Effects of rotor parameter variations on handling qualities of unaugmented helicopters in simulated terrain flight

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, . E ffectsof Rotor Parameter

Vad ations on Handling Qualities of

_. UnaugmentedHelicopters :- in

Simulate d Terrain Flight

Peter D. Talbot, Daniel C. Dugan, , !

RobertT. N. Chen, and RonaldM. Gerdes

// k , _ ; _ August1980 ) National Aeronautics ancl SlOac¢ = Aclmi n_ stration "1 NASA TechnicalMe m orandum81190

Effects of Rotor Parameter

Variations on Handling Qualities of

Unaugmented Helicopters in

Simulated Terrain Flight

Peter D. Talbot Daniel C. Dugan Robert T. N. Chen Ronald M. Gerdes , AmesResearch Center , Moffett F i eld , California

N/ L R A

NatLonal Aeronau t icsand SpaceAdministration AmesResearch Center MoftettF_eld C ahforn_a 94035 i S YMBO L S a blade llft-curve sl o pe a I l o ngitudinal fir s t-harm o nlc flapping co e ff i c ient, r a d BX C lo n gitudinal c y c li c p it c h , t a d (or deg) b 1 lateral f ir st -harm o nic fla[plng c oe ffi c i e nt, tad c blade c h o rd, m e f lapping h i n g e of f set, m hhub rotor hub, height above aircraft center of gravity, m 18 blade moment of inertia about flapping hinge, kg-m 2 k I pitch-flap coupling ratio, _ tan _3 k 8 flapping hinge restraint, N-m / tad LBIC rolling moment due to longitudinal cyclic input, rad / sec2 / deg L rolling moment due to roll rate (roll damping), (sec)-l P L rolling moment due to pitch rate, (sec)-!

q L6 rolling moment due to lateral stick input, rad / sec2 / cm a _IC pitching m o m e nt du e t o longitudinal cy c lic input, rad / sec2 / deg M pitcl_ moment due to roll rate, (sec) -I P M pitch damping, (sec)"'I q M pitching moment due to change in vertical velocity, (m-sec) -!

w M8 blade weight mom e nt about flapping hinge, N-m M6 pit c hing mom e nt due to colle c tive stick input, rad / sec2 / cm c M6 pitching moment due to longitudinal stick input, rad / sec2 / cm e p aircraft roll rate, rad / sec q aircraft pitck rate, rad / sec iii " "!NC P.,_r , BI.' ...... _' " -, ...... l FII..%II : . _ i R r oto r rad i us , m V tr u e a i r sp ee d , m / sec Z vert i cal d am pin g , ( s e c ) - 1 w 7 Lock n u mb e r, _ (oacR_) I B 6 3 pi t ch- fl a p c o upl ing, deg ; n ose-d o wn fea t he r in g with i n c r e a sed flappi n g is po sitive 6 later al s t ic k d eflec t i on, c m a 6 c ol lec ti ve st i c k d e fl ect io n , c m c 6 lon gitud i n al stick de f lec t i o n , cm e 6 pe d al d e f l e c ti o n, c m P e E R 8 p i t c h a t titude, tad (or deg) 8 1 total b la d e twis t (tip w i th r es p ec t to r oo t ) , de g 0 a i r d e n si ty, kg / m 3 o ro tor s oli d it y r a t i o ro ll a ctlt u de, rad (or deg) £ ro t o r- system a ng ul a r veloc it y , ra d / sec ABBR EV IATIONS NOE nap-of-th e e a r th P IO pilot-indu c ed o sc illation S P S P s ho r t peri o d s tab il ity p ara m ete r i v EF F ECTS O F RO T O R PAR A M E T E R V ARI ATI O NS O N HA NDL I NG Q UA L I T I ES OF UNAUGM E NT E D HEL I COP TE k I N S I MULATED T ERR A I N FL I GHT P et e r D. Ta l b ot, Dan ie l C. Du g an , Ro b ert T. N. Ch e n, a nd Ro na ld M. Gerd e s A m en R e s ear c h C e n t e r SUMM ARY A coo r din a ted a nalysis and ground si m u la to r exper im en t w a s perfor m ed t o inves ti gate the effects on single rotor he l icop t er handl i n g qual iti es of sy s - t e m a : ic va ri a ti ons i n the m a i n ro t o r ni nge res t rain t , hub h i nge offse t , p lt ch- f lap coupling, and blade Lock nu m ber. Teetering rotor, ar t icula t ed ro t or and h l ngeless ro t or he li copters we r e ev a lu at ed by research p i lo t s in special low-l ev el flying tasks involvin g obstacle avoidance a t 6 0-1 0 0 knots a i rspeed. T h e results c f the exper im en t a re i n t he form of pilot ra t ings, p i lo t co m mentary and so _ e obje c tive perf o r man ce m easures. Cr i te r ia for da m p- ing and sensi ti vi t y . . e reexamined when combined w it h t he addi ti onal factors of cross-coupling due t o pitch and roll ra t es, p it ch coupli ng with collec tl ve pitc h an d l ong itu d ina l s tatic sta b ilit y . R a tin g s ob tai ned wit h a n d wit h o u t m o t i o n are co m pared.

A cce p t ab le fl y ing q u alitie s we r e o b t a in ed wi t hin each r ot or ty p e b y su ita b l e ad j ust m e nt of th e h ub param e t ers ; how e v e r, pu re t ee t erin g r oto rs wer e found to lack control p o wer for th e task s . A l i mit fo r th e co upli n g p aram e t er __ I L q / L p l o f 0. 35 i s s u gge st ed.

INTROD U CTION C u rre nt ta c ti c s whi c h make use of h e li copters as a n int egr al p a rt of groun d f o rces a nticip a te t he i r u se aga in s t a r mo r an d as a defe n se f or attack h eli cop t ers e m p l oyed aga in s t fr ie n dl y f orces. A s t ro n g e m phas i s ha s been placed on f l yi n g a t l ow al t i t ud e to t a ke adva nt age o f conceal m e nt afforded by vegetat io n and var i a ti o n s in ter r ain h e i ght , B y U .S. A rmy de finiti o n in F ie ld M a nu a l i- I ( r e f . I), t e r r ain fly ing i nc lud es the following mode s of flig h t: I . Low Level - Fl i ght conduc t e d a t a s elect e d altitude, generally over a s t r a i ght route, to m i nimiz e or avoid d e tection o r observation. Ai rs p e ed and indicated al t itude rema i n constant.

2 . Contou r - Flight at low al t itud e con f orming gen er al ly and in clo s e p r o x lmity to the contour s of the e ar th. Ai rs peed and alt it u d e v ary _s v e ge- tation and ob s tacle s dictat e .

3. Nap-of-the-Earth (N O E ) - Fligh t as c lo se t o th e earth's s u rface a s v eg e t a tion o r o b stacles pe rmi t , w h il e generally fol l o wl ng t he ear t h ' s co n - tou rs . Airs pce d a nd al titu de are v ar ied a nd r oute s are w ea ving a nd d e viou s.

T o t a k e m ax imum a d van t a g e o f th e c o ver a nd c on cea lment aff o r d e d by the t e rrain a n d v ege ta t ion , N OE fli g ht h a s b een char a c t er i z e d by ex tr e m e ly low an d s low flight. At Fort Ru ck er, Al ab am a ( th e Arm y Aviation Center ) , NOE i s liter a lly flown in and a mong the trees a nd is c har ac t e rized by we a ving through t he t ree tops at speeds too sl ow t o reg i s t er o n t he ai rspeed in d i cat o r o u t to p erh a ps 2 0 knot s in d i ca t e d a i rspeed.

T h e f l y ing t a sks a sso c i a t e d with t er rain f l y ing p l ace st r ong d em a nds on t he ma neuver in g an d prec i s ion c ont r o l c a p abi l it ies of th e he li c opt er a nd h ave rai se d qu es tions c on ce rnin g t he f lyin g qu a li t i e s n ee d e d fo r such tasks a n d t h e m ea n s to ac hi eve th e m. Among tho se c h arac teri s ti c s f ound to b e d es ir a bl e are: (I ) a d e qu ate c ontrol p ow er a t a ll l eve l s of l o a d f ac tor e n c ount ere d in m a n e uv e rin g fl i g ht ( p o s iti ve a s w el l a s n e g a tiv e); ( 2) a prop e r b a l a n ce of d amping and c ont r ol s e n s itivity in r oll a nd p it c h; a nd (3) a n a bs e n ce of c ou p lin g b e tw ee n axes r e sulting eith er f r om c ontro l in p uts or mo t ions uf th e a ir cra ft. If th e s e c h a r ac t er i s ti c s a r e to be ac hi e v e d without s ta bilit y a ug- m e nt a tion th e y must b e obt a in e d e ntir e ly through va r iation s in m a in a nd t a ll r oto r d es ign a nd oth e r p hysi ca l pa ram e t er s whi c h h a v e a dir ec t influ e n ce on th e ve hi cle 's st ab i l ity a nd c ontrol charac t e risti c s. Obt a inin g a sa tisf ac - to r y l e vel o f flying qu a lities w ithout st a bility a nd c ont r ol a ugm e nt a tion i s d e sir a bl e from th e st a n d point of both c ost a nd re li a bility .

Th e main rotor's c ontribution to th e he l i c op t er's dyn a mi c behavior is s ubstantial: Throu g h vari a tions in th e d e sign of the hub and bla d e s, lar ge c hange s in parameter s which dire c tly affe c t flying qualities s uch a s control pow e r an d dampin g can be obtained. The s e chan g es ar e quite limit e d in th e c a s e of a pure teet e rin g rotor, which d e rives its control moments entir e ly through tilt of th e main rotor thru s t vector about the hub. At the low g -l e vel s e ncounter e d in man e uvering fli g ht, the low thrust l e vels during tran s i e nt man e uvers result in a serious r e duction in available control pow e r.

T he main rotor's contribution to pitch and roll dampin g is derived from it s fl a pping response to a ircraft pitch and roll ra te s. Thi s r esponse offset s th e thru st v e ctor a nd gen e rat e s moments that are proportional to the blade Lock numb e r , _, s o that the helicopter roll and pit c h damping are c orr e spond- i ng ly in f luenced. In the c ase of the teetering rotor, how e ver , both the a bso lute magnitude and the rang e of th e damping that can b e obtain e d are limit e d.

H ingel ess r otor s and rotors with offset flapping hinges are capabl e of g e n e ratin g hub moments that add to and can be mu c h larger than those avail- abl e from thrust tilt a lone. Consequ e ntly, th e y are capable of providin g gre atly au g mente d control power and damping to the helicopt e r. Similar increases can be obtain e d by stiffening the flapping hinge of teetering or off se t hin g e rotors. These direct b e n e fits are attended by incr e ases in couplin g t e rm s su c h as rollin g moment s due to pitch r_te and an unstable con- tribution to th e an g le of attack s tability term, Mw, which may h a ve undesir- abl e e ff e ct s on the handling q ualities of th e helicopter.

To underst a nd t he effects o f these phys i c al p a r am e ter s o n fly i ng qu a l i - ti es i n t h e c on t e xt of t er r a i n f li gh t , a coo rdi na te d a na l ys i s and gr o un d - based s i m ul a ti on exper im en t w as under t aken. T h e genera l obje cti ve o f t h i s exper i men t w as t o m ake an i n iti a l _ xp l ora t ory i nves ti ga tion o f t he t er r a i n fl ig h t reg im e t o provi d e a c o ntr i b uti on t o t he unders t a n d ing of s i ng l e-ro t or, unaug m en t ed he li cop t e r hand li ng qua liti es. A m o r e spec i f ic g oa l w as t o he l p c l ar i fy t he re l a ti onsh ip be tw een var i a ti ons i n t he des i gn o f cer tai n im por- t an t fea t ures o f m a in ro t or geo m e t ry and t he resu lti ng hand li ng char a c t er i s - ti cs. The re sult s of t he a n a l yt i c al s t udy, as repor t ed i n reference 2 , illu s t ra t e t he i nf l ue pce of t he maj or ro t or des i gn fea t ures -- f l app i ng h i nge offse t , _la pp l ng h i nge res trai n t , blade Lock nu m ber, and p it ch-f l ap coup lin g -- o n t he heli :op t er st ab ilit y and c on t ro l charac t e ri s ti cs t ha t w ere expec t ed t o s i gn i f i can tl y i nf l uence hand li ng qua liti es, na m e l y p it ch and ro ll con t ro l effec ti venes s a n d d am p i ng, p lt ch-ro ll cross coup li ng, ver ti ca l damp i ng an d co ll ec tl ve-ya w co u p li ng. Th i s i n f orma ti on prov i ded a bas i s for se l ec ti n g va lu es of t he des ig n parame t ers so as t o appropr i a t e l y vary t he s t ab ilit y and con t r ol charac t er i s ti cs f or eva lu a ti on i n t he s im u l a ti on exper im en t . The eva l ua ti on w as bo t h qua lit a ti ve, t hro u g h s u bj e c ti ve p i lo t assess m en t s, a n d quan tit a ti ve, t hrough t he effec t s of t hese parame t e r s on im por t an t deriva ti ves kn ow n t o h a ve a fu ndamen t a l bear i ng on con t ro l responses or a i rcraf t s t ab ilit y.

The s i mu lati on exper i men t had t hr e e spe cifi c obje cti ves. In an ear li er s tu dy (ref. 3 ), Edenbo r o u gh and W er ni c k e h a d prop os ed d es i rable ran g e s f or t he da m p i n g an d s ens iti v it y in pitc h an d r o ll o f th e h el i cop t e r f o r NOE o p er- ati o ns. Their work, li k e this , con s id e r ed NOE to b e mainly ccntour flying as it is c urrently d e fin ed . On e obj ective of this study was t o investigate c o m- b inations o f da m ping and s e nsitivity c o nsiderably beyond the rang e s discuss e d in ref e r e nce 3 in order to hav e a thor o ugh mapping o f the damplng-sensitivl t y plane in t e r m s of pilot ratings, ea c h point bein g directly identifiable with a spe c i f i c set of rot o r syst e m desi g n parameters. Th e criteria o f r e fer- e n c ¢ 2 ar e illustrated in fi g ur e s ] and 2 from wh i ch the dominant influence s o f hinge restraint, hinge o ffs e t an d L o ck numb e r o n sensitivity and dampin g , as identif i e d in r e f e rence 2, can also b e s e en. Configurations associated with t h i s t e st obje ctive were e stabli s h e d b y s e l e cting appropr i ate com b inations of th e thr e e n o ted design param e t e rs.

A s ec o nd objective was t o inv e stigate the eff ec ts of coupling due to air c raft pit c hin g and r o llin g rates (L q , _ ) on helic o pters which othe r wise had go od damplng-sensitivity characteristics. This woul d help determine what l eve ls o f c o uplin g w e re noti ce able ana ob je cti o nable in th e s e flying tas k s.

Fi g ur e 3 pr e s e nts an e x ampl e fr om r e f e r e nc e 2 that shows that th e influence on pit c h-r o ll c o upling of hin ge offset in combination with Lock numb e r. It is e v id e nt that wi de variations of c oupling ar e ass o ciat e d with these rotor parameters. For an equival e nt flapping frequency, th e contributi o n of hinge r e straint is simila c but l e ss pronounc ed than that due to hin ge o ffs e t. Of cours e , cr o ss- co upllng du e t o cy c li c contr o l inputs can be alter e d through appropriec e phasin g o f pitch and roll cyclic pitch.

T h_ thi r d o b j e c tive w a s t o de t er mine if th e a n g l e of attack in s tability ass o ciat e d with stiff-hln g ed rotors in forward flyin g was o\ J ectiona b le f o r th e sp ee ds and tas k s flown. Wh e th e r or _ot it is ben e ficial to add a l ar ge eno u g h a moun t of pi tc h-flap co u p ling ( 63 ) t o n e u tr alize th e angl e o f a tt a ck i nst abili t y wa s al s o e x amined. Pi tc h-f l ap couplin g al s o c on t ribu t e s t o p i tc h a n d / o r r oll dampi n g (fig. 4) and t o pi tc h-roll c oupling (fig. 5).

The fou r de si g n param ete rs w e r e vari e d o v e r a b roa d r an ge to s imula t e d iff e r e n t rotor s y st ems on a co mm o n fus e lage, t ail rotor and e mpennage.

Ba se d o e t h e analysis, 44 helicopt e r c onfigura t i o n s representin g teetering_ arti c ulat e d and hin ge l e s s r o t o r s w e r e s elect e d an d e valuated in sp e cial tasL s des i g n ed t o be repr ese nta t iv e of t e rrain flig h t. Th e initial par t of t he e x periment wa s p erforme d with two re s ear c h pilots on a flxe d - b a s e simulator, a nd a -u m mar y of tha re s ult s was re p orte d in reference 4.

A f o ll o w-on exper im e nt wa s a lso per f o rme d o n a la rge m ot i o n b a s e s imu- lator w i t h selecte d confi g ur a tions from the i nitlal experime nt that were e va luate d as havin g p arti c ul ar l y g oo d or p arti c ul a rly b a d handli ng qu alities.

T hi s s im u lation permit t ed s ub jec tive evalua t ion s of the s am e c onfi g u r ation s to be made c on s e c u t ively with an d without motion.

Thi s repor t d i sc u sses the res ul ts of t h e ini t ial e x_o_iment and extend s the re s ul ts to c over t he c ompara t ive evalua t ions of six s ele c =ed c onfigura- tions on t he mo t ion ba se s imula t or. Sin c e a la r ge r ange of r o t o r sys t em t ype s c oul d be e x amined in a relativel y shor t period of t ime, it wa s expec t ed t haL some gene r al c on c lu s ion s could b e reached rega r ding t he best flying quali t ie s of an unaugmen t ed helicop t er. It was al s o planne d t ha t t he expe r i- gent would help e st abli s h a good ba s i s fo r choosing t he kind s of augmentation s ys t ems t ha t a r e mo st desirable fo r t erra l n flying t askz.

EXPERIMENT DESIGN Co n f iguration T es t Ma t rix The test configurations that were evaluated c o nsisted of t hree main groups and two subgroups of specially modified configurations. The three main groups are referred to for convenience as teetering (two-bladed, hinge offset zero, spring restraint variable), articulated (four-bladed, hinge offset 5%, spring restraint variable), and hingeless (four-bladed, hinge offset and spring restraint variable). Characteristics of the configurations are shown in table i. Each main group was further subdivided into sets of thr_e with Lock numbers of 3, 6 and 9 within the set. One of the three sets of both the teetering and articulated rotors was a pure example of the type; the other two had increasing amounts of spring restraint added about the flapping hinge in such a way as to maintain the value of augmented rotating natural flapping frequency constant within t he set. The sprinB restraint greatly augmented the hub moment available for maneuvering compared with the pure teetering or articulated rotors. The "hingeless" rotors were approxi- mated by an equivalent hinge offset and spring restraint, also keeping the flapping frequency constant within each seL of three Lock numbers , Natural flapping frequencies varied from a minimum of 1.0 £ for the pure teetering rotor to 1.14 _ for t he stiffest rotor within each group.

B y a dj u st m e n t of t he co n tr o l gea rin g , a b r oad r an ge of v eh i c l e dampi n g and c ontr o l se n sit ivi t i es was achi e ved wi t h the s e 2 / diff e ren t rotor type s as s hown in figure 6. Within the se typ es , large v ariation s in couplin g e x i s t ed , notably roll coupl i ng du e to pi t ch rat e , pit c h c ou v llng due to roll ra te , and pi t ch coupling due to colle c tive p i tch i nputs. In this experiment c ontrol c ro ss -couplln g (e .g., rolling moments due to lon g Jtudln_l co ntrol input s) , which occur s with offse t o r stiffened flappin g hinges, was elimi- na t ed by c on t rol cros s f e ed.

To evaluate the eff e c ts of coupling du e to rol l and pitch rat e s while hol di n g d a mping and se n s itivity co n s tant, four s e t s of three c onfi g ura t ion s w e t , se l ec ted fr o m the main group. The ra t io of damping to control s en s i - tivity wa s mad e c onstant by a dj u s tmen t o f th e i nertia and control g e arin g o f each configuration as shown in figur e 6. T he ratio wa s s elected to li e mid- way b e tw e en th e range s u g g es ted a s op t imum by Ed e n bo rou g h in refer e nc e 3.

Within e a c h se t of thre e c onfiguration s th e damping i n roll and pitch wa s appro x imat e ly con sta nt, but the couplin g param ete r s Lq and Mp varied sig- n i f ic antly, so tha t the ef f ect s of c oupl i ng could b e evaluatea separa te ly.

Th e last group c onsi s t e d of confi g urations with significant values of un s tabl e pitchi n g moment due to angl e of attack, to wh i ch a s uffici e nely large value of pit c h-flat coupling wa s add e d to make th is d e rivative z e ro or s lightly n e gative. Gearin_ and inertia were also ad j ust e d to k e ep the ratios of damping to control sen s itivity cons t an t for each of th e se configuration s .

H e licopter Math Mod e l The helicopter math model (ARMCOP) developed for this simula t ion con- sisted of equations for the separate aerodynamic force and momen t contribu- t ions of the m ain r o tor, tall ro t or, fuselage, fin, and horizontal stabilizer.

The aerodynamics of the fuselage and empennage and the inertlas were based on characteristics of the AH-IG "Cobra" helicopter. Equations for the fuselage were based on those presented in reference 5. Equations fcr the fin and horizontal stabilizer were adapted fro m st a ndard equations for isolated wings of medium aspect ratio. No interference effects were included except those due to main rotor downwash on the horizont a l st a bilizer and tail rotor inflow on the vertical fin.

The tail ro t or model included t hree orthogonal forces with a quasi- static representatio,_ of flapping.

The m o del of _:he main ro t or (ref. 6) was derived from a linearly twis t ed rigid blade with flapping degree of freedo m only, having an offset flapping hinge with a spring restraint a bout the flapping hinge. Rotor speed was assumed constant. Inflow was assumed constant across the disk, and reverse flow, stall and compressibility effects were ignored. A set of differential equations was used to represent the dynamics of the three degrees of freedom of rotor tip path plane motion - coning, longitudinal and lateral flapping.

The tip path plane dynamic equations are identical to th e flapping equation of a three-blad e d rotor system in nonrotating coordinates with p e riodic t e r m s droppe d. The effects o f ai rcr af t an g ula r rates and a cce l er ati ons w e r e in c lu ded i n t he der i va ti on o f t h e f l app i ng equat i o n s, a s desc rib ed in refer- ence 6. The r o t o r forces and m o m e n ts i ncorporated te rms d ue to f l app i ng r a tes, acce l e rati ons, and f l app i ng ang le s.

A limited attempt w as made to vali d ate th e simplifi ed ge ne r ic math m od el (ARMCOP ) for this s tudy. The constant s of the generic math model r e pre s ent- i=g heli c opter geometry and aerodynami c characteri s ti c s were changed to tho s e for a UH-IH, OH-6, and BO-I05, repre se nting, re s pectiv e ly, a t e etering, an articulated, and a hingele ss rotor heli c opter. Th e trim attitudes, control po s ition s , s tability and control derivatives, and eigenvalues were then c om- pared with existing calculated d ata for t h ese helicopter s. Figure 7 show s a c ompari s on of the aircraft trim attitude s and control positions calculated from ARMCOP and C-81 for a UII-IH aircraft. Th e collective s tick and ped a l po s ition s matched rea s onably well, but the air c raft pitch and roll attitude s and th e longitudinal and lateral stick po s itions show some discrepancy in t h e two computer simulation s . S ome maJoz stability and control derivatives from C-81 and ARMCOP are compared in table 2 alon B with thos e of a Bell 2 05 mod e l o b tained from flight d ata u s ing a parameter identification procedure (ref. 7 ) .

These comparisons show that derivative s exhibite d by the AF_CG2 model are compara b le to those obtained from oth e r reliable sources. Reasons for indi- vidual discrepancies have not been pur s ued. It is believed that variation s of these values, with cent e r-of-gravity changes for the ARMCOP model, are similar to varlations obtained from other sources also. Figure 8 shows a comparison of eigenvalues of the six-degree-of-freedom rigid body modes from C-81 and ARMCOP at 60 knots . The basic characteristics of having three pairs of complex roots and two real roots associated with a typical basic teetering rotor belicopter are present, but some slight discrepancies in frequency of the oscillatory modes exi s t in the two computer simulations.

The ARMCOP generated data for simulation of BO-105 configuration gen- erally matched well with Boeing-V e rtol data (ref. 8). The trim attitudes and control positions are shown in flbure 9 to have good agreement. A compari s on of major derivatives at 60 knots is shown in table 3. The significant coupl- ing from collective to pitching moment for hingeless rotor helicopters at forward flight does exist in both computer simulations. The coupled six DOF rigid body elg e nvalues are shown in figure I0. Some discrepancy in frequency between t he two computer simulations does exist, but the basic characteris- ti c s of two pairs of complex roots and four real roots are consistent.

For the articulated rotor helicopter configuzation, the ARMCOP simula- tion of the Hugheq OH-6A also matched reasonably well with the Hughes data (ref. 9).

Simulation Facility Due to facility limitations, simulation fidelity was less thal, deslred for NOE tasks as follows: limited helicopter ma_[_ model in the very low speed region; field of view restrictions, both to the sides and downward; two-dlmenslonal perception of obstacles and clearances; scaling; and lack of rotor disk perception. Within the capabilities of the simulators and visual fa cilitles, a comp ro mi se ha d t o be m ad e a n d ta_ks w ere d e s ig n e d to ev a luate t h e a gillt y of t he h e l i c opt e r. T h e r e sul t wa s a c om b in a tio n o f N O E an d c o n- to ur fl ig h t a t h ig her s p e e ds and a ltit udes t han t h o se n o rmall y ass o c i a t e d wi t h p ur e NOE.

Th e ma j or part o f t hi s st udy wa s p e rformed on a fi xe d-ba s e s i m u_ , v c.

The s imu l ato r ca b c on s i s t e d of a B e ll U H - I A fo r ward f u s elage se c tion _,avi_ th e ori g inal h e licopt er co nt ro l syst e m, includin g h ydrau l ic actuator s at th e s washplate ( fig. II). T h e force di s plac e m e nt c haract e ri s tics, forc e -rel e a se feature and magnetic brak e trim system of th e p e da l s and c yclic c ontrol sti : k w er e th e refore ident i ca l to those of the U H -IA. Th e pilot's instru m ent pa nel i s shown in figur e 12. Torqu e pr es sure and p e rc e nt RPM wer e not r e quired s i n c e con s tant roto r spe e d wa s assum e d for the simulation mod e l. Littl e ref- e r e n ce was mad e to instrum e nta in this task sinc e th e pilot' s att e ntion wa s c on st a n tly focu s ed outside th e cockpit. This is representative of NOE flight a s practiced by th e U.S. Army. For the motion e xp er iment, th e F light Simu- lator f o r Advanced Aircraft ( F S AO was us e d.

Th e FSAA ( f ig. 13) is d es c ribed in d e tail in r e f e r e n c e i 0 . Th e c ab is normally configured f o r fixed-wing transport-type aircraft. For this exp e ri- m e nt th e right-hand s e at was provided with helicopter controls and a basic set of instrum e nts (fig. 14) c onsisting of an altimeter, rate-of- c limb, radi o compass indicator, attltude-dlrector indicator, turn and bank, airspe e d, and engine torque. The pilot's visual display is a c olor TV m onitor with a 420 scan line capability, plac e d behind a collimating lens that both magni- fies the image and makes it appear at infinity.

T he collective stick wa s provided with sufficient static friction to overcome its weight moment about the rotational axis and had light frictional resistance to motion with no force gradient. The force-f ee l chara c teristics of the cycle stick and pedals were provided by a McFadden el e ctro-hydraulic unit with adjustabl e breakout, stati c gradient and viscous damping values.

Th e gradients and c o ntrol travels are shown in table 4. Viscous da m ping was ad j usted until it app e ared satisfactory to the pilots. A force-rel e ase switch on th e panel enabled the stick to b e moved without force gradients and all o wed the helicopter t o b e retrlmmed in a mann e r analogous to th e magnetic brake system on a UH-IH helicopter.

The simulat o r linear mo t ion travel limits ar e ±1.05 m l o ngitudinally, ±12.19 m lat e rally, and ±1.28 m v e rtically. A c omplete d e scription o f th e motion logic is giv e n in reference i0. The general objective of the motion logic is to pr e sent as faithful a reproductio_ of the helicopter linear accelerations and angular accelerations and rates to th e pilot as possible.

A Redif o n visual system was used with a rear-projection system which presented a 48 ° horizontal by 36 ° vertical color television -cene on the back of a translucent screen placed 3.66 m in front of the pilot's window. The terrain model was scaled 1:400 and was based on a section of Hunter-Liggett military reservation in ceutral California. It has natural features such as hills, river beds, and wooded areas combined with roads, telephone poles, vehicles, and other man-made objects to help create some kind of reference scale for the viewer. Three speclal obstecle courses were placed on the model to create repeatable terrain flying tasks for the pilots. A sketch of the course layout is shown in figure 15 and a photograph of part of the c o __ses Js shown in figure 16.

The longitudinal or hurdles course consisted of barriers 15 m high spaced irregularly from 213 to 416 scale meters at 1:600 scale. (The motion was scaled 1:600 for this part of the experiment to give an adequate course length and flying time for the evaluations.) The course was desJgned to focus attention on the longitudinal flying qualities of the hellcopcer, empha- sizing pitch and vertlca_ flightpath control.

The lateral-dlrectlonal or slalom course was a straight line of 6rees spaced slmil_rly to the barriers, requiring the pilots to fly a cur_ing s-path alterna t ing left and right aro u nd successive trees to negotiate _he course, emphasizln£ lareral-dlrectional flying qu a lities.

Trees were placed down the centerllne of a second set of barriers to form a third course combining the fl_ghtpach variations of the other two, referred to as the combination c_.urse.

Except for details, the equipment and pr o cedures of the motion experf ment were the same as for the flxed-base e x periment. The combination course created from the longitudinal course by placing the trees from the s_alom course down its centerline was used for the comparative evaluations. The visual motion scaling was changed from 1:600 to 1:400, so the motion scaling and terrain model scaling were properly matched. The pilots felt that the subjective impressions of speed and altitude were then more like the instru- ment _ndicatlons. At this scale the barriers were i0 m high and SF_ced between 142 and 284 m apart.

Task Description Two research pilots were used for eva2uauing all 2onflguration_ on the flxed-base sim,:lator. Pil o t A had a _aJority o f hours in conventional and V / STOL aircraft and approximately 800 hr of h_ilcopter time. Pilot B had a fixed and rotary wing background which included over 1500 hr in helicopters of many types, both pilots had participated in numerous helicopter and fixed- wing simulator and variable stability aircraft experiments Each pilot was required to fly each configuration through the three courses and give a sep- arate evaluation for each course. In each i tance, the initial condition was 40 knots level flight a_ approximately 18 m a bove ground level. The instructions to the pilots were to fly "as fast as possible and as low as possible" through the course. The longltudinai course was flown first, the lateral-dlrectlona! second, and the combination last in order to allow con- centration on one seL of air,:raft axes at a time before attempting a coordi- na t ed task.

Two additional pilots participated in the moti o n phase of this experi- ment. Piiot C had over 200 hr in helicopters including combat experience and preliminary Army evaluations of pronotypc he!icopL=rs. £ilot D had over 2300 hr in helicopters in a variety of utility missions with some flight test experience in addition. Both pilots also had a large amount of fixed-wing time. The initial condition of the course for the motion ezperiment was 60 knots level flight at approximately 35 m a'ove ground level.

Prior to each Lest period, a few minutes of familiarization were allowed each pilot. Also, each pilo_ was allowed a trial run on each configuration away from the course to get a general impression of its flying qualities independently of the task. Thus, each pilot was individually familiar with the configuration before a final rating was assigned. After completing the course, a pilot rating was given for the task. General and specific comments were written on a Di]ot ques ' .ionnaire (table 5), and voice comments were _ e corded .

The main group configurations were not presented to the pilots in any specific order. For the subgroups (coupling and 63 evaluation) configura- tions related by common values of damping and sensitivity were given as a sequence of three so that effects of the parameter v_ried could be directly compared.

The Cooper-Harper handling qualities rating scale (fig. 17) was used to rate the helicopter and task combinations.

RESULTS OF EXPERIMENT Effects of Damping and Sensitivity Variations in Pitch and Roll Axes The ratings assigned by each pilot to the main group of configurations, for the longitudinal and lateral-directional tasks, are shown with the asso- ciated values of helicopter dampin= and sensitivity in figure 18. "Accepta- bility" boundaries from two references are also shown. The "satisfactory" region for NOE (actual, contour) operation suggested in reference 3 is shown by dashed lines. The upper and lower boundaries correspopd to fixed ratios of control sensitivity to damping and also to fixed values of steady-state r a te response of an idealized first order system to a unit control input.

The lower horizontal boundary is a minimum damping value cutoff.

Despite the broad range of sensi.ivity and damping investigated, th_ ratings of the pilots did not show a corresponding wide variation. For pilot A, satisfactory ratings w e re obtained well outside the bounderies sug- gested by reference 3. Pilot B's ratings were 4.0 or greater for =he la_eral- directiono! task, and some configurations with damping-sensitivity values outside of the boundaries received be=ter ratings than those within the boundaries.

Pilot rating is plotted versus L_ / L__ in figure 19. These dat_ alone would suggest that, fo_ this kind of ta_k ?lateral-directional), the r_tio of damping to sensitivity does not ha ' ,e a primary effect on the handling quail- ties. Pilot A found acceptable configurations across the entire range of L_ / L 6 values tested; pilot B's ratings seem equally insensitive to this r_tio_ The pilot ratings for the longitudinal and lateral-directional courses are plotted versus dam@ing and sensitivity in f_gures 20 through 23. The pilot ratings do not correlate well with control sensitivity for either task.

For pilot A, the longitudinal task ratings appear to be consistently good when Mq decreases below about -3.0 sec-I, but since pilot B's ratings do not show similar behavior, the result may be fortuitous. Pilot A's ratings for the lateral-directional task seem to indicate a cons_._tent variation with LD, with an optimum range for Lp between -i0 and -25 sec-I. Pilot B's r&tings are a little lower in the same range than at the extremes of the test Lp values, but the influence of Lp is not as convincing as pilot A's results suggest.

In general there was a lack of agreement between the two pilots in the numerical ratings that were assigned to the main group of helicopter config- urations, whose characteristics were varied so as to cover a broad range of damping and control sensitivity combinations. One reason is that each pilot used different evaluation criterla for his numerical ratings.

Pilot A used the criteria of maneuverability and suitability of the |eli- copter as a gun platform. His ratings reflected pilot compensation required to correct deficiencies such as low control power, overcontrol tendencx, cross-coupling, or low stability. A rating of 8 or greater reflected a con- trollability problem or near collision; ratings from 5 to 7 indicated high workload; ratings from 2 to 3, a good gun platform. He felt his extreme ratings were most significant.

Pilot B looked for agility and precise control while flying at maximum speeds and lowest tolerable altitudes. His per_option of exaggerated pitch coupling due to collective pitch consistently biased his pilot ratings by i to 2 rating points toward unacceptable. The resulting longitud_nal cyclic pitch changes to correct the coupling overshadowed inputs required for speed changes. The coupling affected performance of the lateral-directional and combination tasks as well.

The pilot comments indicate an awareness of the helicopter's damping and sensitivity and are more in agreement than tbe numerical ratings. Factors affecting the ratings and a comparison of pilot comments are discussed in a special section on pilot comments. To ascertain whether the subjective impressions represented by the written comments were relatable to the sensi- tivity and damping of the helicopters, the configurations with a common characteristic description were plotted on sensitivity-damping plots similar to those used to show pilot opinion ratings (figs. 24 and 25). The categories plotted are io_ sensitivity-sluggish, too sensitive, low dampJng_ and adequate- to-good handling qualities.

There is substantial =greement between the two pilots about the low sensi¢ivity-sluggish configurations, especially in the roll axis. Most of those so noted had very low ratios of sensitivity to damping and consequently i0 ..d a very l o w steady-state rate response per unit of stick deflection. These results, while qu a litatively in agreement with _he Edenborough criterion, indicate that the upper boundary is too _onservative. Those combinations of damping and sensitivity, substantially to the left of the upper boundary and also too near the corner of the damping-sensltivity plot, are to be avoided.

Few e r co n figurations wer e criticized f o r b e ing t oo sen s itive in ro ll a nd pit c h. Pilot A made reference to high sensitivity more frequen t ly th a n pilo t B, especially in pi t c h. Only the c onfigura t ions wi t h the highes t values of sensitivi t y to damping rati o , or very low values of d am ping and sensitivity, were so described. F o r these tasks, b o th pilots seemed to pr e fer a very responsive helicop t er t o one t hat app e ared to be unresp o nsive or h a ving low control power. Th o se configura t ions that were des c ribed as having low damp- ing exclusively or in addition t o other qualities are shown as filled symb o ls.

These p o in t s e mphasiz e the need for adequa t e pi t ch and roll damping and, in the case of roll, that a higher v a lue o f damping than t hat called for by the Edenb o rough criterion may be accep t able.

Those configurations that were either called go o d or deemed to be ade- quate by the lack of negative comments are al s o shown for reference. As was indicated by the pilot opinion ratings, they are found over a rather bread range of sensitivity and damping and often near clearly undesirable config- urations. In pitch, both pilots seemed to find sensitivi t y to damping ratios in the region of the Edenborough criterion or lower to be adequate. In roll, both pilots _ound adequate configurations with higher and lower sensitivity to damping ratios than the criterion. Pil o t B, in particular, did not object to those helicopters with the highest ratios of sensitivity to damping.

These results indicate that the damping-sensitivity region for satisfac- tory handling qualities is broader than that indicated in reference 3.

Because unsatisfactory ratings were obtained even within the boundaries and some very unsatisfactory ratings were found immediately adjacent to satisfac- tory ratings, factors other than damping and sensitivity were evidently sig- nificant to the pilot's perception of the flying qualities.

Both pilots found the pure teetering rotor (KS = 0) configurations to be unacceptable and substantiated these ratings with comments criticizing pri- marily the lack of control power, the low sensitivity and damping, and the low agility. WftL this exception, no particular type of rotor system (augmented teetering, offset hinge and hingeless) was found to be uniformly superior to any other.

The lack of a clear preference for any one type of rotor system is sur- prising since the hingeless rotor was acclaimed at its inception for its superior flying qualities. Part of the reason for its acceptance seems to have been the extremely low time constant and flat rate resp o nse in roll or pitch of the hingeless rotor helicopter. In a British evaluation of the Lockheed XH-51N "rigid" rotor helicopter, mechanical changes were made to the rotor control gyro which directly affected the sensitivity and, to a lesser extent, the damping (ref. ii). The "rate command" nature of the control response seemed to be particularly pleasing to the pilots. It is contrasted II with that of a UH - IH teetering rotor helicopt=r (taken from ref. 12) in figure 26. At low values of sensitivity, however (Mq = -2 see-l; M6e = 0.079 rad / sec2 / cm in pitch; Lp = -7 sec -I, L6a = 0.55 rad / sec2 / cm in roll), the pilots described the XH-51N as "ponderous." This result agrees with descriptions of some of the 300 configurations of this simulation. The m_st sensitive combination tested on the XH-51N in roll was evaluated as being "ever-geared." Configurations in this simulation were not described as too sensitive until the roll control sensitivity was more than twice that of the "o_er-geared" XH-51N.

Effects of Cross-Axis Coupling Coupling between axes was frequently criticized as an additional source of workload that caused the pilots to downgrade configurations. The three types of coupling most often referred to were: p_tching moments due to col- lective pitch inputs, Mac; yaw coupling due to collective inputs, N_c; an_ pitching and rolling moments due to roll rate and pitch rate, respectively (Mp and Lq). Of these, only the pitch and roll coupling due to rates were examined in a systematic manner, and some quantitative estimates for limiting values made.

Pilot B found any noticeable pitch coupling, due to collective inputs, to be highly objectio:lable. A rather large range of this derivative was in the configurations tested, from M_c = 0.008 to M6c = 0.151 rad / sec2 / cm.

The pilot rating did not appear to be related to the value of this derivative, however, as shown in figure 27. Pilot B's comments were reviewed to deter- mine which configurations were criticized for this characteristic. Based on the wording used, comments fell into three rough categories: very exaggerated, large, and noted. Then the values of pitch coupling M6c , the ratio (M6c / Mq) and the ratio (M6c / M6e) were examined to see if their magnitudes corresponded to the categories used to describe the coupling. Both high and low values of these parameters were found in each category. Values of M_c as low as 0.071 rad / sec2 / cm (at 60 knots) were described as "exaggerated," and values as high as 0.122 were described as "reduced." A teetering rotor with a value of 0.0079 rad / sec2 / cm was described as having "a lot." From these results it is not possible to determine what value of collective to pitch coupling is objectionable. Qualitatively the coupling can be a significant source of increased pilot workload. Pilot A also remarked about the presence of this coupling but was either not concerned about it or felt that it actually helped in the performance of the longitudinal task, since the direction of the resul- tant pitch motion was consistent with his technique of pitching up as the barriers were approached and pitching down as they were cleared. Records of speed also showed that pilot A allowed the helicopter to slow while climbing, whereas pilot B attempted to keep the speed constant a task made more diffi- cult by the coupling.

Another significant factor that interfered with the task perform , _L1ce was collective to yaw coupling, N6c. Pilot B fe)t that the simulated helicopter_, had exaggerated collective to yaw coupling and inadequate yaw damping which was most noticeable at speeds below 60 knots. It was considered by h i m to be I a major problem in terms of pilot compensation required, resulting in poor precision in the control of heading and sideslip angle and decreasing confi- dence in his ability to fly low and fast. The yaw dynamics were the same for all the test helicopters as this axis was not changed during the experiment.

In configurations with reduced damping in pitch and roll, however, the dis- turbances induced by collective to yaw coupling were more obvious and inter- fered more with the task. Pilot A also commented on this coupling but did not seem to consider it to be as seriously degrading as pilot B felt it was.

Roll coupling due to pitch rate and pitch coupling due to roll rates were evaluated with 12 special configurations selected from the main group.

By adjustment of fuselage inertias and control gearing, the ratios of sensi- tivity to damping of each configuration were made the same as shown in fig- ure 6 (Series A configurations). The test configurations consisted of four groups of three helicopters with constant pitch and roll damping in each group having a variation in coupling among the three. The pilots evaluated the handling qualities on the combination course.

Pilot ratings are plotted versus the rolling moment due to pitch rate derivative Lq in figure 28. No distinct trend is evident. Good pilot ratings were obtained even at fairly large values of coupling, and some with low values of coupling received poor ratings. According to reference 13 the pitch and roll damping of the helicopter must be taken into account in eval- uating the coupling. The ratios of the coupling terms to the damping • (Lq / Lp and Mp / M_) appear to be more important than the values of the coupling terms themselve=.

The data in figure 28 were rep]otted in figure 29 to show the variation of pilot rating with the parameter Lq / Lp. The pattern of pilot A's ratings supports the hypothesis that the ratio of coupling to damping is the signifi- cant coupling parameter. Pilot B's ratings were almost all unfavorable and did not convey any distinct picture. His pilot commentary on these config- urations indiceted that collective tc pitch coupling was a dominant feature that lo_ered the ratings. Also shown in figure 29 are boundaries discussed in reference 13. The boundaries are meant to indicate that unacceptable ratings (PR > 6.5) are to be expected if Lq / Lp exceeds 0.5 and no better than acceptable ratings (PR > 3.5) can be expected if Lq / Lp exceeds 0.3.

Ratings of all configurations of the main group for beth pilots were examined to see if these boundaries were valid for the results of this experiment.

Generally, the agreement was good. The data suggest that the boundary between satisfactory and acceptable ratings is closer to La / L D = 0.35 than to Lq / Lp = 0.30. The par a meter Mp / Mq was also examined _or-a correlation with pilot raging. Its effect could not be clearly isolated from those due to roll coupling Lq / Lp, since both kinds of coupling occurred simultaneously in these experiments. Values of Mp / Mq as high as 0.35 received a pilot rating of 3. For one configuration with Mp / Mq = 0.75, the pilot noticud extreme roll coupling rather than pitch coupling.

Effects of Pitch-Flap Coup]ing A selected group of five configurations were modified (Series F, fig. 6) by adding 39 ° of pitch-flap coupling (63) to the main rotor. Inertia and con- trol gearing were also changed, as with the A group, to maintain a constant ratio of damping to sensitivity. The value of 6 3 was made large enough so that the angle-of-attack stability term, Mw, was zero or negative. This was thought to be important because of the direct effects of Mw on the longi- tudinal short period stability characteristics. Hingeless rotors make an unstable contribution to Mw which increases with speed. It was thought that reducing the angle of attack instability might have a significant effect on the flying qualities of the stiffly hinged configurations. These configura- tions were evaluated on the combination course.

The pilot ratings and pilot comments for these cases were largely unfa- vorable. The ratings in some cases were worse than configurations from the main group with similar damping and sensitivity and large unstable values of M w.

Pilot ratings for all configurations flown on the longitudinal course were plotted against their respective values of Mw to see if the ratings reflected the variation. No dependency on Mw was evident, and good pilot ratings were given to some cases with the highest unstable values of Mw (to M_ = 0.023 m-sec-l). Having some unstable (positive) value of this derivative, therefore, appears to be acceptable in this kind of task. The addition of 63 to the main rotor also had unwanted side effects in the form of increased pitch-roll coupling (up to Lq / Lp = -1.18) and decreased pitch and roll damping. These factors evidently were more detrimental to the han- dling qualities than any benefit that might have been felt due to decreasing A ground-based simulator study reported in reference 14 established a relationship between the pilot rating and a parameter which represents the "spring" term in the short period mode of the longitudi_.al dynamics -- the so-called "short period stability parameter" (ZwMq - M_). The data of that study showed a rapid deterioration in pilot rating as the v lue of this param- eter approached zero. Calculated values of this parameter 60 knots for the main g_o,_p of configurations and also for the 6 3 group are shown with the corresponding pilot ratings in figure 30. For the 6 3 group, the short period _.ability parameter (SPSP) values are quite low (below 2.0 sec-2), the reduction in M_ due to the addition of 63 being more pronounced than the red,Jction in M_. Pilot A's ratings deteriorate as the SPSP value decreases b , =low 2.0 sec-2. This result is qualitatively in agreement with reference 14, but the deterioration of pilot ratings begins at lower values of the SPSP and the ratings degrade more rapidly than in that reference. Also, pilot B's results are ouite dissimilar. Because of the experiment design, it is not possible to interpret derivative-pilot rating cause-effect relationships unambiFuously.

The approach taken here deliberately chose to compare physical config- urations directly rather than independent variations in derivatives.

Pilot Performance A typical flightpath record for the longitudinal course is shown in flg- ure 31. The minimum ground clearance was limited by a crash pro_ection device on the Re'iron visual system to approximately 6 m. One pilot's perception of speed and scale was distorted, perhaps by the combination of 1:400 modeling scale and 1:600 motion scaling of the visual system. Pilot B commented that occasionally the speed appeared to be one-half of that shownn by the airspeed indicator, and the barriers had the appearance of being much smaller than 15 m high. Pilot A did not comment on this discrepancy.

Time and average height through the courses were used as measures of per- formance. Figure 32 shows the relationship between mean height and course times for the icngitudinal course. Pilot A's neight performance was very con- sistent, and for almost all configurations he flew lower and slower than pilot B. The data did lot show any height-speed tradeoff for either pilot on this course or on the combination course. Records of height versus time were integrated to give a kind of exposure index for the run (a large value indi- cating either a very high mean altitude or very slow speed) to see how expo- sure was influenced by average speed. In all cases exposure was decreased by flying faster. In particular, there was no speed for minimum exposure, where increasing speed and decreasing flightpath excursions resulted in an increased exposure.

Frequently the pilots commented that a configuration with good flying qualities enabled them to fly faster or lower or more confidently. Pilot ratings were plotted versus time to complete the course and mean altitude to verify these impressions. No relationship between pilot rating and course time was found, but for pilot B a definite relationship between his rating and his mea_ altitude performance seemed to exist (fig. 33).

In trying to rationalize the pilot ratings that were given, a number of fllghtpath and control variables were examined to see if they correlated in any way with the ratings. Among these were standard deviations of longitudi- nal, lateral, collective, and pedal control movements, and standard deviations of lateral excursions, heading, altitude, sideslip, and angle of attack.

Except in some extreme cases (e.g., large control motions used for a teetering rotor helicopter with low control power), these measures did not prove to be a good index of pilot rating.

Pilot Comments The differences between the numerical rating3 of the two pilots were dif- ficul_ to reconcile partly because they represented a mixture of deficiencies.

An additional source of information on the differences between the main group of configurations was available from the questionnaires completed by the pilots at the end of each run (table 5).

When the individual pilot comments were reviewed for a few cases with particularly poor agreement (e.g., 301, 308 and 306) it was clear that a "minor but annoying" deficiency (see Cooper-Harper scale) to pilot A might become a "very objectionable" to "major" deficiency to pilot B. Yet both pilots might agree qualitatively on the problem: "Lacks rapid response for NOE" (pilot A) compared to "very poor for NOE -- low sensltivity-low damping" -- for pilot B. The comments may therefore be compared, but the rating numbers have meaning relative to each other only fcr each pilot.

The comments on the questionnaires were reviewed and are briefly summar- ized in table 6, using descriptions that paraphrase the actual commegts written by the two pilots. The comments are segregated into groups relating to longitudinal handling qualities, cross-coupling, and lateral-directional handling qualities.

The data for six configurations - three good and three bad -- that were flown fixed-base in the lateral-directional or combination tasks and had well correlated ratings were examined to see what the pilots most liked and dis- liked about them. They were selected subsequently for the motion experiment.

The "good" configurations (201, A204 and A308) all had fairly high val- ues of damping, with damping to sensitivity ratios near or within the boundar- ies of the Edenborough criterion. Pitch-roll coupling was noticed but not objectionable. Values L. / Lp were less than 0.30. Both pilots liked 201 because they were able to'fly the course fast with it. Pilot B remarked on the good pitch end roll damping of A308. Two configurations had Lock numbers of 3 and one had a Lock number of 6. Both pilots seemed to like A204 best -- pilot A for its good response and insignificant pitch-roll coupling and pilot B also for lack of coupling. Both pilots independently compared A204 to A301 and found A204 much superior.

The "poor" configurations (i01, 301 and 203) were quite dissimilar in their damping and sensitivity. Both pilots complained about the excessive sensitivity of 203 and the associated tendency toward overcontrol and PIO.

This configuration is identical to the popular 201 except for decreased blade inertia that made its LocL number 9 instead of 3. In fact, 201, 202, and 203 represent a steadily worsening pilot rating and an increase in Lock number.

Configuration i01 was criticized for its slow roll response and low sensi- tivity -- pilot A called it sluggish. Configuration 301 was called sluggish by both pilots. Pilot A found 301 hard to coordinate in turns. One pilot considered it to have low sensitivity, the other to have low control power.

Based on these and some other observations it would appear that, to have desirable characteristics in roll, the unaugmented single roto_ helicopter should have L, between -12.5 and -30 sec-1; a ratio of sensitivity to damp- ing between 3.6 and 10.6 deg / sec / cm of stick; and cross-coupling ILp / Lql less than 0.30.

Effects of Motion All of the foregoing results derive from a fixed-base experiment, which may be a significant limitation considering the emphasis on agility and maneuverability in these tasks. Beth pilots indicated that they probably took more risks in the experiment than they might have in an actual helicop- ter, such as approaching very close to trees and using high bank angles at very low levels. A limited motion experiment was therefore conducted which removed additional differences between the simulator and flight tasks.

nesigned and carried out on the FSAA, a large motion simulator, the experi- ment was intended to compare directly subjective impressions of the same con- figurations with and without motion.

For contrast, three configurations rated "good" and three rated "poor" in the flxed-base experiment were selected for the mot_on, no-motlon compari- son. The "good" configurations were 201, A204 and A308; the "poor" were i01, 301 and 203. Each pilot first fl w the helicopter, with the simulator in the fixed-base mode, down the combination course. After assigning a rating and recording comments, the task was repeated with motion.

The pilots felt that the motion fidelity for this sim u l a tion was reason- ably good. However, travel limits (particularly vertical limits) were some- times encountered during a run because of either extreme maneuvers being attempted or because of individual pilot technique in flying.

There was a marked difference in the motion amplitude of the simulator between pilots flying the same task and configuration. With two of the pilots (A and C) the simulator motions appeared to be relatively mild and small, particularly in lateral displacement. The behavior for the other two (B and D) was characterized by impressively large and rapid lateral excur- slo,Ls accompanied by full amplitude heaving motions that carried the simulator to at least the software limits in vertical travel. When these excursions actually resulted in contact with the limits, the resulting false motion cues interfered with the assessment of the motion.

The difference in the pilot rating for a single configuration between motion and no motion never exceeded 1.0, with only one exception, for all pilots. For most configurations the rating with motion was either the same or better than that for the fixed-base run. As with the previous experiment, ratings of the different pilots were often poorly correlated.

d The cooments comparing the fixed-base and motion runs were mixed. Each pilot's comments are reviewed separately.

All of pilot A's ratings improved with motion. He was able to handle all couplings better because of motion cues, particularly collective to yaw coupling. In one case, a tendency to overcontrol in pitch, due to high sen- sltivity or low damping, was reduced with motion. One configuration had a very uncomfortable ride quality, perhaps due to roll-pltch coupling that only revealed itself with motion. He felt he was helped most by motion in the vertical axis (i.e., by the simulation of vertical g forces).

Most of pilot B's ratings improved with motion. He felt the cues were helpful and not misleadlng, particularly the sensation of vertical accelera- tions. A difficult configuration having a combination of low pitch damping combined with strong collective to pitch coupling was more easily controlled with motion than without. He felt that motion caused him to slow down through the course, with an attendant improvement in his ability to perform the task. The data confirm that he flew slightly _1ower with motion than fixed-base, in all cases. In another case, motion did not help in the task performance, but made the lack of damping in that configuration more apparent.

Pilot C's ratings were the same or worse with motion. _i_ general impression of the addition of motion cues was very favorable. In two cases he felt forced to fly the helicopter more gently with motion because of jerky motions resulting from hitting motion stops, making the helicopter more dif- ficult to fly compared with fixed-base operation. Motion cues helped with control coordination, making one helicopter slightly easier to fly in another instance.

Pilot D's rati:_gs were almost all the same or worse with motion. Motion cues were a big help with directional control in one case. In another, the presence of motion contributed toward a PIO tendency. He also complained of hitting motion stops at times. More than one comment seemed to imply that control of the helicopter was more difficult with motion than without.

These comments indicate that some important characteristics were revealed by the presence of motion, and it is therefore desirable to include it in low level helicopter simulations. The small differences between the ratings for fixed-base operation and motion, however, indicate that the essential trends of flying qualities with the dominant aircraft characteristics can be deter- mined quite adequately without motion in this type of task.

Some of the recorded variables were examined to see if there were differ- ences between the fixed-base and motion values, among them average height and speed; standard deviations of cyclic and collective control positions; and standard deviations of pitch attitude, roll attitude and normal acceleration.

For all pilots, average height through the course was about the same with and without motion. For pilot B average speed was consistently lower, by about 5%, with motion, and both higher and lower for the others. The standard deviations of control positions with motion were lower in almost every instance for all pilots, and markedly so for collective pitch for pilot B.

Normal acceleration standard deviation was also much lower with pilot B with motion. Many values were about 60% of the fixed-base amount. For the other pilots, there were no consistent dlffpr=nces for this quantity between fixed- base and motion runs. Th u standard deviations of pitch and roll attitude were similar with and without motion, except for pilet B's roll attitude which was less with motion.

Maximum and minimum values of several variables were also examined, including normal acceleration; pitch attitude, rate and acceleration; and roll attitude, rate and acceleration. For almost all pilots and configura- tions the extremes of the fixed-base values of normal acceleration were greater than those for the motion runs. For pitch and roll attitude, maxi- mums and minimums were about the same for motion and fixed-base. For pitch rate and acceleration, however, fixed-base extreme values were greater than those with motion in most cases. For roll rate and acceleration, values we-e about the same in all cases.

The similarity of the attitudes between flxed-base and motion runs indi- cates that the same kind of maneuvering performance was demanded by the pilots.

The decrease in control excursions, accelerations and rates appears to indi- cate that motion feedback inhibited the pilots from _sing large inputs, but their comments did not show a conscious awareness of this.

DESIGN IMPLICATIONS The damping versus sensitivity criterion for pitch and roll axes is a rather restricted view of the complex flight dynamics of a single main rotor helic o pter.

For the generic math m odel used in this study, the flapping frequency was varied over a wide range to cover teetering rotors as well as stiff, hingeless, single main rotors. The stability characteristics as well as the direct and the cross-coupling response characteristics to control inputs varied significantly as the four-rotor system desJ_n parameters were varied over a wide range, as shown previously in table i. Never t heless, t he damping- sensitivity criterion may serve as a necessary condition for the short-term @irect response requirements in the pitch axis and roll axis individually, especially in the demanding tasks such as those evaluated in this study, wherein less a t tention was pa_d to long-term response characteristics. While being a good candidate for a necessary condition for terrain flight, it is by no means a sufficient condition; many qualifications such as stability char- acteristics of the vehicle and the cross-coupling response characteristics need to be defined to achieve s o me form of necessary and sufficient require m ents.

Another point needing clarification is the quasi-static nature of the damping and sensitivity parameters that were shown in figure 12. With rotor dynamics included, the apparent vehicle damping and control sensitivity can be substantially different from the quasi-static values given in these figures.

Closely related to the damping in roll and pitch are the roll subsidence mode and the pitch subsidence mode (or the "longitudinal short period mode" in the case of a teetering rotor helicopter). The eigenvalues of the roll subsidence mode and the pitch subsidence mode of the coupled 6 DOF zlgld body m ode are approximately equal to the roll damping, L , and pitch damping, M_, respectively. Modal characteristics requirements o_ other rigid body mode_ are less 6..enable to quantification, however. Figure 34 shows the root loci of a teetering rotor (configuration i01), an articulated rotor (201) and a hingeless rotor (301), all with a heavy blade ( 7 = 3). Airspeeds are indi- cated in these figures. The eigenvalues for the Dutch roll, heaving mode, spiral, and the phugoid o f configuration 201 are similar to those for con- fig u ration 301, yet pilots rated 201 much better than 301. The reason for the difference in ratings is attributed to cross-coupling in pitch and roll due to aircraft angular rate and the low contcol sensitivity of 301.

These couplings are shown in figure 35, comparing the aircraft responses u, w, q, e, v, p, _, r to a 2.54 cm step input in the longitudinal stick for configurations 201 and 301. The short-te_m response characteristics in pitch rate and roll rate can be estimated with a reasonable accuracy using the values of M., -M_e / Mq, and (M_e / Mq)(Lq / L p) (as shown in table 7) for the inverse of patch subsidence time constant, short-term pitch rate and roll rate peak, respectively. Note that for configuration 201, the aircraft mildly rolls to the left initially with aft stick input, but for 301 the air- craft strongly rolls to the right instead, as predicted in the sign and magni- tude of (M6e / Mq)(Lq / L p) for the two aircraft. If the value of M6^ / M q were the same for the two aircraft, the roll coupling to the aft stick _nput would have been even more pronounced for 301 as indicated in the sign and magnitude of Lq / Lp. (Note: As discussed previously, a proper control phasing was used, therefore the control couplings L6e and M6a are approximately equal to zero.)

The initial pitch-roll responses of these two aircraft to a step lateral stick input can also be estimated using the values of Lp, -L6a / L p, and (L6a / Lp)(Mp / Mq). The initial pitch rate response to a collective input can also be estimated with good accuracy for these two aircraft using the value of -M6c / M q as shown in table 7.

As indicated earlier, the parameters, L / L_ and M_ / M_ play an important roll in the initial response in tn_ pitch and roll coupling. Their importance has been shown in figure 29. To fully explore these coupling parameters, the values of Lq / Lp at 60 knots were plotted for all the 27 basic configurations, including configurations 201 and 301 discussed above, as shown in figure 36.

As evidenced in this figure, the effects of the pure hinge offset, the flap- ping restraint and Lock number are rather strong and independent. This point was discussed earlier in the paper for the hover case with the main rotor contribution only. The equivalent hinge offset based on the flapping fre- quency for the combination of pure hinge offset and flapping hinge restraint does not serve as a combined parameter to achieve a one to one correspondence to the coupling parameter Lq / Lp for all the three families of rotor systems.

This figure indicates several interesting and important points related to blade inertia and rotor type: i. For a _eavy blade, a high inerti_ rotor system (y = 3) with a large equivalent hinge offset the aircraft will have a strong initial right roll tendency in response to aft stick, even though a proper control phasing has been used (L6e = 0). For y = 3, an equivalent hinge offset of less than 12% should be used for all the rotor systems to keep Lq / Lp _ 0_35. _n optimum range of equivalent hinge offset to minimize Lq / Lp is 3.5% to 6% -.,ith the lower value for teetering rotor with flapping hinge restraint and the higher value for a pure hinge offset (articulated rotors).

2. For a moderately heavy blade (V = 6), the optimum equivalent hinge offset varies widely with rotor systems: 6% for teetering rotor, 11.5% for articulated rotor, and 14.6% for hingeless rotor helicopters.

3. With a light blade (y = 9), t i le o ptimum equivalent hinge offset var- ies even mo_e widely with type of rotor systems: 8% for a teetering rotor at the lower end and substantially higher values for articulated and hingeless rotor helicopters.

Adding a pltch-flap c oupling (in the sense of reducing the blade pitch with up flapping) produces a well known effect of improving the static sta- bility in pitching moment due to angle of attack in forward flight. Adding a large amount (_B = 39° considered in this investigation) produces many poor side effects, however. Effe c ts on the elgenvalues and major parameters dis- cussed in this section are shown in tables 8 and 9,'respectively. Note that the damping in pitch (and also roll) decreases drastically and the magnitude of the pit c h-roll coupling parameters, Lq / Lp and Mp / Mq, incresses substantially.

A recent simulator study by the RAE (ref. 15) on helicopter agility is worth comparing to this one, because of the similarities in tasks and test configuration variables (Lock number and spring restraint) e x amined. The 12 configurations investigated by them (fig. 37) were from m_d to low ratio of ! damping to sensitivity compared to this study (i.e., there were no very slug- gish conflguratlon_). They found that no great preferen c e was shown for any one type of rotor, but that the stiff rotors (high K6) were disliked. One of them (D3) was very close to the AI07 configuration of this study in terms of sensitiv_ y damping, and the value of (K6 / 16_2) (table i0). The pilot ratings were also similar in the two experiments. In this study, some of the stiff hinged rotors (107, 108, 109, 207) exhibited poor flying qualities for the lateral-dlrectlonal task only. Reference 15 also found an occasional wide variation of up to three pilot rating units in repeat evaluations of the same configuration flying the same task by the same pilot. A similarly wide variation in ratings was occasionally seen in this experiment for configura- tions that were almost identical - for example, A's ratings for AI09 and 109 in the combination task.

Another point of comparison was the coupling of collective pitch into pitching moments. This was mentioned in the RAE report as a handling quali- ties problem because of its tendenc y to destroy pre c ise pit c h control in turning maneuvers, by exciting unstable longitudinal modes of the test con- figurations. On a hurdles course, in contrast, their pilots found it notice- able but in the c orre c t sense to aid the anti c ipated maneuver - and therefore, presumably, not especially harmful. Pilot A participated in both experiments.

Our pilots found the low damping, low sensitivity configurations (i01, 102, 103) to be very unsatisfactory. No FAE configurations are quite com- p a rable. The closest ones (A1 and A2) with low K 8 were found to be quite satisfa c tory. They are c omparable to our configuration 106 which our Filots found unsatisfactory in pitch.

The RAE study Identif l ed speed as an important factor affecting the pilots' ratings. Slower flying resulted in "significant improvement in the ratings." Higher mean course speed may thus have been the c ause of pllot B's c onsistentl y poorer ratings in the fixed-base experiment of this stud y .

CONCLU _ I n Tn_ conclusions that follow m, st _sidered in the context of the _as_" Ic-7 level, rel_tively high spe, .,_euv e ring around obstacles. Also these concl u sions are oased on exDeJ : _ : in simulat o rs, both fixed and m o v- ing b a _ obtained from o nly a _ ........ : .

No one ty_e o : ro t or s' : s_e_ : _ .,ifozmly superior to the others for thuse tasks. Goo o [ o adequ a te i f . n g qualities were found in m ore than one member o f each rotor gro u p. A\: pilo t s were, however, un a nimous in downgrad- ing the pure tee t ering rotor cor_ligdra t ions, primarily for having insuffi- cient control power.

The ra t io of control sens l givity to d a mping was used as a guide in selecting test configura t ions and was t hought to be a significant h a ndling qualities parame t er. The results of t his experiment imply that control sen- sitivity / damping is not a strong determinant of pilot opini o n for t hese tasks within the range of values tes t ed. A t minimum, the acceptable range of sen- sitivity and dampi n g is considerably broader than that indicated by t he Edenborough cri t erion. The lev e ls of m inimum damping and mini m u m con t rol power probably are important b u t cannot be determined from these t est resul t s.

Coupling in the form of rolling moments due to pitch rate is importan t .

The absolute v a lue of Lq / L p should be less than 0.3 5 . Collective to pitch and collective to yaw coupling were also very objecti o nable a nd sh o uld be min i mized.

U ns ta ble values of M w (angle-of- a ttack sta t ic stabJ]ity derivative) did no t seem to be objectionable within the range t ested. The SPSP was examined as an index of pilo t opinion rating. At least for o _e pilo t , values of this parameter less t han 2.0 sec -2 resulted in poorer p_lot ratings. These find- in_s tentati v ely support its use as a design pa r ameter.

The evaluation of selected configurations on a motion-base simula t or did n o t greatly a lter t hese conclusions. ._ithough moti o n g a ve iosights and had a definite effect on control mo t ion amplit u des used f o r the t ask, valuable i n forma t ion was gained from t he fixed-base simul a clon.

R E FERENC E S I. D.S. Army Field Manual i-i, Oct i, 1975.

2. Chen, R. T. N.: Effects of Primary Rotor Parameters on Flapping Dynamics, NASA TP-1431, 1979.

3. Edenborough, H. K.; and Wernicke, K. G.: Control and Maneuver _equire- ments for Armed Helicopters. Paper presented at 20th Annual National AHS Forum, May 1964.

4. Chen, Robert T. N,; and Talbot, Peter D.: An Exploratory Invest±_ation of the Effects of Large Variations in Rotor System Dynamics Design Parameters on Helic o pter Handling Characteristics in Nap-of-the- Earth Flight. Paper presented at the 33rd l_at_onal Forum of the American Helicopter Society, Washington, D.C., May 1977.

5. Davis, Joh_ M.: Rotorcraft Flight Simulation with Aeroelastic Rotor and I_proved Aerodynamic Representation. U J D _%MROL-TR-74-10B, June 1964.

6. Chen, R. T. N.: A Simplified Rotor System Mathematical Model for Piloted Flight Dynamics Simulation, NASA TM-78575, 1979.

2. Gould, D. G.; and Hindson, W. S.: Estimates of the Stability Deriva- tives of a Helicopter and a V / STOL Aircraf t from Fligl,t Data.

AGARD CP 172, 1975, pp. 23-1 to 23-9.

8. McLaughlin, J. J.: Stability and Control Datd for the BO-]05 Heli- copter. Boeing-Vertol Rept. D212-I0035-I, May 1975.

9. Siegel, E.: Stability a _d Control Data Summary for Single Rotor He]i- copter, Hughes OH-6A. Hughes Helicopter Rept. 369-V-8010, April 28, 1975 .

i0. Sina c orl, J ohn B., e t al.: Resear c her's Guide t o t he NA S A-_mes Fli g h t S i m ula t or for A d van c ed Aircraft (FSAA), NASA CK-28 7 5, 1977.

Ii. B r o therhood, P. ; and J a m es, C . A.: Some Fli g h t Exp e rimen t s on the XH 5 1N Heli c op t er. RAE Aero. Te c h. M e mo 1342, Au g ust 19 7 1.

12. Talbo t , P . D.; and Corliss, L. D.: A Ma t hematical For c e and Moment Mo d el of a [_-IH He]i c opter for Flight Dynamics Simulations. NA S A TM-73,254, 19 7 7.

13. Huston, R . J .; and W a rd, J. F.: Handlin g Quali t ies and S tru ct ural C har- a c teristi c s of t he _lingeless-Ro t or H_li c u pter. Pro c eedin g s of the V / STOL Air c raft C onferen c e , April 1966.

14. Blake, B. B.; an d Alansk y , I. B.: Stabili t y and Control of the YU H -61A.

J . _m. H e li c o p t e r So c ., v o ]. 22, n o . i, J anuary 1977.

& ...... , .......... A m 15. £adfield, G. D.; Tomllnson, B. N.; and Wells, P. M.: Simulation Studies of Helicop t er Agility and Other Topics. RAE Tech. Memo, Structures 927, FS 197, July 1978.

I I TABLE I. - PHYSICAL CHARACTERISTICS OF THE CONFIGUR A TIONS (a) Configuration test matrix Configuration Y E KB / I8_2 6 3 Remarks i01 3 102 6 0 0 0 Con t rol throws for 103 9 ALL configurati u ns: 104 3 6 = ± 13 . 97 cm 105 6 0 .15 e 106 9 6 = ±13.97 cm a 107 3 6 = 0 - 25.4 cm 108 6 0 .30 c 10 9 9 6 = ±8.26 cm 201 3 P 202 6 0.5 0 203 9 204 3 205 6 .05 .075 206 9 207 3 208 6 .05 .225 209 9 ' 301 3 302 6 .i0 .03 303 9 304 3 305 6 .14 .63 306 9 30 7 3 308 6 .18 .03 309 9 AI09 9 0 .300 A209 9 .05 .225 A303 9 .10 .030 A306 9 .14 .030 Within each group of A205 6 .05 .075 3 configurations, AI08 6 0 .300 damping and sensitiv- A305 6 .14 .030 ity were held constant A308 6 .18 .030 in both pitch and roll.

AI04 3 0 .150 AI07 3 0 .30 A301 3 .i0 .03 A204 3 .05 .075 TABLE i.- Continued (a) Concluded Configuration Y e K6 / 16_ 2 6 3 Remarks L p.

F108 6 0 0.30 38.6 = F205 6 .05 .075

I

F204 3 .05 .075 + ,M < 0 F301 3 .10 .030 I w F104 3 0 .150 _V

J

TABLE i.- Continued _alected stability and control derivatives of the test configurations [Lateral-directional characteristics: mid c.g., 60 knots] L_a, Lp, Lq , N6c , Configuration rad / sec2 / c m sec_ 1 sec_ 1 rad / sec2 / c m Remarks i01 0.154 -2.85 -0.97 0.059 Nr = -1.20 sec-I 102 .153 -1.46 -.49 .061 103 .152 -i.00 -.34 .061 104 1.304 -23.60 8.62 .048 105 .727 -7.38 1.00 .062 106 .535 -3.73 .14 .065 107 2.464 -31.69 24.66 .022 108 1.300 -11.99 4.32 .054 109 .915 -6.17 1.28 .063 201 1.517 -18.35 -.71 .056 202 .911 -5.70 -2.09 .060 203 .709 -3.00 -1.78 .061 204 2.741 -29.20 8.58 .044 205 1.524 -9.37 -.56 .058 206 1.124 -4.75 -1.26 .061 207 5.209 -36.57 28.98 .015 208 2.749 -14.88 4.18 .047 209 1.938 -7.82 .44 .058 301 1.117 -39.06 14.38 .032 302 .605 -13.58 -1.28 .021 303 .436 -6.91 -2.45 .059 304 1.550 -46.92 29.13 .012 305 .817 -19.53 1.00 .044 306 .576 -10.31 -2.50 .055 307 1.158 -27.77 21.03 .005 308 .605 -14.44 i.ii .015 309 .736 -14.32 -1.49 .048 AI09 .582 -5.09 1.05 .063 A209 .316 -2.79 -.08 .031 A303 .329 -2.90 -1.52 .032 A306 1.176 -10.23 -2.48 .056 A205 1.185 -10.28 -.62 .023 AI08 1.172 -10.23 3.68 .054 . . _05 2.364 -20.55 1.04 .044 A308 2.342 -20.41 4.53 .036 AI04 2.362 -20.53 7.49 .050 ' TABLE i.- Continued (b) Continued L6 a , Lp, L q, N6 c , Configuration rad / sec2 / c m sec_ I sec- I rad / sec2 / c m Remarks AI07 3.551 -30.87 24.03 0.023 N r = -1.20 se_-I A301 3.526 -30.65 11.28 .038 A204 3.670 -30.91 9.09 .043 FI08 . 500 -_._5 5.04 .020 F205 .507 -4.35 2.33 .033 F204 1.246 -10.85 10.7! .033 F301 1.262 -10.98 6.33 .02_ FI04 i. 0 -10.91 12.84 .043 _ (b) Continued [Longitudinal characteristics: mid c.g., 60 knots] M6e , M M, M SPSP, q' . Mbc' w Configuration rad / sec2 / c m sec_ I se , _ 1 rad / sec2 / c m (m_sec)_ 1 sec_ 2 i01 0.036 -0.665 0.207 0.007 -0.021 i.i01 102 .036 -.397 .102 .007 -.021 .921 103 .036 -.307 .067 .007 -.021 .861 104 .327 -5.02 -1.732 .112 .009 3.120 105 .183 -1.640 .194 .070 -.002 1.169 106 .134 -.881 -.028 .052 -.007 .818 107 .612 -6.724 -5.017 .156 -.023 5.204 108 .327 -2.607 -.874 .I17 .012 1.390 109 .231 -1.393 -.259 .089 ,004 .821 201 .380 -3.934 .171 .076 -.0004 2.650 202 .229 -1.294 .439 .048 -.008 1.112 203 .178 -.730 .371 .038 -.011 .819 204 .270 -6,207 -1.733 .122 .012 3.786 205 .384 -2.062 .127 .079 .0008 1.356 206 .282 -1.097 .265 .061 -.004 .862 207 1.288 -7.758 -5.912 .158 .023 4.485 208 .691 -3.218 -.846 .128 .015 1.695 209 .489 -1.741 -.085 .]01 .008 .931 301 .279 -8.287 -2.916 .142 .018 5.007 302 .153 -2.952 .275 .I00 .007 1.770 303 .iii -1.554 .512 .077 .001 1.015 TABLE i.- Concluded (b) Concluded M6e, Mq M , _M_c M SPSP ' p ' w ' C on f ig ura tion rad / s e c 2 / c m s e c- 1 s ee - 1 rad / s e c 2 / c m (m _s e c)_ I sec- 2 3 04 0.384 -9.951 -5.937 0.155 0 . 002 6.000 305 . 2 06 -4.207 -189 .130 .015 2.354 3 06 . 1 46 -2. 1 72 .5z C .103 .008 1.263 307 ,284 -5.952 -4.289 .086 .005 3.824 308 .152 -3.150 -.216 .089 .006 1.919 309 .186 -3.119 .312 .129 .016 1.602 AI09 .171 -1.503 -.279 .096 .004 .885 A209 .094 -.870 .028 .049 -.003 .672 A303 .097 -.908 .420 .044 -.007 .816 A306 .344 -3.003 .688 .137 .011 1.669 A205 .343 -3.003 .184 .115 .001 1.975 AI08 .343 -3.007 -1.009 .135 .013 1.604 A305 .572 -5.008 -.225 .154 .018 2.803 A308 .571 -5.007 -1.049 .139 .020 2.746 AI04 .572 -5.002 - -!. , 25 .112 .009 3.060 AI07 .802 -7.004 -5.226 .162 .024 3.965 A301 .898 -6.998 -2.462 .120 .015 4.228 A204 .838 -7.008 -1.956 .137 .014 4.274 FI08 .116 -1.020 -1.062 .040 -.006 .884 F205 .117 -1.016 -.470 .344 -.011 .848 F204 .273 -2.379 -2.214 .045 -.006 1.494 F301 .272 -2.379 -1.264 .058 -.009 1.530 FI04 .273 -2.381 -2.63R .051 -.007 1.810 I TAB L E 2. - A COMPA R ISON OF A RMCOP G E NERATED DE R IVATIVES FOR A UH-]-H WITH OTHER DATA SOURCES ARMC O P Bell C- 81 NR C Bell 205 Derivative J60 knots 80 knots 60 knots 80 knots V = 70 knots I param, i.d., Xu, sec-I -0.017 -0.0202 -0.024 -0.034 -0.1117 Xw, sec-I .032 .035 .012 .057 .0084 _, m-sec -I .610 .615 .043 .053 1.064 Zu, sec-I -.0048 .0102 .066 .079 -.009 Zw, sec-l -.696 -.775 -.875 -.946 -.875 Zq, m-sec -I -.070 -.152 .036 -.058 .427 Mu (m-sec) -I .0049 .0039 .121 .127 .020 _7 (m-sec)-I -.033 -.046 -.121 -.203 -.022 Mq, sec-I -.556 -.633 -.523 -.612 -.848 Lp, sec-I -1.436 -1.445 -.987 -1.002 -.806 Np, sec-I -.289 -.278 .132 .0175 -.037 Lr, see-I .282 .293 -.741 -.704 .1 7 4 !N r, s ec - I - 1.0 7 7 -i.149 -1.42 -1.64 -1.303 iL v (m-sec) -I -.026 -.020 -.499 -.62 -.048 Nv (m-sec) -I .091 .096 .066 .088 .058 L6_ I (sec2_cm)_ I .218 .219 .206 .20 7 .iii N 6 a_ . 0 33 .033 .0 0 14 .001 .015 Ldr 1 (sec2_cm)_ I .104 -.114 -.439 -.506 -.102 N6r, .284 ._09 .589 .678 .194 TABLE 3.- A COMPARISON OF ARMCOP GENERATED DERIVATIVES FOR A BO-105 WITH DATA OF REFERENCE 7 ARM C OP BO - 105 B oelng -Ve rt o l B O - 7 . 0 5 Der i vatives 60 knots 60 kn o ts Xu, sec-I -0.0339 -0.0338 Xw, sec-I .0128 .0311 _, m-sec -I .607 .6 3 8 Zu, sec-I -.0362 -.0 5 63 Zw, sec-I -.6568 -.7885 Zq, m-sec -I .250 .056 Mu (m-s e c)-I .048 .059 Mw (m-sec) -I -.020 .042 MMq, sec-I -3.3077 -3.6151 6e (sec2-cm) -I .321 .392 M6c (sec2-cm) -I .149 .203 Lp, sec-I -8.46 -9.35 Np, sec-I -.7119 -.022 Yp, m-sec -I -.662 -.716 Lr, sec-I .1151 -.0251 Nr, sec-I -.8849 -.6627 Yr, m-sec-I .277 .181 Lv (m-sec)-I -.022 -.226 Nv (m-sec)-I .119 .083 Yv, sec-i -.1469 -.091 L6a (sec2-cm) -I .894 1.03 N6a (sec2-cm) -I .092 .012 Y6a' m-(sec2-cm)-I .0766 .0926 L6r (sec2-cm) -I .369 -.426 N6r (sec2-cm) -I .667 .581 y6r , m-(sec2-cm) -I -.246< -.2081 T A B L E 4 .- HE LICOP T E R CO NTR OL TR AVEL S AN D F O R C E GR AD IENTS Bre akout, C o n t r o l Tr a v e l , [ Gradi e n t, cm N / cm N (appro x. ) C o l l e c t i ve 0 -25 .4 0 2.2 2 P e dals ±8.26 3.50 8.90 Long i tud i nal c yclic ±13.97 2.92 4.45 Lat i tud i nal cycl i c ±13.97 1.75 4.45 3 2 * - ............... m_., . -_ ................. _ i..............

.... r ' _.................. A, - - m _ _ • • . ,_ .F .... _ _ P r - n _ .......... - .. ' .... w,, TABLE 5.- PILOT QUESTIONNAIRE Longitudinal task Lateral-directional task I. Overall Cooper rating i. Overall Cooper rating 2 . V ertic a l response to collective 2. Yaw response to pedals a) Sensiti v ity? a ) Sensitivity?

b) Damping? b) Damping?

3 . Pitch response to longitudinal c) Speed o f response?

_,clic 3. R o ll response to lateral cyclic a) Sensitivity? a) Sensitivity?

b) Damping? b) Damping?

c) Speed of response? c) Speed of response?

4. Coupling 4. Symmetry of response a) Roll-pitch? 5. Coordination of stick, pedals b) Collective-yaw? and collective required?

5. Dynamic stability 6. General comments • General comments Combination task i. Overall Cooper rating 2. General comments

I m m i l u in n m

TA B LE 7 .- A COMPARISON OF SOME MAJOR PARAMETERS FOR THREE TEST CONFIGURATIONS [Mid c .g., 60 knots] Conflgura_lon P arameter - : - - -= - lOl 201 301 , .... , . : -- , , .

!

13 3 3 ¢ ( 0 0.05 0 . i0 KB / (IBfl2) i0 0 .03 M6 e, r a_ / sec2 / cm I 0 . 035 0 . 381 0.280 M_, sec-I Z.- . 67 -3.93 -8.29 -M__ _Mq, rad / sec / cm [ .053 . 097 .034 -.018 -.004 .012 (M_L ql eZMq)(Lq / Lp) ,_R ' . 341 . 044 - . 364 rad / sec / cm L6a, rad / sec2 / cm 0.150 1.468 1.083 Lp, sec-I -2.70 -17.85 -38.00 -Lda / L p, rad / sec / cm .056 .082 .028 Mp / Mq -.313 -.043 .3 5 1 (n6a / Lp)(Mp / M q) .017 .004 -.010 M6 c ' rad / sec 2 / c m i 0.008 0.0 7 5 0. 1 42 -M_c / Mq, rad / sec / cm , .011 .019 .01 7 Pi l o t r a ting (comb. c ourse) Pilot A 8. 5 3 7.5 Pilot B 7.0 5 7.0 I • • • • • 0 I cN ,-- I o' _ '4 :) I ° / I l _ 0 O c ,,I oO r,') o " ) I' _ Z .- ,1 " • • • • • I e [-..i I P-I _ °

_ _, ' _ " _,. , -'1 ," I I

u m 00 _ D u"l _ _I D "_ 0 • , I_, 0 1 _,--_ _ _0 I" _ a Q 0 _ D 0 00 i..rl 0 ,'-4 I: : 0 0 ....

I I I ,-t _0 _0 c _3 0 " ' _ _" n:_ 0 I c _ ,- _ _ D _' l" P- ,i . _ I ::E: I I I I . , , 1 " r, _ 0 ¢ m 0 umO r _ II _ ,, o _ , -- i i . _ r _ . _ c m _ I u.- _ c ,_ _ ,,o i _ o I

d _

_ _ _0 _ ,- 1 _ _ I ::: I 0

'

_ J _ O O m _ 0 0 _ I _; OJ _ , -_ , , _ E ! co _ , H , ,-_ TABLE i0.- COMPARISON OF NASA AND RAE TEST CONFIGURATIONS i Lock no. K6 Damping, Sensitivity, Config. Mq, M6e, : Y IBm2 sec-1- rad / sec2 / cm,

I

RAE D3 4.10 0.30 -7.00 0.787 NASA AI07 3.00 .30 -7.00 .802 RAE A2 8.20 .05 -.60 .i_$ RAE AI 11.71 .05 -.40 .122 NASA 106 9.00 .i) -.881 .134 NASA i01 3.00 0 -.665 .036 • L REQUIREMENTS ACCORDING TO .... REF. 3 12 VFR / MIL-', ;-85 0 1A .... IFR f 10 BIC / S e = 0 . 59 deg / cm

/

/

/

8 /

I _ ; A'_._Y .,"

0 .2 .4 .6 .8 1.0 PITCH SENSITIVITY , MB , lc . rad / s ec2 / cm -1 Figur e i.- Effect of spring re s traint and L o ck number o n pitch damping and sensitivity.

REQUI REMENTS ACCORDING TO -- - REF. 3 V.F.R. I MIL-H-8501A •_ - I.F.R.

12- I I I I ,.J 0 .2 .4 .6 .8 1.0 P ITCH SE N SITIVITY , MB , lc , rad / sec2 / cm -1 Figur e 2. - Effect of hinge offset and Lock number on pitch damping and sensitivity.

A I 1 . 5

._ .s

°' 1 " 0c c _ -. 5 - 1.0 i I I i 0 .05 .1 0 . ! 5 .2 0 HI N GE OFFSET, Figure 3.- Effect of hinge offset and Lock number on pitch-roll coupling at hover.

I .6- -r 6.45 _" 9.675 I I I 0 .1 .2 .3 PITCH FLAP COUPLING PARAMETER , _ ( tan _3 t I I I I 0 5 10 15 20 P ERCE N T EQUIVA L ENT HINGE OFFSET (BASED ON F L A P PING FREQUENCY) Figure 4. - Effect o f pitch-flap coupling o n pitch damping (hover).

............. J -1.2 -1.6 I i = 0 .1 .2 .3

_tan_3

Figure 5.- Effec_ of pitch-flap co u pling on pit c h-roll c o upling at h o v e r.

FLYI N G QUALITIES CRITERIA BOUNDARIES .... REF . 3 NOE -1 0 -- MIL-F-833 00 (HOVER) - -- --- -- AGARD 577 (STOL) / DERIVATIVE VALUES COMPUTED AT 60 knots / / E)301 / f -8 -- / /i 207 0 / f u / A A4 / / 107Q / i r _ . 6_ // 0 204 f _ z / O 308 / / / Q / E)305 . . , // . ./ / . / O F1 = F108 , F205 (. 1 3: - 4-- / / U 201 , / / _ F2 F204, F301, F104 t -- / 302 03 39 /I AA2 // / 0 2 0 8 Z _ A1 A1 0 9 , A209, A303 E / " Z-' _ _ A2 A306 , A205 , A108 A3E)306 L u 108 -2 . / 303 Q 3 l n : F2 E ) 205 Z ^, . . A305, A308, A104 I E) _) u _ , 109 / _z uu A4 A107 , A301 , A204 ....

o, 'i,Z i ,- , , , C 101 .2 .4 .6 .1:1 1 , 0 1 . 2 103 PITCH SENSITIVITY, rad / sec2 / cm (a) Pitch damping and longitudinal control sensitivity.

Figure 6.- £ est configuration vaJues.

FLYING QUALITIES CRITERIA BOUNDARIES .... R E F. 3. NOE -50 ' MI L F-83300-HOVE R - - --- - ---- AGARD 577-HOVER DERIVATIVE VALUES COMPUTED AT 60 knots -40 O 301 -30 O 107 I A A4

I o o . "°' . - / .-

/ ol o4 , _ / I

- 20 J - . i O 305 201 , " * A A3 / t /

/ I / O 309 P _, / _ ., . v O 208

/I / 0_ 02 , = 2 . _ 10,, _ : _ --

i " I / Q 306,,A2 ,_ l'd _ __ _ r "10 Jl / 3 9 3 4 0510 _ ' _ , _ 205 0 209 I , I I I 0 101 1 2 3 4 102 ROLL SEN S ITIVITY, rad / sec2 / cm (b) Roll damping and l a t e ral contr o l sens i tivity .

Figure 6.- Concluded.

i F . , _ lZ ----_A .MCO P .

_ > _ 10 ........

N C'81 t _= l _ , == , = , o

8 6

q _

_: -2 L Ep:2 z 1 o i I I I ,I I- " 0 I

II '

°

v C-81 r,j - 2 - -- AR MCOP E

c.8 1

2 1 _ - _ ' - _ : : ' __ - - - = ' _ ' _

0 ' _ ' P "_ " _ I ' -I - 1 ARMCO P r_ - 2 I I I I I lJJ a . 0 20 40 60 80 100 AIRSPEED, knots (a) Trim c o ntrol po s itions.

Figur e 7 .- Comparison of ARM C OP and C - 81 mod e ls of UH-I H.

i

_4 _ c- 81

a I- _ _ _ _ ARMCOP I- 2- p. 1 0 I I I f J = 1 - "0 ARMCOP k -

F- 0 .. _ / __1------ 4 - -- I .... I - -- ----I

k.

- J C _ 1 -- , , , , _ :-1 0 20 40 6 0 80 100 AIRSPEED , knots (b) Trlm attitudes.

F i gure 7.- Concluded.

UH-IH at 60 knots (STABILIZING BAR OFF) O BELL C-81 DATA • ARMCOP E) i _

3 "

°== t E) - 2 - 1 0 o, l / sec Figur e 8. - A c omp a rison of e ig e nvalu e s of rigid body mo des of UH-IH o bt a ined from C-81 and ARMCOP simulations.

4 8 v, 0 BOEI N G-VERTOL

E 5 14 r A R MC OP MO D_ L

g ; 1o r .. -

.J 8 t I o _ , _ 10 _ " u . 8 - _ . _,_ E 6 " BOEING-V E RTOL .-- p . c a 2 - ARMCOP _ _.... "_ '- " * _ , . i. , I O I -- 0 . _. - -, , , , , " J _ -2 " _

_ o _ "

-J -2- BOE IN G-VERTOL 2 -- _ - 2 ' , i ,i , i , = - " _ 0 20 40 6 0 80 100 AIRSPEED , knots (a) Tri m c ont r ol p ositi o ns.

F i gur e 9. - Comp a riso n of ARM C OP a nd Boeing-Vertol mod e l s o f B O -I05.

2 3 _ BOEING VERT u L .

a _ 0 ' p .

= ARMC O P o - 2 " F - X" -3 -4 - ¢m 0 I I i I !

"o -1 BO E II _' G VE RTOL .J - '-4 I I i I I 0 2 0 4 0 60 8 0 10 0 AIRS P EED, knots (b) Trim attitudes.

Fig u re 9.- C o ncl u ded.

B0-105 AT 60 knots O BOEI N G / VERTOL DATA A R MC OP • 2 . _ E) " _ 1 3 " , ul C--" ' ' l , E) i _= I _= _ . _ -9 -8 -7 -6 - 5 -4 -3 - 2 - 1 O a , l / sec FiBur e i0.- A comparison o f eig e nvalue s of rigid body m o des o f BO - 105 o bt a ined by ARMCO ? a nd Vertol simulations.

Figure ii.- Ames S - -L 9 simulator.

52 O_1_,,,_,\E p_ , ,_: 1_ S 1I " I'_,)L, _,_ ' IY

1980022901-05

1980022901-

- 1000 -- E = _ CO M BI N ATIO N COURSE 15 m / __\_ _ 0 n

= I

O LATERAL-DIRECTIONAL CO

"' I ®

Z I / ( _L ONGITUDI N A L COURSE _ / -- _ _' --C: _ 0 - - "

_ - I1 / 0 0 0 II l) 0 0 0 0 0 n

q • • • • • • • • >.

l OOO I I 1 I I

0 1 0 0 0 2 0 00 3000 4000 X-DISTA N CE FROM REF. , m Figure L S.- Cour s e lay ou t, s cale 1:600.

Figure 16 . - Terrain mode[, 1:400 scale.

5 / F L YING QUALITIES CRITERIA BOUNDARIES -- " REF. 3 (NOE) -10 -- MI L -F-83300 (HOVER) -- - -- - - - - -- AGARD 577 (STOL) f / / 0 3.0 // / jr ' / -8 -- 3.5 0 / 3.5 // f_

I o / / i

z _ - _ 6 / // 0 3 . 0 // _ , I O ' .o_. o ,,, , ,- _ .

I -'_ : " / 6. o / / ....

- 2 / 0 3"5 g ' -EL "0 30 _ _7_._-_- 7,- ............

. u 6 . 0 J L

0 .2 .4 .6 . 3 1.0 1.2 PITCH SENSITIVITY . rad / sec2 / cm (a) Longitudinal-vertical ta s k --pil o t A.

Figure 18 .- Pilot rating s.

• -- T- 'ZI' _ I" II II ..... ,= , .........................

(b) Lateral-directional task - pilot A.

Figure 18.- Continued.

FLYI N G QUALITIES CRITERIA BOUNDARIES .... REF. 3 (NOE} MIL-F - 83300 (HOVER) - 10 - - -- - - - - '- -- AGARD 577 (STO L ) / / . Q 7.0 // // -8 I / / . I / E _ 6 . 5 /

I

-4 / /

_ " /I k .5 o ° ''° ,, i o ,,- - " _ o 5.5

[ = I 06 . 5 - E) 6 o -

/ [ L .._ 6. 0 __ 0 .. _ . / ........

0 7 . 5 . 2 . 4 . 6 .8 _ . 0 _ . 2

8.0 7 .5 P I TCH SENSITIVITY, rad / sec2 / cm (c ) LonRitudlnal - verticaltask --pil o t B.

Figur e 1 8 .- Continued .

1980022901 - 065

FLYING QUALITIES CRITERIA BOU N DARIES - 50 MI L-F-83300 - HOV ER AGARD 577 ( S TOL) II REF. 3 (NOE) -40 -I Q 7.0 I

° I

I 7.5 _ - 3O q Q 4 _ _.5 // "if' I _ / 0 55 4.0 i t

I / //

l ! D o0 o .. .

.0 _/ ___ 4.0

..11 / i"7. o _- o".' ..- ._ - _

-I U _-- I / O " " -- ¢ / . , , "_ e':

I, / e _o s_ o / e o /- , P _ 5. 0 , , . _

h I -6 . _ A ; - o _ _'°

0 7.5 1 2 3 4 7.0 7.0 ROLL SENSITIVITY . rad / s e c2 / cm (d) L ateral-directiol,al task--pilot B.

Figure 18 . - Concluded .

6 " 10 - O PILOTA 9 - D PILOTB 8- ® E] 7- ® E] I':1 El C) El z6 - r.-n EI] r:l 1 3

e e

m 5- _ C) El E] _-J4 - C) F I El Q Q® Q ® ® ®® 3 - Q OQ Q 2 - SATISFACTORYREGION 1 - FROMREF. 3 l I I ' , I 0 5 10 15 20 25 - Lp / L S a,cm / (rad / sec) Figure 19. - Eff e ct of ratio of roll d amping to roll sensitivity on pilot opinion rating.

10 F LONGITUDINAL TASK 9 t PILOT A 8 - m E] O 7- z I-" 6 - E] I' : I I-- 5 _ -J 0 F3 = '4- E]E} r:Ir_ E} E] E] r:1 3- f':'l Q f'_ f':'i r : t [] 2 - a ) 1 , __ I ! I I 10, P ILOT B 8 -0 OO Q 7 - O O O O O O O O z6 OO O O OO _- G O G O O c¢5 O O a.

b) 1 I I I I I 0 2 4 6 8 10 -M q, 1 / sec,60 knots Figure 20.- Pilot rating vs pitch damping.

• o,--------- ...........

10 LONGITU O INAL TA S Y 9 PILOT A El El ( ' 97 Z I- <6 0 B rr _ --5 0 E] ..J _-4 [] I ": 1 B E } E] E] E] [ 3 3 F I _ I ":1 1 ": 1 _ E]

B

a)

1 I I .L-----. _ i 9 PILOT B

8 "G

E) O

_ 7 E) E) (2)

- E: D E) E)

I-- 6 QQO QQ <

a : GE) Q Q

D -5 E) o Q

- - -4

a.

1 I I I 0 . 2 .4 .6 .8 1.0 MSe , rad / _ c2 /c m, 60 knots Figure 21 .- Pilot ratin B vs pitch sensitivity .

I 1 0 - L ATERAL DIRECTIONAL COURSE PI LOT A 8 - [] [] _ 7 -[] z [] _ 6 < =: [] [] I - - 0 5 El El EJ 0 " 4 _ _ [] Im []] ! _ [] 3 [] E] [] [] a) 1 i I I I I 10- 9 PILOT B (3 (3 ( 3 7 )(2) E) E)

z o

_ -6 E _ E _ E) ,¢ @ (3 Q t - o 5 Ex_@ - J @ @ 0 " 4 Q Q Q 3P - b) 1 I I I l I 0 10 2 0 30 4 0 50 - L p, ! / sec Figure 2Z. - Pi l ot rating v , _ rull damping.

LATERAL DIRECTIO N AL COURSE 1 0 - 9 PILOT A

8 3

E]

_7 F'I Z - 13 1- 4 6 13 1 3 _ 5 13 13 1 3 o .. J _ 4 1313 13 EEl 13 1313 13 3 El O E] El

a)

I0

F

9 _- P I LOT B Z7 E) O

p ®

6 - G (: -_ E) 0 0 a: 000 5 O0 0 o • J 0 E) E . 4 0 0 0 2 -

b)

I l I I I I 0 1 2 ;_ 4 5 L S a,rad / sec2 / cm Figure 23.- Pilot rating vs roll sensitivity.

P I L OT COMMENTS O LOW SENSITIVITY-SLUGGISH O TOO SE N SITIVE A ADEQUATE TO GOOD SHADED LOW _ AMPI N G REF. 3 PILOT A -10 .. o A

-8 // _ 0

- 4 0 A , / f _

_ - _2 1 - o " -' _

_- - / O & , _ El

Io ION&

0 t I . t ._L . I I I " a) P ILO T B -10[ ® - 8 - / 0

Z - i /

, 01 //

I- -2 I , _ I I I l I I 0 ,2 .4 .6 .8 1.0 1,2 P ITCH SENSITIVITY , rad / sec2 / cm b) Fi g ure 24 , - Relationship between pilot comments and helicopter pitch damping and sensiti v ity.

PILOT COM M E N TS O LOW SENSITIVITY-SLUGGISH -50 O TO O SE N SITIVE A ADEQUATE TO GOOD SHADED LO _ VDAMPING - - -- - -- - R E F . 3 _ 40 i 0 PI L OT A

- 3 0 - A // 1 1

/ /

-20 Q / 1 1

o A Q // _ , , I f

-lO O /

o _' 4ii "A

0 .... I . I I i

al

-50 _ - PILOT B

-4 0

ZO '-30 0 e , _ //

/

/ -20 O / I'

A // z . . 1

o , _O / / A

_._-- A

I t I . J 0 1 2 3 4 b) RO L L SENF _ . IIVITY , rad / sec2 i cm Figure 25 . - Relati o n s hip be t ween pil o t c omm e nts and helicopt e r roll damping and sensitivity.

_ " 4 LOC K HEED XH-51N r _ E _ / / ,/ BELLUH'IH

F -

r -- 0 t i J " 20 F XH - 51N ee 10

o , //- - -_ _ _ _ _

I J / I I I 0 1 2 3 TIME, _ c Figure 26 .- Comparison of hingeless rot o r and teetering rotor roll response in hover.

7O 10 - O PI L OT A 9 O PILOT B

D E] O

7 . _ O 0 z6 O E) O C O B E) Q

P 0 0 0

< _

= :5 " 0 O0

P E]

O

J4 r_E) 0 0

_ " B I::} _ o @

3 E _ r:l E]ra E ] _ 2- __ I , . I I I 0 .05 .10 .15 .20 M 8 c , rad / s e c2 / cm Fig u r e 27.- Pil o t r ating vs pitch c oupling du e to collective control.

Lp L _ a O- 5 0.57 I - I -10 1.14 O -20 2.29 10 - _ -30 3.44 9 - COMBINATION TASK ( 9 8 - PILOT A z _ 7- A

_5-

_ 5 - -

°° % o Q -4 -

3 - -_ A A

I":1

w _ .

li a) i I I I I __l 10, - -- 8 = PI L OT B Z

T -- . 7 . ®

.dr .1 _

A

I-

O 5- G

- " o

E 4 w - - 0 _ m .

b)

1 I J I I I I -,_ 0 4 8 12 16 2 0 24 L q , 1 / see Figure 28. - Pil o t rati n g vs roll coupling L .

q 7 2 COMBI N ATION TASK 'A ' CONFIGURATIO N S = • PILOT A O PILOT B 10 = .-- -- REF. 13 = _ 7 E) z - _ E) I I- <C 6 Q I Q (D (D I ¢ I E) t _ -5 I ( 9 E ) • I o • I I I • ' L _l _ g g , 3 i_ ll I- l I I 2 I i I I 1 I _ I = I I I I i -.8 -.6 -. 4 -.2 0 .2 .4 .6 .8 Lq / Lp Figure 29.- Pilot rating vs coupling parameter (Lq / Lp ).

P I L O T

AIB

9 " . _ O L O N G. COURSE _ 3 = 0

_ , 8 e% =

10 f I -_ J _1} COMB. COURSE tan _ 3=0.8

°

¢6 • • @

z a O • • •

°- s B •

z

E O Q Q

04 Q} e . ra

_ 0 Q O O

o; _ 0 00 O O Q

1 J I I I J 0 1 2 3 4 5 6 SHORT P ERIOD STABILITY PARAMETER (ZwM q - M e ) , sec -2 Figure 30. - Effect of short - period stability par a m e ter o n pilot opinion rating.

PILOT A, CONFIGURATION 105 LONGITUDINAL COURSE 80 / C.G. HEIGHT E u _ LOWEST OBJECT ¢J Z ON VISUAL ! HELICOPTER uJ 60 = c , SYSTEM (CRASH uJ PROTECT RI N G) w 4 0 > o T_.

_ 20 I.-- ,.1

.... i ,I

0 2000 3000 4000 5000 RANGE . m Figure 31.- Flightp a th history of helic o pt e r.

_ > ,: E) O PILOT B , . 40[ E) _ E) COURSE L ENGTH D 3660pI L OTmA O j Q Q ® ( 3

_, o oD_ o

I I i I , --I 2060 8C 100 i20 140 160 COURSE TIME, se c Figur e 32. - Mean altitude A C L vs tim e to compl e te cour s e - longitudinal - vertical task.

' 10 O PILOT A 9 O PI L OT B

8 Q

E] Q

_i Q Q Q

- 0 0

1- 4 6- 0 Q Q 0

_: O EX 3

i - -5 - O

O

13 (D

E4- l": 1 1 7

13 I.'l":'n

3 - 1 7

Q rn

2 -

1L I I I ' I 56 60 64 68 72 76 MEAN A L TITUDE ABOVE REF. , m Figure 33.- Pilot rating vs m e an a ltitude through longitudinal course.

r _ ° , a) CONFIGURATION 101 1 3 knots

- f = 3 c= 0 /

_ 100 / K / _ / J _ 2 = 0 DUTCH ' . 80 |

. o.. +6 0 1 2

1 0 0 _ L 40 i LONG. 80- _ I o 3 " SHORT 6 ;_ "1' "-' ROLL 80 40 PERIOD 40 _ 4 0 1 SUBSIDENCE _ 1 100_ PHUGOID - 3 - 2 - 1 0 = 3; _ = 0.10 K _ / I_ 2 = k not s DUTCH ROLL b ) CON F I G U RATI O N s , _ 0. 03 30 1 t 3

S U B S, D EN C E S U S, DE N C = _ "

• 8 : 0 0 _ 2 _ 406080100 80 6040 0 80 40' RO LL "PIT C H 11

, i l f J ' v A , [,I .! (, _. , , , , L_I_ _ PHUGO , D

-38 -37 -12 -11 -10 -9 -4 - 3 -2 -1 0 c) CO N FIG U RATION 201 -- 3 = 3 , _ = 0.05 knot s K _ / I , _ 2 = 0 100 p, / J DUTCH 80i -- 2 ROLL 60- < 40 - E ROL L " P IT CH -- 1 ._ SU B SIDENCE S UB SIDENCE"HEA V ING 6040 1000 80400 MODE 100 4|)_ J J Ill I j _ [ [ I (I[ I J J J[] / 0 " _ PHUG O ID -19 -18 -17 -7 -6 -5 -4 -3 - 2 60 -1 SPIRAL o , 1 / s e c Figur e 34. - Rigid body root loci.

CONFIGURATION - - --- - -- 201 ---- -- 301

, 1 . ° F

0 " ' I I . 2 . I r e __ --- ' 7 "_-- .... _ J o " 0 L i , _ -- , - .1 0 3 6 9 T , sec Figure 35 . Aircraft r esponse to a step _ input .

e _9 CONFIGURATION 201 4N - 3 01 E =" - 4 0 Z -80 L .

. 2 0 'l . . .. .

, - " o I

-.10 - .30 - f .10 _ ' _ I ' _ ' = " ==" _ ,, = I I _ I -.10 , i I 0 3 6 9 T suc Fi_:ur_ , 35.- ConL':lued, _ () CONFIGURATION • 201 . 0 5 - .10 - __ , _ . _.. , _" " "

o ' ' -- I I

- .O FF, I I - I _ E -10 -20 I t 0 3 6 9 T , sec Figure 35.- Concluded.

2 0 1 5 -1 . 0 0 5 T 10 EQUIV A LEN t t INGE o FFSET , % Figure 36 . - E_ect o_ r o t o r s y stem parameters on Lq / Lp ,

1980022901

-8 D3 _ N ASA P ,107 -6 C3 E ) z a.

< B3Q e_ - 4 I-" a.

A3 E) D2 -2- NASA E ) B2 Q C2 Q 106 B1 A 1 Q4 1'E)A2 E) Q C? D1 ' I i I I 0 .2 .4 . 6 .8 1.0 CONTROL P OWE r'l , r _ / sec2 / c m Figur e 37.- Configurstions t e sted in reference 15.

1 Repo rt No / 2 Governmeflt A c ¢es s ton No 3 Rec = p l ent ' s C a ta l og No NASA TM-811 9 0

I

4. T 6t ie a r i d Subt,tle o Report D a t e EFFECTS OF ROTOR PARAMETER VARIATIONS ON HANDLING QUALITIES OF UNAUGMENTED H ELICOPTERS IN SIMULATED s P e_ f o . m.ng O rea. ,zat ,onC _e TERRAIN FLIGHT 7 AuthOr( S ) 8 Perform , ri g O rgantzat , o n R e pc. ( No Pete r D . T a lbot, D aniel C. Dugan, Robert T. N. Chen, A-8158 and Ronald M. Gerdes 1 0 wo . k u . , t No 9 Peffotrn,ng Or gan i zat i o n Name a nd A ddre ss 505--42-- 21 NASA-Ames Research Center u Co n.a c t o r G :a , t No Moffett Field, Calif. 94035 13 T yDe of R epo rt a n d Per t od Co ve red 12 Spon f_ octng Ag e ncy Name and Addr ess Technical Memo r andum N a tional Aero z , a utics a_d Space Administration 1 4 Sponsor , n 9 Agency Cod e Washington, D.C. 20546 1 5 . Supplementary N ot e s 16 Ab s trac _ A coordinated analysis and ground simulator ex o eriment was performed to investigate the effects on single rotor helicopter handling qualities of systematic variations ±n the main rotor hinge rcstraint, hub hinge offset, pitch-flap coupling, and blade Lock number. Teetering rotor, articulated rotor and hingeless rotor helicopters were evaluated by research pilots in special low-level flying tasks involving obstacle avoidance at 60-100 knots airspeed. The results of the experiment are in the form of pilot ratings, pilot commentary and som e objective perfomnance measures. Criteria for damp- ing and sensitivity _re reexamined when combined with the additional factors of cross-coupling due to pitch and roll rates, pitch coupling with collective pitch and longitudinal static stability. Ratings obtained with and without motion are compared.

Acceptable flying qualities were obtained within each rotor type by suitable adjustment of the hub parameters; however, pure teetering rotors were found to lack control power for the tasks. A limit for the coupling parameter ILq / Lpl of 0.35 is suggested.

17 Ke y W ords (S ug ges t_ b y A ut h or( s )) I 18 Ot s t r_hJt, on S t a tement Helicopter R_tor systems Handling qualities Coupling Unlimited criteria

I

Piloted simul a tion STAR Category - 08 Terrain fJ_qt _ [ 1 9 _ c u r ,t v C l ass ' , _ ', s,,po r tl 120 _ ' b C u r ,t v Cl a _, f ( of t h,i _) 1 21 No o f Pa _s I 22 P , ( e " I I

[

Un c l_i"ied J Unclassified j 8_ $6.00 _ For sa l e b y t he Nat, T_ , N _. C _ I I n f or m . ho n _ e rw c e _ pr mq f _e l , J . Vi rq ln_ 22 1 G t