCe t i e necessary inc rease in angle of a t tack

and t o a s s i s t i n decelerat ing the aiPCrbft to nore comentiorull speeds.

euver, the p i l o t could e a s i l y con t ro l rotor rpn by simply increasing o r decreasing h i s

rate of tum.

entered t a f u r t h e r assist the decelerat ion to lower f l i g h t speeds.

at ta ined, t h e awr i l i e ry thmst was modulated t o Continue flil:ht and penn i t an apt-

landing site t o be selected. k i n g tests sucn as these it was snown that l e v e l f l i g h t

could e a s i l y be maintained and the nel icopter could be operated qu i t e s a t i s f a c t o r i l y as

an B"tOgn-0.

Dvring t h i s man-

Auxiliary engine thmst was reduced to i d l e soon a f t e r the autorotat ion was

Once these speeds were

Blade s t s l l was not not iceable a t any point during the tes t program.

f l i g h t s up to dens i ty a l t i t u d e s of 1 2 , W feet shoued the value of t h e conqowd a,ipmazh.

It i s under these condl t ions that b l a i z stall l imi t a t ions and Snbsequent decw of hand-

l i n g q u a l i t i e s and c o n t r o l ape usua l ly encountered in conventional r o t o r sjstems.

the m-5U conpound, a speed of 263 mph was r e a c k d a t aoproxinately 12,OW i e e t densi ty

a l t i t ude .

level conditions, and aga in there was no apparent blade stall pohlem.

High-altitude

In

S t a b i l i t y and c o n t r o l c h a r a c t e r i s t i c s remained e s s e n t i a l l y uncharged fro* sea-

Although Specific tests were not conducted to evaluate wiwd s t a l l , many conditions were

encountered *ere t h e e f f e c t s c? high angle a t t ack w e r ~ evaluated.

procedure following uei-tical takeoff invo1;ed acce l? r i t i ng t o speeds of 55 - do knots

using j e t t h m s t . A t t h i s speed t h e col lect ive control is lowered t o i t s raise se t t i nz .

The ac t ion i s m c h l i k e r a i s ing the : l a x on air>l?ne ani t.as about the 5- effect .

Some fuselage r o t a t i o n i s necessarj ~o l lowing l0weTin; tne c o l l e c t i v e handle and dming

the program r e su l t an t p s i t i v e angles Of at tack as r i g > as 200 were obtained a t low speeds.

These zigh angles were n o t accompanied by an) of the adverse e f f e c t s u sua l ly associated

with Wing stall. Even i n f l i g h t conditions *ere the s p o i l e r s were extended to destroy

the l i f t on the wing, no unnsual t r L i s.ufts 01. handling c h a r a c t e r i s t i c s yere evident.

The a b i l i t y of the r o t o r a i d wing t o share l i f t appropriate to the f l i g h t condition is

<or exm4le, the

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FIG. 1 MAIN ROTOR CONTROL

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j , . . flG. 8 G E N U L MlANGMNl Cy X H J U COMlOUND

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- XH-51A 4 DUDES

x-' I - 3 sum 4 1.6 1.8

FIG. n UML FLIGHT ERFCWWKE

3-0 A

AIRSSICEWNOIS

FIG. U FUGHT D!ZERMNFD SWD AND LOAD FACE4

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FIG. I5 FLIGHT DATA COMPARISON OF HELKOF'ER AND COMWUND

LOAD FKToIGq'i

FIG. 16 WNEWLUNG S T u l u T y OF COMPOUND CONFIGUUTlON

FIG. l7 XHdlACoUFUIND W FLIGHT

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