A Flight Investigation of the STOL Characteristics of an Augmented Jet Flap STOL Research Aircraft

r r , NAS _ TECHNICAL NASA TM X-62 , 334 MEMORAN DUM 1 - -?

SY M B OLS - A X a cce ler a ti o n a lo n g l o ngit u d inal ax is L A Z acce ler a tio n a lo ng ve rti ca l axis BLC bo un dary la ye r c o n t r o l < : ' . c mea n ae rody nam i c ch or d, 1 2. 1 ft (3 .6 8 m) [ . - CO d r a g coe'.f fi c i en t . q(widng:rea ) CDA dr a g c o ef fi c i en t exclud in g hot t h r us t c ontr ibu tio n " CDG dra g coe ffi c i en t in groun d effec t co ld t h rust " - < Cj jet m omentu m c oe f fi ci ent , isentropi c, " q (wi n g area) C _ / rolling momen t co eff ici e n t , rolling mom e n t q ( w in g a re a )(w in g span ) ,=. . C _ fl aC ___ :, ¢ aft lif t ' • CL l i ft coef fi c ient , _ - (wi n g a re a ) - CLA l i ft coef fi ci ent e xc l u ding h ot th ru st co nt ri b u tion ; . .

• :: CL G li f t co e ffic ient in gr ou nd ef fe c t . . _ : ..

W nz ?_ CLT t ota l l i f t c o e ffi c i en t , q (wi n g a r ea ) _ < . p i t ch ing moment : ,_ CM p it chi ng mo m ent coe ffi c ie n t , b - F ( wing a r ea ) ' ! " , CM G pit c hing mo m en t co effi c ient in gro u nd e ffe c t : thrust ;'_ C T thr us t (h ot ) c o e ffi ci ent, _ -( w i ng area ) ; , ; ?

,) FCOL c ol um n ( sti c k) fo r ce, lb • FW w h eel f or c e, i b ii " K degrees K e l vin .I " ME ma ss flow o f ai r i n e ng i n e 1 ;' M B m ass flow o f bypass ai r : _ A-54 ! 8 v :_ , P RI ,_CI,;DING P AGE BLANK NOT F I LMED _i ?

¢ Nt f high press u re e n gir e rotor s peed, rp m P n z vertical load factor " q - fre e - s t re am dyn am ic I. ." e ss u re , i / 2 p V 2 , Ib / ft 2 i P p e riod of osc il lati¢,- l, sec '" P I engine inlet p re ssu re , psi _ : PT engin e byp a s s air total pr e ssu re , p si • • yawing rate, rad / s e c rpm r e volution p e r m inut e SAS Stability Augm e ntation Syst e m t ti m e, se c T _ 12 t ime to I / 2 amplitude , s e c 7 " 2 time to double amplitude , s e c g T _ engine inlet temp era tu re, ° K • V velocity, ft / s e c or knot s I VE e quival en t air sp ee d (EAS), knots _ VS s ink rat e , f t / s e c VFR visadl fli gh t rul e s :. VSS Variabl e Stab il ity Sy s t e m W w e i gh t, lb _ ' F angl e of attack of fus e l a ge, positiv e nos e up, d e g i - : I S a ngle of sid e slip, positive no se l e ft, d e g / " _ / fli gh t pat h an gl e, posi t iv e up, d ei ; [

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6 A 1L a i l e ron d e fl e ction, positive T.E. down, d e g 6 CH choke d e flection, po s itiv e T.E. up, d e g t l ,

; v i - s4 1 s ! -

: i . 6 CO L control column deflec tio n in p it ch, p o sitive a i r , de g _: 6 c e l e v a t o r d e fl ect io n, po si tiv e T .E. down , d eg 5 / fla p d eflection, p os it ive T.E. do wn , d e g - - 6 p r udd e r p e d al d efl e ction, po sit i ve l ef t for wa rd , i n.

6 sp sp oil e r deflec t i o n, p o s i ti v e T .E. up, d e g ? 6 w co n t r ol whee l defl ect i on, po s i t i ve clo c kw ise, d e g 0 pitch angle, positive no s e u p , deg , : _ b raki ng c o e ff icien t _i v n ozzl e d e fl ec tio n, po s iti ve d o w n f rom f u l l a ft . re l a t ive to fuse l a g e d atum line. de g , 2 .' _ d am ping ratio _" p ambi e nt a ir de n sity , s l . ugs / ft 3 _ r A a pp a re nt roll m od e t i me c o n st an t ¢ , r oll an gl e, posi t iv e rig ht win g do w n, de g ; _ _ ro ll an g l e after ! sec, d eg ¢ , _,i g , y a w a n gle , posit i v e nose ri gh t , d e g _. wn na tur al fre q ue n cy, r ad / se c

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l ' A-541 8 vii 1 _ A F LI G HT INVES ] ' IGATt O N O F T HE STO L C H ARACT ER I S TI C S O F " AN A UGMENTED JET F LAP STOL RESEARCH AIRCR AF T Her v ey C. Q u ig l e y an d R o b e rt C. I nni s A mes R esearch C en t er _ and Seth Gr o ssmith . : M inist ry . o f Transport, Cana d a ?

; INTRO D UCTION Several p ow er ed lift concepts are bein g stu di ed by NASA for' p o_ible fu t ure u se on fan j e t STOL transport airplanes. T he augmented jet flap or augmentor.wing concept has been recognized by both government ( ref. 1) a nd industry (refs. 2-4) as one of the promising concepts for further research and development.

, A c o o p erative N A S A / Cana d ian Government research program on the a ug mented je t fl a p c o n- cept began in 1965. The program included analysis and small-scale static and wind-tunnel tests (ref. 5): large-scale tests in the Ames 40- by 80-Ft Wind Tunnel (refs. 6-8) conducted by NASA in " cooperation with th e Canadian Defense Research Board u sing a de Havilland built model: and NASA design feasibility and simulator studies. Research progress by early 1970 warranted develop- ment of a proof-of-con c ept aircraft to test the jet STOL principle in flight. The U.S. and Canadian governments entered into an international agreement whereby the NASA and the Canadian De p art- ment of Industry Ty.ade and Commerce ( DIT C ) would modify a de Havilland C-8A Buffalo to an attgmented jet fla p STOL research aircraft. The D ITC contracted with the de Havilland Aircraft of Canada, Ltd., and their subcontractor Rolls Royce of Canada, Ltd. , to provide and modify the jet engines and modify the nacelles. The NASA contracted with T he Boeing Company to modify the aircraft, pro v ide the augmented jet flap system, install the pro p ulsion system, and perform the _ in i tial fl ight tes t s . Ref e renc e 9 summarizes th e contract o r d e velo p ment pr o gram and de scribes the augmented jet fl ap STOL research aircraft.

F "_ The C -8A Buffalo aircraft was chosen on the basis of a design feasibility study, which showed that with re q u i re d aircraft modificat i ons, the p rimary research obje c ti v e c o uld b e achieved at a reas o n a ble cost an d within a n ac c e pt ab l e t ime sp an. I n a d d i t io n, c ons i d erable design da t a were _ a va ilable from ex t ensive t est i ng in t he A me s 40 - b y 80 - Ft Wind T unnel of a large - scale model ha v ing _. a win g p lanf o rm similar to the C - 8A (refs . 6 - 8). (Simulation in p uts to th e d evelo p ment of the ; ai r cr a ft are d iscusse d in refs. i 0 -12 .)

• The fir s t flight of the a i rcraf t wa s ma de on Ma y 1, 1972, at Seattle, Washington. T he i nitial i , airworth i ness flig ht test pro g ram was con d u c ted by T he Boein g Compan y (ref. 13 ) . D uring these t_ . , , ,tests th e aircraft was flo w n withi n a fl ight envelope of from 5 0 to 180 knot s an d at loa d factors _ ; .

_ 2 . s u ffi c i en t to d emonstra te t hat the ai r craf t fl i g h t load s were with i n des ig n an d the airplane flu t t er _ fi : ee. The aircraft was de l i ver ed to NASA on Jul y 3 1 , 19' 7 2.

' .',_- A - S4 18 Th e fl ig h t te st p ro g r am o bj ec t ive s a re ( ! ) t o det e r m i n e t h e i n - fl ig h t a e r ody na mi c , p e rf or manc e , an d han d lin g qual itie s o f a jet S T OL aircraft inc o rp o ratin g the a u gmente d jet fla p co nce p t; (2) t o - ; compare the results obtained in flight with characteristics predicted from wind tunnel and s imulator test results; (3) to contribute to the de v elopment of criteria for design a nd operation of jet STOL transp o rt air c raft; and ( 4 ) t o pr ov i de a jet STO L transp or t ai rc ra f t f o r STO L systems research and ° de v elopment.

This report presents results obtained during the first 8 months of proof-of-concept fliglat testing of the aircraft in STOL con fi gurations. Included are a brief description of the aircraft, fan-jet en g ines , and systems: a discussion of the aerodynamic, stability and control, and STOL perfor- mance: and pilot opinion of.the handlin g qualities and opera t ional charactei-istics.The tests did ng.t include flight near maximum lift c oefficient beca.use of the limitations of the longitudinal control system, night t ests at high angl es of att ac k at or n e ar CL max will be con a ucted following modi- '- fi c ations t o the aircra f t t o inc o rp o rate a powe re d lo ngitudinal c o n t r o l system.

T he fligh f t es ts w er e c o nduct ed by a proj ec t te am c o nsistin g o f th e fo ll ow ing pe rsonn el ; all are at N A S A - A mes Research C enter unless otherwise sp o cifi e d: Project M anag e ment " Da v id D. Few : Her ve y C. Qui g l ey .

P r o je c t ' P i lot s - - Robert C. lnnis _ Se th Gr o ssmith, Ministry o f T ransport , Canada ° , : P r oje ct En gi nee rs Jerry P. B arrack .." _ " _ .

- Al fred G. Boisse v ain Br _ c e L !! le y, The d e Ha v illa nd Air craft of C ana d a , L t d. _" Ja c k W. R atcl i ff B ria n Sw a n , Dep _.rtme nt of Na t i on al De f e n ce , Ca n ada _ i R i c har d F. V om aske " Jo h n W. We ye rs °, ; Nt : 2 A - S 4 1 8 : -5 ¢ ,, T HE RESEARCH AIRCRA F T _ .

B _ L : T h e rese ar c h air cr aft i s a hig h l y modi fied de Ha v illa nd C- 8 A Bu ffal o military tu r b oprop t r ans- : _- p o rt. I t s high w ing a nd h igh " T " t ail made it especially s u itable for application of a powered lift " system. The landing and takeoff configurations of the aircraft are shown in figure !. Table 1 lists , the geometric and mass cl_aracteristic s of the aircraft. Figure 2 is a three- v iew drawing of the : a ir c r a f t . Sp e c ial f ea t ures of t he air c raft a re describiM b ri efly below; a more complete description is : pr es e nted in r e f e r e n ce 9.

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•! , Engine _ • ,. T w o R ol l s R o y c e S I6_y MK 801- SF sp li t fl ow eng i nes, o ne m o unted in each o f two nacelles i : - ' (fi g . 2), pr o v i de t h e t h r us t for t he aircraf t as well as t he air for the augmenta t ion sys t em . The engine .

i s m oun t ed t o t h e s ame n ac elle s t ruct u re as u sed f o r the T -64 in st alla t io n i n t he original C-g A. The _.

MK 8 01-SF engine ( fi g. 3 ) is a hybrid engine of 0. 6 bypass ra t io, which was as s embled from several _ e n g i ne s by R o ll s R o y c e. I t co nsi s t s o f a Spe y MK 51 !-8 core, a Spey MK 512 low p r essure corn- : '_ pr e s s e r , a S pey MK 555-15 high p r e ssure c o m p ress o r ex te rnal gearbox, Av o n MK 1 0 1 l o w p r essu r e " dum p valves, a new bypass air duct, a Pegasus MK 5 trouser piece, and new v e c torable conical _ n o zzl e s.-Ex ist i n g engi h e h a rdware w as use d t o the e x te nt po ssibl e to redu ce de v el op ment t ime, ri s k, ' a nd"c o s t s . -Th e resul t ing engine re p res e nts a ve ry useful resear c h t oo l, bu t is n ot n e c e ssaril y an ._.

_- o p t imum c o n fi gurati o n f o r future augment er win g aircraft. A m o re detailed descripti o n o f the _: e ngine and a di s cus s i o n o f i t s devel o pmen t is provided in reference 14.

_ The low p r essure compressor f or t he engine has five s t ages, t he firs t and fif t h being made of ti t anium wi t h mi d -s pa n s nubbers to mak e t h e c o m pr ess o r mor e to le r an t of inl et fl ow d is t or t ions _ and deviations from the nomina l compressor w o rking line. Its maximum pressure rat i o is 2.5. _ , _.

The e n gi n e b yp a s s a ir is colle ct ed a nd di . s ch arged t hrough t w o 13-i n .-diame t e r (0.3 m) duc t s -_ l o cated at th e top o f the e n gine (fi g . 4); it is distributed to t he augmenter fu s elage and aileron .

_!i n o zzles b y t h e a i r d i s tribu t ion s y s t em d iscussed below .

t rou s er p iec e an d out .through two coni c al nozzles . The t rouser piece was originally designed for a _ *_ ii Th e fl o w f ro m t he h o t s ec t i o n o f t h e en gi n e is di s charged i nto a s l i gh t l y str eng t he n e d Peg as u s ,; mu c h l ar g er e n g ine, a nd a "colander p l ate " ( fi g. 3 ) betw ee n t he t u r bine a n d t he tr o u se r p iece allows _ , tro u s er pie ce or n ozzle a f fec t ing'e ng ine opera t ion. The colander pla t e _s a l-in. s t eel pla te wi t h ; ' ; 40 0 o n e-i nch-di a m eter (2 . 54 c m ) h o l es, 3 6 of w h i ch are p lugged to acc o mm o d at e t h e en gine d i s - _ } I p r op er m at ching o f t he c o m p ressor s an d t u r bines wi t hou t t he fear o f p o ssib l e d ist urbance s i n t he c ha rge flow ar e a re qui re men t . T h e c oni ca l nozzl es c an b e vectored b y m ea ns of a H a wker S id dele y i _ no zzle c o n tro l s y ste m fro m 0 0 to 9 8 * d o wn relative t o t he engine centerline (6* to 104" relative to ), "_ , t h e f u s elage water li ne) t o pr o v id e fli g h t path c on trol and inc re a se ex pe riment a l ve r s a t ilit y. T h e !

, n o zzle vec to r an gle co ntr o l han d le s , o ne fo r each e n gine, are l o cated in the c o ckpit o v erhead _ -.

consol e a dj a cent to t he t hro tt le s w i t hin e a sy rea ch o f the ,, p il ot. "!i * _ : " T he S p e y M K 80 1- SF e n g in e , operat i ng p ar am e t er s ar e s hown i n f ig u re 5. The b y p a ss (c o l d ,_ _ t h r u st shown is is en tro pic co ld t h ru s t a t t he eng i ne offta ke. _ ._ A -5 4 ! 8 3 I A i r Dis tr ib uti o n Sys tem q The a i r d is t r ib ution system di r ec ts the en g ine b y pa ss a i r to t ile u p per an d lower augmenter noz zles, to the f u s e l a g e b ou nd a r y l aye r b l ow in g noz z les, and to the ail er on bl owin g i_ o zzl e s (fi g s.,6 • i and 7 ) . A c rossover ducting s y s te m i s u s ed'so t ha t approxima t el y 64 p e rcent of the b y pass mass flo w o f e ac h e n gine is du et e d t o t h e augme n t er an d a il e ro n n o zzles on t h e o pposi t e wi n g and to half of the fuselage bou nd ar y- la y er b lo w ing nozzles , , while the remaining 36 percent of the bypass mas s flo w is ducted aft to the augmenter nozzles on the same side of the aircraft as the engine. The air distribution sys te ms for each &lgine are completely separa t e, but identical. This uniqu e arrangement . , _ p ro v i de s f o r en gi n e - out o pe rati o n w i t h ou t la r ge rolli n g o r y a w in g moments. , ; The m ass flow f ro m t he i nb o a rd off t a ke port (3 6 pe r cent of t h e t o t al m ass fl ow ) - of e ac h e ngi ne i s duc t ed af t i n the n acelle t o a tee in t he l ower (i nn er) a ugmen t er n ozzle as s e m bly. Thi s • d u c t f r o m th e engin e af t t o t he l ow e r augmen ter _nozz l e du c t co nt ai ns a cal i b rate d fl o w m e asu r i n g - .

k s t at io n. ' , ._. Th e m ass fl o w from the ou tb oar d off tak e d uct ( 64 per ce nt o f the total m ass flo w) o f ea c h e ngine i_ dir ec t e d through a 1 4 - in.- d iam e t er ( 0 .35 m) du c t alon g th e f r o n t spar o f t h e w in g a nd - , across th e int e rior of the fuselage to the upp e r (outer ) augmenter nozzle duct at th e r e ar spar of th e o pp osit e wi n g . A ca l i b ra te d mass fl ow m easuring station is lo cated in th e stra i ght s e ction of th e du c t along t he wing front spar. A 6 - in.-diamet e r (0.15 m) duc t is tapp e d into the fus e lag e c r osso,6 i duct ; to p rov i de ai r f o r th e fuselag e b o un d aryq ayer blo w i n g nozzle s . O f th e 6 4 perc e nt of the engin e mas s flow carried b y the c rossover ducting system, approximately 7.1 percent is used for fuselage blow- ing, 44 per c ent by the upper augmenter nozzles, and the remaining 12.9 perc e nt by th e ail e ron b o undary - laye r c o ntr ol n o zzles ( fi g" 7 ).

Th e design duct M ath number of the air distribution system was 0.3 to prevent excessive fl ow losses. , The resulting bypass air fl ow los s es are summarized in fig ure 8. For the purposes of this rep 6 rt , all bypass ( hrust levels are d efi ned as the isentropic thrust that would be obtained for fi:lly expanded fl ow based on the pressure r temperature, and mass flow measurements just downstream of the engine offtakes.

At takeoff, th e ma s s fl ow throu g h the air distribution system is 79 lb / sec per engine at 270* F • a n d a pre ss ur e r a t io o f 2 . 5. Th e i ns t a ll a tion thru s t l o sses d ue t o p ressur e dr op s i n the su ppl y d uc ts to the flaps are about 4 percent of the fan thrust. A n additianal th ru st loss of about 7 percent occurs through the hi g h aspect ratio nozzles and the flap ducts. Addition_d.information on duct losses are provided in reference 9.

:: _ A u gme nt e d J e t F l ap : . F i g u re 9 is a s k e t c h o f the gener al a r r a nge m e nt of the aug m en t e d je t fla p an d i t s raaj or ', co m pone nts . The fla p geome tr y i s bas i ca l ly t he sa m e as t ha t of th e l arge-sc al e m o d e l t es t e d in th e .

; , A m es 40- by 80 -Ft Wi n d Tun nel (re f. 7 ) . The fl ap s ha ve c onsta n t c h ord a n d are made in fo ur equ al s pa nw ise seg m en ts , tw o on e ach si de o f th e a ir c r a ft. I' o red uc e th e o v e rall co st of th e modifi c a t i on .- !_ .- program, t he flap as s em b ly wa s , _ o t d e s i gned to r et r a c t i n to th e s trea mli ned airf o i l c o nt ou r as w oul d , . . .b e req ui red for h igh s peed flig ht.

' t 4 A- 54 18 r' i " _ The entire augmenter flap assemb l y, inc l uding the augmenter nozzle duct, is supporte d on % • " beafns external to the wing that attach to the front and rear spars. This arrangement permits "_:_ accurate alignment of the flap and nozzle assembly and minimizes the effects of wing deflection.

= _- Th e fl ap s are de fl ec t ed by hydraulic lin e a r ac t ua to rs m o un te d e:._ernal to t h e wing. The _ miuimum(flaps up) angle is 5.6 ° and the maximum fl ap angle is 72°.

._, _ ; _ The ducts t h at supply the ai_ t o the flaps are m o un t ed just aft of the rear spars and provide air r ¢_ independently from each engine. The inner duct air is . _upplied from the engine o n the same side of the aircraft as the engine, while the Outer duct is supplied from the engine on the opposite side. • , ' . _'_ The fl ap it se l f is m a de u p of t wc surfaces, eac.h of which has s lots for a dded air fl ow and - - _. b o u n d a ry-la y er c o nf r ol. As th e flap defl e c ts, th e h_take d oor on the top s t,,rface al s o deflects >-": - (fi g . 1 0 ) to a l l o w .a s moother out si de a i r entry i n t o the flap sy stem an d to re_ , a rd flo w s e p aration ) _ : . from the upper surface of the shroud. Note also that the flap system pivots about a poirit within the _... Coanda leading edge portion of the lower surface so that the ejected air fro m the nozzles is tangent to t he sur f ace o f t he Coa nda as t he flrp de fl ec t s . S t atic t e sts of a 0.7-scale m o del ( ref. 15 ) sh o wed _ t hat t he loca t ion of t his t angency poin t rela t ive t o t he eflux cen t erline was of cri t ical impor t ance i_ f or o bta i n i n g t he maxim u m au g menta t i on r a ti o. Also, the opti mum Coan d a s ur face p o siti o n f o r _. , e a ch fla p de fl ec t ion is at a diffe r ent l o cati o n. Th e loca t ion chosen for the aircraft represents a ._ com p romise that fa v ors a flap deflection of 45 °. The sensitivity of performance to physical posi- ._'_ tioning of the f lap components resulted in a design of relatively rigid structure to minimize , de f or mati o n _ " Fixed leading edge slats were installed to help maintain airflow over the wingat the high values " _ of c i rcula t io n ob t aine d du ring p owere dq ift o peration.

_. _ . . Flight Control System _ _ Several modifi cati ons _ c; re m ac l e to t he b a sic C -8A Buf f al o fl igh t c o n t rol system. In the _ c oc k pi t ( fi g . 1I ), t he C on tr o l wheel was _e p lace d w it h a w heel i n s tru m ente d t o rea d o ut la t eral and longitudinal c ontrol force s • The only other change in the cockpit fl ight controls was the addition of an electrical lateral and pitch trim switch , which the pilot could c ontrol with his left thumb and the co-pilot with his right thumb; the manual trim was retained.

The lo n gi t u d i nal co t itro l s y ste m is the b asic C -g A Bu ffal o ma nua l sp ri n g tab s_ , s t em modified to reduce th e s tick forc es for "on e hand" pilot operation. The . s tatic stick force variation with " c o n t ro l co lumn positio n is s h o w n in fig ur e 1 2 . T he dyna mic f o r ce ch ar a c ter is t i c o f a s p ri ng t ab 1 system requir es an i n itial "forc e to d efle ct the tab that i s momentarily hi g her than the static for ce characteristic s . As di s cus s ed later, th e se forc e characteri s tics alon g with mass balance forces at low sp e ed s were found to b e undesi r a ble.

,The dire c tional control sy s tem co n s ist s of a two-panel rudder, th e aft panel being hinged to the trailing edge o f the forwa r d panel and geared t 9 it with a 2:1 de fl ection ratio. The rudder is full y • ' pow ere d t h ro ugh an i rre v ers i bl e d ual h ydra ul ic actuat or cont r o lle d by cabl e s f r o m t h e p edal s . T he o r:ly modi fic ati o n to the. b asic Buf f al o directional control system was the addition of a Stability 4 ' • A -5418 5 _ ;

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A ugment a tio n S y s te m ( SAS ) a c tuator m e ch anic al ly s u mm e d i n se ri e s w it h pilot i npu ts. The ru dder pedal I o r c e and gearing are shown in fi gurt. 13.

The lateral control system is completely new. Three separate surfaces are used to produce the requir e d rolling moments: ailerons with boundary-layer control, spoilers in . front of the ailerons, and an augmentor choke. Figure ! 4 shows the position and function of each element. The ailerons : are m echanically programmed to droop as a function of the flap deflection as shown in figure 15(a).

Full d r oop i s 35° a nd is re a che d a t fl ap de fle c ti o n of a bo ut 70 ° . The dif f erential a iler o n d eflec t i o n • : from the droop position is -+17 ° . Blowing boundary-layer control is used on the aileron to increase i t he effe c ti v ene ss elf bo th t h e ailer o n a n d s p o il e r. T he augm e nt o r c ho ke s are d e signed t o co nt ro l the _. lif t o f t h e. augm e nt ed jet fl a p sy s tem b y c h a nging the e xit of th e augmentor. Full lat e ral control . r e duc e s th e e xi t a re a as a functi o n o f flap de fl ec ti o n t o a maximu m o f 55 percent on the downgoing wing ( fig. 15(b )) . The chokes function much the same as spoilers on a conventional airplane.

A lthough there ar e augmentor chokes in ea c hOf the four sections of the flap, only the chokes in the outboard section of each wing are used for l ateral control. All four chokes are activated on the grotmd after landin g for lift dump.

The thre e lat e ral control surfaces are programmed to give nearly linear effecti ') eness with control wheel d e fl e ction (fig. 16(a)). The aileron and spoiler op e rate from 0 ° wh ee l deflection, and the augmentor chok e is phased in at 17 ° control wheel deflection. The s poilers are fully deflected at 48 ° wheel defl e ction.

- Lateral c o nt ro l wheel forces ar e l o w ( fig ! 6(b)) and a r e p r oduced by a s im p le s p ring sys t em.

The lateral control surfaces are activated by a central dual hydraulic power actuator, located on the rear s p ar, which , drives the ailerons through a cable system. The central lateral p ow e r actuator also d t 4ves spo il e r 2 a nd augm ento r ch o k e con tr o l va l ves thr o ugh a s ec o n d c a ble. Th e s po iler _ r d ch o ke actuators are on differ e nt hydraulic sy stems.

S tability Au g mehtation System _ Two i ndepe :l de nt s tabil i ty a u g m en tati o n sys t e m s , o ne fo r th e lat e ral a nd one f o r th e dir ec - ti o nal air c raft axis ar e pr o vid e d. Each s y st e m off e rs the pilot two m o d e s of SAS op e ration: n o rmal mod e ha v in g fix e d g ains f o r n e ar op timum c haract e ri s tics and a variable stabilit y m o de with g ain adjustm e nt. Th e normal mode is us e d for basic STOL o p e rations, atld the v ariable s tabilit y mode is u s ed for chan ge s in gain s etting in handlin g -qualities te s tin g .

, The S AS actuators a r e pos itioned b y cl o sedq o op ser v os a n d s u m m ed in series with t h e pil o t's -_' c on tr o l sys t e m . Th e S AS electr on ic s is s i n gle cha n nel, a nd r elie s o n lim i t e d r a t e an d d i s p l a cement ; ' ' a u t hi_rity f o x' s af e ty , . Th e w he e l l a te ral r a te limit is 5 0* / see while the d i sp lac e m e n t a u t hority i s _ l i m i t ed t o : 1 : 2 0 " ( 27 percen t ) e qu iv al en t w h ee l di sp la ce m en t. T h e maximum di rec ti onal r at e i s 25° / s e e of ru dd er t r a ve l, a nd d is p l ace m en t is l imi t e d t o ±50 ( 2 0 perce nt) of f ore rudd er d e fl ec ti o n.

Th e la ter al S AS i n t he no rmal m od e pe rf o rm s the f o ll ow in g f unc ti ons: " I . Sp iral stab il i t y aug m en t a ti on, u s in g yaw f_ edb a c k to t h e lat er al co nt ro l s . ,, 2. R o ll dampin g augm e ntati o n, u s in_ r o l l rat e fe e dback t o th e later a l co ntr o ls. " 6 A-541 8 S l r , . • .

3 . Lateral control quickening, us i ng wheel position fee d forwar d (-+3 ° max i mum) to the later a l : controls.

• The directional SA S in the normal mode perform ' the following functions: ' 1. Turn coordination, using roll rate and roll attitude feedback to the rudder.

2. D utch roll damping, using yaw rate and roll attitude feedback to the rudder.

N o rrr._l m od e gains (fig. 1 7) are p rogrammed with flap positio n a nd are automat i call y sw i tche d , _ off above 100 knots. Figure 17(c) is the schematic for the lateral control aileron quickener, which doubles t he la t er a l c o ntrol g e a t ; ; '_g f o r t he fir s t t hree deg r ees o f c o ntrol wheel t ra v el to improve t he -" _! control ch a racteristics near zero deflection.

_ " _ ,.

" _ : _i Data Acquisition System , _, . A n on b o a rd data acquisiti o n system g athers data o n abo u t 95 pa r ameters measu r ed during all _ ground and fl igh t te s tin g . A pul s e co d e modulated ( P C M ) digit a l system r e cor ds t he d ata o n : magne t ic t a pe. Reco r ded fl ight t es t data includ e stabili ty and contr o l, S AS signals, e ngine perfor- _._ mance, augmenter performance, and guidance information. Each channel is sampled 100 times per :: , se c o nd. The P C M wo rd l e ng t h is 1 0b it s. The s ys tem ' s 1 4 -tr a ck t ap e recor d e r has a ca p acit y o f _ about 5 0 min rec or di_ g t ime per reel of tap& To allow da t a g a the r ing dur i ng fl ights that exceed _ _; _ 50 m i n du r a tio n, t h e p _ : ot ma y star t a nd st op t he recor d er a t will. _ c alcula tio ns f or de t e rminin g such p arame te rs as en gine l h ru st , lif t an d dr a g coefficients, fl i g h t path I T he I_ M data. is processed on a Sigm a -7 di g ita.l comp u ter programmed w i th t he desired ; " a n gl e , and c o rr ec t e d airsp e ed . _ " ' d ' _? The o nb o ar d dat a acquisi t i on sy st e m a l s o . inclu d es a reco r ding o sc i l!ggr a ph, w hich is use d f or i i _" a dditi o nal r eco r d ing requir e men t s when the available number of PCM channels are exceeded. The " _ o scillogra p h al s o h as a mu c h hig he r ( t o 5000 H z) frequency r esp onse used foi: recording such parame t ers a s l oads an d vib r a t i o n. . , F LI G HT T E S T P R OCEDU R E _!

Th e flig h t test s w e re in itia ted f r o m M o ff ett F i el d an d fl o wn i n tes t ar eas i n t h e vic ini t y o f _ _ M o f f ett F t e ld and at t he Crow s Lan d in g Nav al A u x il i ary Landing F a c i l i t y . The flig hts w er e ma de by _ pro j ec t p i l ot s fr om NA S A , the Can adia n Minis t ry o f T r ansport , The de Ha v ill a n d Aircraf t of _t C a nada . a nd The Boe in g C om pa n y. J i

•

•Th e fl i ght tests in cl uded t he following g e ne r a l c a tegor i es : ca l i brati o n , aerodyn am i cs, stab ilit y _ _' .

_!_ a n d c o nt ro l , perfo rm a nc e, and o peratio nal an d h a ndlin g qua l i t i es.

, p A - 5 41 8 , 7

t

q .. Ca l ibra t i on T es t s : _ The ai rs pe e d and st a t ic p r essu r e p osi t ion e rr o rs we r e de t ermined b y a "pace r " helicop ter t owing a calib r a t ed " tr ailing airsp e ed bomb." F lap angl e s fo r t hese calib r a t ion fligh t s w er e 5.6 ° , 30 ° , 5 0°, a nd 6 5° . in ge n e ra l , t he engine po w e r was that r e q u i r e d fo r l eve l fhgh t w it h n o zzl e s a t " _. minimum deflec t i o n (~ 60). F o r s o me o f t he 30 ° flap tes t s, t he powe r was he l d fixed a t 96 pe r cen t : rpm and no z zle vectored be t ween 50* and 80 ° t o main t ain level fligh t over t he t est speed range.

_' T e sts w e r e als o c o nduc te d wi t h 65 ° flaps, 94 p er cen t rpm , and v ari e d nozzle deflec t ion over the : " s peed range of 60 to 75 kno ts. Th e t e st a l t i t ude w as bet w een 3000 and 5000 ft (900 and 1500 m), _ _ and gro s s w eigh t v aried be tw een 3 7 ,000 and 4 4 , 000 l b (16, 8 00 and 20,000 k g ). _ o T h e angle-of - at t ack p o si ti on e rr o r calibra t i o ns were made in c o n ju nc t i o n wi t h the aer o dynam- :i' : ics tests d e sc r ibed l a te r. I n g en er al, t h e an g l e of a t tack was det e rmin e d by t he diffe r ence be t wee _ . irplane att i t u d e a n d fli gh t p a t h. T h e air p la n e a tt i t u d e was d e t ermi ne d fr o m the p itch a t ti t u d e g yro and the l o n g i t ud i nal aceel er bmeter. Fligh t p ath wa s de t ermined from ra t e o f change of a l t i t ude and ,._' air p la n e ai r s pe ed.

The p o s i t i o n error da t a derived in t he s e te st s.for ai rs peed, al t i t ude, and angle of a t tack is pres en t e d i n t he ap p e n dix. : C ' Aero dynam i c Te sts . , The airpl a ne aer o dy na m i cs tests we re c o nduc t e d b y fly ing in u n accele rat ed fli ght f or ab o u t '_ : 10 sec, h old in g a ngl e of a tt ack s t ead y at the v ari o u s t e st condi ti ons. Ex te nsive t e s t s w ere p erforme d - :: in wi ng_ -l e v el fl ig h t , w hi le effo r ts at tu r ning fli g h t ("wind - up t urns") t es t ing we r e sharply limi t ed by p il o t d if fi c u lt y i n e sta bli s hin g s t e a d y con di t io m . Reco rd ed t e s t da ta were u sed in determirdng lift, dra g , thru s t, and c ont r o l charac t eri st ics. A e ro d y n a mics t ests !ncluded t he following configu_'a t ions : ' : a n d c onditions : gros s wei g h t s, 3 7,C30 to 4 5,000 Ib (16, 80 0 to 2 0 ,40 0 kg); fl a ps , 5 . 6 °, 3 0% 4 0 °, ' 50 ° , 6 5°, 7 2 ° ; al t i t ude, s ea l e' , . o 10 , 000 f t ( 3000 m); engine power, 89 t o 100 pe r cen t rpm i nc l uding si ng l e eng i n e; eng i n e '_° to 10 4 "; an d angl e o f atta ck , .-to to 1 8 " .

, !

,_ St ability and Contr ol Tests Sta bili ty a nd contro l t es ti n g m a y b e sep arat e d i nto t w o cat e go ries, lat era l-di r e ct i o nal and Ion gl t ud in ,d . S o me st abili ty a nd c on tr o l test data w er e gath e red durin g a e ro d y na m ics a nd calib ra - tio n test i ng . Test c on fi gu ra t i o n s were g ene ra l l y cruise (5 . 6 ° fla ps ) at 15 0 k no t s , l a nd in g a pproa ch (65* fla ps ) at abo u t 65 k nots, and ta keo ff ( 3 0 ° ". ap s) a t a bou t 80 knots. P o we r for l e vel fligh t w a s :_ u se d f o r the c ru i s e and landi ng approa ch test in g, whil e m axi mum cont i n u o us p o wer was use d f or ) t h e ta keo ff test s . E n g in e-o ut t est m g was co n d u c t e d us i n g t ak eo f f p o w er o n o n e e n gine a nd idl e on : t he ot h er.

St a bil i ty an d contro l te s ts were pe rf orme d to d eter mi ne the f ol l o win g : control s y s tem s c ha r - ' " _ ac t eris ti c s , cont r o l power; st at ic s ta b il it y ; dy n amic st a b i li ty ; acce l er at io n ch a r a cter is t i c s; a n d trim , ch_n g es. ., ' , < .

, L " , 8 A-541 8 Longitudinal t est maneuvers included the following: t rim change with speed (two and one engine): trim change with power and with nozzle delaection; elevator steps, reversals, ana d, 'ublets; .:, rapid p i tch attitu d e changes; wind-up turns; engine p o wer lever step,_:engine n o zzle lever steps; trim _,: change with flap deflection; phugoid; and trim change in ground effect.

_ L ateral-di, ectional test maneuvers included the following, with VSS _nd S, i S both on andoff: ' lateral control steps and reversals; directional control steps and reversals; Dutch-roll; spiral stab!lity: r apid bank angle changes; and t rim change with speed, o ne engine o perating.

P L !_, T e s t s w e r e als o c o nduct e d t o de t ermine the effects of partial hydraulic failure on lateral !_ control and _a SAS operation. The lateral control augrnentor chokes and the spoilers were deacti- !_ " vated separate.y t, _ te_*,their effect on c ontrol power: The directional and lateral control channels i_ of the SAS were deactivated separately, again to determine their effect on stability and control.

9 , P erformance i" The STOL take-off climb, transition, approach, landing, waveoff, and simulated slngle-engine _ performance characteristics were measured and evaluatec , by the project pil o ts. A Pulse Coded f: Op t ical Lan d i n g Aid ( P COLA) was used by t he pilo t s fo r gui d ance o n the appr o aches. A Fairchild _ Analyzer Camera was used for redundant measurements or takeoff and landing performance.

_ Operational _nd Handling Qualities I i_' ' ' Operational and handling-qualities testing was conducted in co n junction , v'th the other tests. _ ?_ ' _ In addition, STOL landing, transition, takeoff, and waveoff operation, and ground effect testing i _!_ were conducted as part of the evaluation of operational and handling-qualities characteristics, t .

_- S i ngle-eng'ne l a n d ing ap pr oaches and t akeoffs were also c o hduc t ed.

_,_ _ FL I GH T T E ST RE S UL T S A N D DISCU SS ION _*, _ Th e res ults o f th e fl ight inv es ti ga tt o n o f t he S TO L characteristics of the augmented j et flap " S TOL re se arch ai rcraft are d i s cus s ed und e r five g eneral categories: (1 ) aerodynamics, (2)stabhity an d c o n t r o l , (3 ) h an d l ing q u a lit i es, ( 4) p e rfo rmance an d c _ , ' , ra t i on al cha r ac t eristics, an d (S)engine- o ut co nt ro l a n d perfo rm a nc e .

A irc r aft Aer o dynam i cs '_ The aer o dynamics characteristics for the au_rnented jet flap aircraft, like those of conventit_ n al aircraft, are functi o ns o f flap deflecti o n ; .nd of angle of attack. Unique to the augmented jet flap aircraft, h o wev e r, are the additional parameter s of c o ld (fan or bypass) engine thrust, of the , a u g m e nt ed je t fl ap nozz l es , hot ( primary ) engin e thrust, an d h o t th ru s t v e ctor an gl e. The s e effe c ts are dis cu sse d h e r e , tog eth e r wit h t h e la n din g and tak e off operational envelopes t ha t resul t from t h e A - 5 4 1 8 9

l

,_ ._ . _ . _ __ ., . . r ._ • ._ _ _ _ _ o_ * - 4 L.. _ ' _ - ' _ .I ,._ -' Z _ _ . r., _ aerodynamic c haracteristics of the augmented j et flap aircraft, and the eff ec t of the ground prox- [ !

imity on the aircraft aerodyna m ics. . " - The ae r odyna m i c garameter _ - lift ,' drag, jet m o men tu m , thrust c oeffici e nts, and angle of attack - can b e co m ptlt e M from flight-m e asured quantities. It is difficult to derive generalized l ift _ and drag curv e s fro m individual data points, however, becaus e the jet momentu m an d thrust i Consequently, the lift and drag cu,-v e spr e sented are for varying v al u e s of C j a n d C T .

c oe f_cients( C . _ an _ CT )varywitheach_i gh tp_inta s ai rs peed an d ` at m _spheri cc_i t i _ns c ha_ge_ i _ Effects of ai r plane conf q _ u m tion on aerodynamic characteris t ics - Th e eff e ct of flap deflection ; on aerc l ynamic charact e risti cs is shown in figure 18 for flap an gle s of 5 .6° (flaps up), 3 0° (tak e off), " _ ; _ and 6 5 ° (landing). The engine nozzle an gl e for thes e data was about 1 50 , the en gi n e rp m was _ 94 p e rc e nt ( 6 f -- 5 .6 ° and 3 0 °) and 9 5 percent (6 f = 65°), and the alti tu de v ari e d betw ee n 2 4 00 : an d 9 4 00 ft (7 3 0 and 2900 m). Th e data ar e shown as a func ti on of aer od ynamic lift coef fi cient _ " _ CLA , which is d e fine d as th e fotal trimm e d lift c oe fficient minus the hot thrust ( CT ) contribution.

Similarly, th e d r ag c oe ffici e nt includ e s the thru s t of the a u gmented jet fl ap but n ot the hot thru s t.

_ The flight da ta wer e obtained during trimmed steady-state fli gh t at s e l e cted airsp e eds r es ulting in _ _ th e CT and Cj variation b e ing n e ady identical for each " flap setting shown. The Cj, valu e s in • _ fi gu r e 18 ar e lower than thos e us e d in STOL land in gs. B e caus e of the method for obtaining th e s e - data ,- the Cj incr e as e s with decr e asing airsp ee d ( in creasing an gl e of at ta ck) , which r e sults.in a l ar g e r _ . CL than would be produced b y a change in an gle of attack alone - th at is, with constant Cj . ?

The eff e ct of en g ine rp n : on th e a ir c ra ft _ erodynamics is show _ in figure ! 9 for a P.apd e fl e c- t _ n of 67°. A t a given angle of a t tack, a.n / incr e a se in e ngi ne rpm r e su lt s in an incr e a se i n Cj . , _ r e sulting in a large i nc r e as e in CLA an d some incr e ase in CD . Also , the lift curve slope incr e as e s wi th incr e asing engi . 'n e rp m .

_ Th e eff e ct 'of e n gi ne nozzle deflection on the to :-4 CLT and aerodynamic CLA of th e airpl an e f o_ -a _6 7° fl _ p angl e i s shown in figu re 20. Althou gh th e Cj and CT variations ar e not th e sam e for the nozzle an _e s shown ; a large chang e in total CLT occurs when the nozz le s ar e defl e cted. B e ca l_se _ - _ , . th e . C . / wa s not match e d for_ a l l nozzl e d e fl e c ti ons, th ee ffe _ t of nozzl e d e fl e ction on CLA cannot b e , , r e adily determined from thi. _ fi gu r e .. How e ver, pr e limina-y flight test data suggest that the lar ge r nozzle defl e ctions may caus e some reduction in a e r od ynan'ic lift or incr ea s e in d ra g, or both.

_ Landin g and takeoff aerod y namic cha r acteristics - Fli gh t-d e rived a e rodynamic charact e ris ti cs , - .

e ngine thrust, g eo m e t ri cal characteristic s , and the downw as h (r e f. 16) w e r e u se d i n the t ri m algo- l * rith m of re f ere nce i 7 to co m put e, the typi ca l tak e off and land i ng a e rodynamic charact e ristics shown in f_ ur e 21 for a s e a 4 e vel sta ndard day a t co n stant e ngine rpm . Th e cu rve s shown in this ' _ fi g ure re pr e s e nt th e ch an g e in aerodyn am ic param e ters produc e d by var y ing airsp ee d in l-g flight only. The s e curv es we re d e riv e d from fl ight t e_ t data s hown i n figures ! 8-20 . For th e l an ding - confi gu ration ( _ = 6 5 ° ) , 92 and 9 5 percent e ngin e spe e d ( N _ ) . charact e ristics ar e sho wn since th e . e ngin e pow e r us ual ly vari e d betw ee n th e s e two v al u es on a : -7. 5 ° approa c h path ( _ ). Th e charact e r- , is ti cs shown for th e tak e off conf'qguration( _ f = 3 0°) r e p _ nt conditions for n e ar m inimum take- , . off d is tan ce using takeoff pow e r (99 p e rc e nt NH ). _ t .; / 10 A- 54 18 ',..

" _ La n ding op erational envelope - Landing operational envelopes computed for various configura- -= tions for a sea-level standard day are shown in fi gure.s 22 through 27. These computations followed : ' _ the sa m e mat h ematical model as those for the li f t and d r ag curves discussed above. Figure 2 2 i • __ s hows the effect of engine rp m on the l a nding ope r ation al enve l ope ( _ f -- 6 5 °) a t a gross we i ght of : " __ " , 40,000 lb (18 , 000 kg) , and for the nozzles de fl ected 90 ° relative to the fuselage w aterline. A t a nominal STOL approac h condition of 60 knots and _ , = -7.5 ° , a ~ 3° and N H ~ 94 percent. At _= these conditions, the fli gh t path angle is changed by about , 0.3 ° per percent change in N H .

_ , Th erefore , large chan _ es in N H are require d for large c h anges in 7- Ffght p at h contro l m ethods are _ . _: Th e effect of nozzle angle o n the land in g ope ra tional envelope is shown in fi gure 23 for r _ S f = 65 °, gross wei gh t of 40,000 Ib (18 _ 000 kg) , and N H = 94 percent. For the nominal approach : discussed f urther i n the section on pe rf ormance and operat i onal charact e rist i cs. • condition o f 1 ' = -7.5 ° and a speed of 60 kno ts , t h e angle of a t tack and n o zzle angle a re about 3° and 85°, resp ee tivol _ . T h e nozzle is a v ery effective means of controlling flig h t pat h an gl e. : .

ar e based on a mat h ematic al m od el of t h e airplane. For verification pu rp oses , t h e _ e data were _ _ compared with speci fi c fl ig h t test poin ts . Flig h t test data on fli gh t pat h an gl e are co m pared in i T h e computed varia t ions in aircraft flig h t pat h a n gl e wit h airspeed shown in figu res 22 and 23 - figure 24 with th e computed performance of t h e airplane at the sa me gross wei gh t, flap angle, !_ " _ al titude , airspeed, engine rpm , and ambient temperature. T h e solid sy m bols"in t h e figure are fli gh t test points , and t h e open sym b ols at t h e same airspeed a r e t h e comp u ted v al ues. Wit h t h e nozzles at : : - " 12°, the fli ght test and computed val ues of'l _ s h ow very good agree m ent. Wit h t h e no z zles de fl ected to abo u t 9 0°, the measured fligi _ t test v a lue for 7 i s per h aps a little belo w the co m p u ted v al ue, . in di cating that there may be a nozzle interferen ce effect. T h is effect will be investigated fu rth er in futu r e fli gh t tes t s. (For the m os t part, h owever, t h e data s h own good agree m en t .)

Takeoff and climb operati onal envelopes - The t akeoff ope ra tion al envelope for the ai rp lane was computed using t h e m at h e m a t ic al model of the airplane d is cussed earlier. The effec t of engine rp m on the ai rp lane fli ght path angle for a wing flap an gl e of 30°, engin e nozzles a t 6° (fully up) , and sea-level standard day is s h own in fi gure 25 for gr oss wei gh ts of 40 , 000 and 45,000 lb (18 , 200 and 20.500 k g? . At full t akeoff power, t h e airplane h as the capabili t y of climbing a t an gl es of abou t 12° a t 45 , 000 lb (20,500 kg) and 16 ° a t 40 , 000 ib (18,200 kg). Th e airplane re ta ins significant climb p erformance even at reduced power setting.

T h e effect of en gi ne r pm on the climb operation al envelope at a wing flap _ n gl e of 5.6 ° (fully retrac t ed) for a sea 4 evel standard d a y at 40,000 and-45,000 lb (I 8,200 and 20,500 kg) grdss wei gh t is shown in figure 26. Wit h t h e 5 . 6 ° flap angl e , t h e airplane h as about 2° to 3° greater climb angle Fi gu re 27 shows fli gh t pat h a ngles computed from a ma th e m a t ical model of th e aircraft for speci fi c fli gh t t es t data points compared with measured fli gh t t es t fli gh t pa th angle at flap an gl es of capab il ity a _ t he engin e power .s e ttings s h own t h an wit h t h e flaps at 30°, even at a greater ai rs peed.

30 c ' and 5 .6 ° The measured valu es for fligh t pa t h angle were about 1° greater t han the computed • valu es for 30° flap an gl e . But at 5.60 flap angl e , the m easured values of fli gh t path an gl e were about I ° l e ss t h an t h e-co m puted v a lu es .

" Ground effect - Th e groun d effect with t h e airplane in t h e landing confi gu ration ( d r= 65°) shows a mark e d d e cr e ase in drag, a nosedown pitching moment, and a slight increase in lift as the !

A- 5 418 ! I l_

i -

i!

| airplane de sc e nds to ground level. Figur e 28 shows t h e eff e ct of ground proximity on t h e lift, drag, and pitching moment co e fficients obtained from several landings in whic h the airplane d e sc e nd e d - | slowly to touchdown. The discrete d a ta po in ts shown were derived from t e st d a:a (assum ed quasi- i s te ady _ tat e ) collected with f laps 6 5 °, CLA = 2. 3 and 3 .0, and engine nozzles d efl e ct e d betw ee n , : 5 0 ° 'and 80°. S e parat e flight condi ti ons were us e d for d eterm i n in g lift, drag, an d pitching mom e n ts gr ound eff e c ts. Th e e ngine power varie d b e tw ee n 93 an d 9 6 perc e nt NH , dep e nding on flig h t !i ; 7 condition s and gr os s w e i g ht. Figu re 28 also pres e n ts fairings ( CLA = 2.6 5 ) bas e d on a l e ast- s quar e s |_ , .

fit to th e e ntir e data for four ru ns us i ng a r e gr e ssion param e t e r id e ntification t e chniqu e . The da ta [ _ show good a gree m en t be tw ee n the two d ata me th ods. With th e airplane flying at ground level ( CLA = 2. 65 ), th e drag is about 6 0 p e rc e nt and lift 10 5 p e rc e nt of th e basic valu es . The n o s e down • pitching mo men t at gr ound l e v e l for th e sam e conditi o ns is e quiv ale nt to about 8° of e l e vato r ' d e flection (A CM =-0. 3 ). The v al u es of CLT s hown ar e bas e d on a nominal f light conditio n of 65 knots, 65 ° nozzl e d e fl e ction, and 9 4 . 5 p e rc e nt e ngin e rpm . There is e vid e nc e th at the gr ound e ff e ct vari es a s a function of CLA . The param e t e r id e ntifi ca tion proc e ssing of th e t es t data indi ca t e d the followin g c hang e s of gr ound e ff e ct with CLA at ground l e v e l:

:A c£dc ' A ) = -o .12

A CLA

|

W ° G/ CD a) = -o. 0 5 "

A C LA [ i

A cm o

! A CLA = -0 .3 0 Th e e xpon e ntial d e e _ y of th e ground e ff ec t with incr e asing altitud e ( h ) is e xpr e ss e d by e " kh .

Th e th e or e tical v al u e f o r the scal e h e i gh t ( k ) is approximat e ly 0.0 6 bas e d on th e asp ec t ratio (r e f. 18). Para me t e r id e ntification proc e ssin g _ i e ld e d a s ca l e h e i g ht valu e n e ar the th eo r e tical v al u e for CLG , and approximately 0.07 f o r CDG and for A CMG .

_ rcrafl Stabilit y and Control Longitudinal control - Th e l ongi t udin al control system us e d for th e s e t e s ts was t h e I _ asic Buff a lo sprin g tab syst e m modifi e d to r e duc e th e control forc e s . During th e initi al flight t e s ts , the pilo ts found that the lon g itudin al control dynamic f ee l c hract e ristic s w e r e un sa tisfa c tory, and the maximu m c on tr ol that could b e obtain e d at STOL a i rsp ee ds with r e asonabl e forces was only about + !0° to -! 7° e l e vator d e fl e ction ( 6e) , a lthou gh th e full trav e l of th e e l e vator is + 1 5 ° to -2 5 °. (This probl em is distain e d in mo re d e tail in r e f. 10. For sub se qu e nt t e sting, the ele vator control ry st e m _ wa s m odifi e d to a full y pow ere d syst e m. Data discuss e d h ere , how e v e r, p e rtain to th e ori gi n al spring tab syste m.) Th e staticforc e variation w i th d e fl e ctionissho wn i n f'qlure12. " 12 : A- 5 418 , . Elev a tor control pow er w as ass e ss ed i n lo n git u di na l c ontrol rever sa l ma ne u v e r s. A typi c a l c o n trol rev e r sa l a t 65 kn o ts i s s ho wn i n fi gu re 2 9. T h e v a ri ati o n o f p it ch an g u l a r acce ler a tio n w ith elev a tor a ngle at v a rio us sp ee d s for 6 5°. 3 0°, a nd 5 .6 ° fl a p s is g i v en in fig u r e 30, a nd m a xim u m t • pi tch r at es ach iev e d in th e r e v e r sa l m aneu v e r s a r e sh ovm i n fig u r e 31 .

I niti al p it ch sensi t i v i t y (pi tch acce l e r a tio n pe r un it d eflec t i o n , O/ _ c , s ho wn in fig . 30) is ab o ut 0 . 0 9 7 c a d / se e 2 / in, a t 65 to 69 k n ot s, 0 . 0 88 a t 6 0 to 64 k n ot s, an d 0 . 07 6 a t 57 t o 5 9 k n ot s, i n pi lot o p i n io n , th e sens it iv ity w as sa ti s f ac tory .

> . T h e m a x i m u m contro l p o w er co u ld not b e assess ed , of co urs e, bec aus e of th e l i m i t a t i on on ach iev in g m axi m u m c o nh' ol .- Th e p i lot s fo un d t ha t n or mal S TOL t a k e off s an d l an d in g s co u ld b e " a c h ieved with the a v a ilable elev a tor con tr ol . T h e dat a indic a te th a t a m ax im u m c ontrol p ower of , a bout 0 . 5 rad / s ec 2 was a vail a bl e a t t akeoff rotation and l a nd i ng a ppro a ch s pe e ds of 6 0 to 65 knots .

. Elev a tor deflection s as hi gh as -18 ° ( a bou t 1 5 ° a bov e trim) we re us ed in s ome takeoff s a nd fl a r es, ., b ut the pilot s did not r e por t a l a rge in c r ease in forc e or oth e r limit a tio ns d u e to t h ec Qn tfol s y s tem i n a ny of thes e m an e u ver s. I t i s not known wh e th e r knowledge of re s t ri cted e l ev a tor de fl ectio n , in flu enced pilot use of t h e co n trol . T h e ai r p l a ne w as re s tricted fro m conductin g s tall s b e ca use of th e :. dynamic feel characte ris tic s and low co n trol power a t very low a ir s peed re s ul t ing from the r e du c ed elev a tor de fle ctio n with low force s.

Long itudinal s tatic st abilit y - Th e sti c k- fi x e d s tatic longit u din a l charact e ri s tic s of the aircraft w e r e ob ta in e d by fl ying s t e ady ! g t rim C ondition s a nd m ea su ring elevator d e fl ec tio n. Test s w e r e " ! . c ond u ct e d a t fla p se ttin gs of 7 0 °, 65° , 5 0°, 3 0 ° , an d 5.6 ° with v a rio ufp o we r se ttin gs ov e r a r a ng e i :_ - of n ozzl e posit i ons . T he. c en t e r of g rav i ty ( CG ) v a ri e d b e tw e e n 28 . 9 and 3 0.7 p e rc en t of th e mean i : " a erodyna mic chord, and te s t weig hts ra nged from 3 5,800 to 4 6 , 6 001b (1 6,3 0 0 to 21,200k g ) .

¢ _ V a riation of CG po s ition with w e i gh t i s s hown in fig u r e 3 2 .

Figur es 33, 34, a nd 3 5 ill us trat e typic a l v a ri a t i on s of e l e v a tor d e fl e ctio n with fl ap d e fl e ction , n ozzle po s ition, and pow e r se tting r e sp e ctiv e ly. The s e dat a corr e spond to th e aerodynamic data given i n fig u re s 18, ! 9, and 20. F i gure 33 pr e sents measured el e vator-to-trim v a lu e s . F or t he se flight cond i tions, ! g tri m i s a chiev e d with elevator deflections of le ss t han 4° for the range of a irsp ee d s te s ted . L a rg e r u p e lev a tor de flec tion s were - re qu ired with lower fl ap a ngle s. Ma x i mum tri m el e vator u sed w as a bo ut - 7 ° with fl a p s u p ., N o a ttem pt was m a d e to dete rm i ne elev at or f o r m i n i mu m - .

"ai r spe ed because of the h aza rd o f app ro ac hi n g s t al l c o n ditio ns. .... _ St a ti c lo ng it u d ina l s t a b il ity l e v e l s w e re fo un d to be low , as e xpec ted . El e vator re qu ir e d to tr im a t a g i ve n a ir spe ed is giv e n i n fi g u r es 3 3 to 35 ; va ri a t i o ns i n s t ic k for c e , e l e v a tor d efl ecti o n, a n d an_ e o f atta c k with airsp ee d ar e gi v en in fig u r e s 36 a n d 37 f o r th e fl a p angle s of 67 °, 3 2 ° , a nd 5 .6 ° . W ith " : fl aps u p, both stick-fi xe d and s tick-free stability ar e positive . At 3 0 ° and 65 ° fl a p dd fle ction stick-fix e d stability is positiv e e xc e pt a t 3 p ee ds b e low abo u t 65 knots wh e re stability t en ds to b e co m e n eu tr al. Stick-fr e e stability ( s ti ck forc e required to c h a n g e ai rsp e ed) at c o n s tant _ ow e r i s non exis t e nt at both 3 0° and 6 5 ° flap d e flection s . At 3 0 ° fla p s and sp e eds below abo u t 80 knot s and wit hi n t he fri c tion b a nd . A t 70 ° fl ap d e fl e ctio n , s t i ck-fr ee s t ab il i ty d e t eri or a t e d f u rt he r, b ec om in g ". " !i i at 6 5 ° fl ap s below 6 0 knots, s tability e ven becom e s s lig h tly neg a tive alt h oug h t h e for _ e variation i s _ ' n e gativ e througho u t the landi n g approach s p e ed rang e. Th e low 1 9 ngit u din al s tability a nd "s pongy " I e levator r e spons e a t low spe e d s , combined with.poor sticl _ -c e nt _ 'ng capability (large friction b a nd), i mad e it difficult for pilots t o m aintain both pitch attitud e a nd ai rs pe e d durin g la n di n g a p proach e s. _, A- 5 418 ! 3 i _ 4r _ . . _ - .... i -, . J L on gi tudin al d y n amic st abil ity '- :l 'hc dy namic l on gi tud inal s ' h o :- t pe rio d sta bi l i ty c ha ra ct er- istics w er e approxi m ate d by ana l yz ing the respons e o f t h e a ircra f t to contr o l p u ls e s a n d step inputs. " Figur e 3 8 is a typical tim e his tor y of an e l e vator st e p at 6 2 k n ot s . Th e dat a i llu s trat e th e low short-period stabil i ty of the a i rcraft at th is _s pee d. Pitch ra te i s almost constant for th e approxi- , mat e ly 3 .0 s e c whil e th e controls ar e h e ld fix e d , and th e pitch rat e r e spons e of th e airplan e to the c o n trol input app e ars to be a lm ost first order w i t h a t im_- c 0 n s tatrt of abo u t 0. 5 se c. F igu r e 39 is a • - tim e h is tory of a phu g oid os c illation from a trimm e d c ondition of 65 kno ts with th e nozzl e : defl e cted 77 " . Airc ra ft short-pe ri od and phugoid c h aracteristic s ar e s u mmarized i n table 2 for _ se veral c onfi gurati on s .

L_ T h e p ilots w e r e n o t s atisfi ed wi t h the s h or t- p er i od dyn a m i c ch a r a c t e ris t i cs o f t he a i rcr a ft . - _ .

:: _ b ecause , of the hi g h p il ot workoad to c o ntr o l p i tch att i t u de and fl igh t pa th an gl e at ai rspeeds b e l ow / , . , 65 to 7 0 knots. Approaches _un d e r v is ual flight rules ( VFR ) w er e perform e d wi thout d iffi c u lty i n _ - winds as high as 3 0 knots w i th gus ts to 4 5 knots becaus e t he low s t ability r e sults i n only small _ . low-frequency disturbance; un d er such conditions, h ow e ver, t h e pilot w orklo a d i ncr e as e d.

[ Longitudinal maneuve r ing characteri stics - Th is secti o n covers the vert i ca l an d longi tu din al acce lerati on - responses o f the airc raf t to control inputs used f or fli gh t pat h an d ai rs peed c ontr ol.

_ . Th r ee control s ar e available to th e pilot _ in this propulsive lif t airc ra ft to achieve th e se r es pons e s : _ " . . (!) c h ang e s of angle of attack w it h elevat o r con tr ol, ( 2 ) c h anges i n engine t h rust w i t h th e t h rottles, ,- ?

and (3) chang e s in h ot thrust vector an gl es wit h nozzle control l eve rs . All t h r ee of th ese c on tr ols _ af fe ct t he lift, dra g , and pitc h ing moment of t he air cra ft.

'l / , The aircraft exh i b i t e d stable man eu ve ri ng c h aract e ristics at constant sp ee d, thrus t , an d nozz l e -_ angle. F i g ur e 40 compares pr e dict e d v al u e s f ro m re f ere n ce 1 9 wit h maneuve .r ing . data , .wh ic h are , " : J _ ' _ 4 / given in te rm s of angle o f atta c k, e l evat o r position , a nd st i ck " f orce per g va c at i on w it h airsp ee d o btain e d from w in d-up turns and pitch attitud e st e ps. Specific d.ata p o ints ar e pr e s e nt e d in _ ig- ur es 4 ! and42 for th e l and in g and tak e off configuratio ns, respectiv e ly. E levator-per-g i s c l os e to t h e predicted v al ue bas ed on . the w i n d-up turn data w it h h i gh valu es o f b et w een 2 5 ° and 3 0° per g at : *, airsp ee ds be t'_vee n 6 0 and 65 kn o ts. Fi gu r e 4 1 shows that'at 6 7° flap d e fl e cti o n stick forc e p e r g is clos e to p .e dict ed values up to 1.1 g, abov e w hi c h 'it falls be l ow t h e pr e dicted value. At 3 0 ° flaps, , t he stick forc e p e r g is well below th e p re d i cted leve l . Th e se data wer e ob tai ned w ith , the spri n g-tab e levator contr o l syst e m. In t h e fully po w er e d e l evator contr 9 1 system inst al led subs e quently , t he , f o rc es ar e a function of airspe e d and ele vator d e flection f ro m trim. . / ; : _ Load fact o r p e r unit an gle of attack b e com e s ' quit e Id l y as a irsp ee d _ r e duc e d (fig. 4 0). The : m e asur e d value of 1. 6 g / rad at 63 knots, which is r e p re s e ntativ e o f a landi s. . , appr(>ac h conditi o n, is . i clos e to the prechct e d v al ue; a mea m r e d point at 100 knots is ab o ut 4 g / rad. _ i • i _e Th_ _ riation of v e rti cal acc e l erati o n with c h ang es in angl e of attack obtain e d during pitch att i tud e c h ang e s ar e pr e s e nt e d in f i g ur e 4 3 . The d a t e s ho ws that for an airsp ee d of 6 3 knots, th e !!

sk , p e of th e variati o n o f & o r F wit h g is about th e sa me as that obtain e d in w i nd-up turns - for n e gativ e ¢ m d low positiv e v al u e s of A a F . At v al u es of AC _ F abov e 6°, h ow e v e r, v e ry littl e if any "i n.cr e its ei n "- .

, _ load fact o r is achi e v e d wit h incr e as e d an g l e of attack. A m aximum l oad f _ ct o r of only 0.14 i s -, _ : ._ ; avail a bl e at 63 knots w ith th e p i tch att i tud e chang es u se d in th e s e tests. At 6 8 k nots,, the maxim u m _ _ i " ,t al ue o f loa d f actor is 0.2 3 and th e curv e aga i n flatt e ns at ab o ut A¢ _ F = 6 °. At 98 knots, the ' " _: t _ " variation o f _ with A_ F is qu i t e linear to a A n: v al ue of 0,4 g. r _4 - ' . A- 5 418 / " , ') . / , Figure 4 4 shows tim e histor i es + of 5 ° and 10° pit c h r:ti itude c h a n g es a t a tr i m speed o f 63 k n _ o t s. Note that as attitu d e i s "increased t he a ir spee d decre _ es such tha t the pea k va l ue of i " , vert ic al accele r a ti o n does not occur a t the pea k value of an g le of _ * , tack for the 10° c as e, w hi ch _ accounts for the fla t tening c f the va ri ation of A n z with _ U z F in figu re 43. With at ti tude ch an g e s of : 5 5 ° , acceleration follows angle of attack more closely and airspeed ch an ges are much smaller and do " _ _. 5/ not a f f ec t the pe ak l o a d fa ctor . F i gu r e 45 p _e se nts si m ilar ti me hi sto ry d a t a for nega t iv e at ti tude ._ ch : l ges _ Th e r es po n se of the a ir p la ne wi th pitc h at tit ude ch an g es i s ty p i ca' o f ope ra t i on o n the b ac k : s i de of t h e t hrust requ i red curve .

_ F igure 46. show s aircra ft r esponse t o noz zl e rotat i on do wnwar d fr om th e af t posi tion, w it h t h e pilot controll i ng e l eva t or to _ hold a ng l e of at ta c k nearly constan t . Th_ effec t is somew h a t s i m il a r to _: a decr e a se of pow e r i n ?c- co n ven ti o n al a i r c ra ft beca u se t he m a i n e f fec t of vector i ng t h rust o n t h e _T ,-flight ch a ract e ri s tic s is i n th e 'd_ g d ir e cti on. Wh en th e no zzl e is mo v ed fr om l u ll aft t o n o rm a l t o .

_ " t he fl igh t path o n t h e app i 'o a c h, t h e li ft c han ge i s only ab o ut 2 0 perce nt o f th e we i gh t of the ; i_ a ir p lane. A ir_ eed dec r e a ses qu i c kl y a s the thru st vec t o r i s rot a ted from af t to ve rti c al p o s it io n , then _ f _ - i nc rea s es a s t he air c raft pitch es dow n a n d r a te of sh,k i nc r ea se s, e xciti ng th e phugoid. Th e new tri m : _ speed is lower t h an i ni t ial tri m . I f a ttit ude w ere he l d c on st a nt, a ngle o f a tt ac k wo u l d incre_ e and " J _ spee d wo uld f urt h er d ecre ase. An i n iti al in c r ease in posi ti ve ve r tical acce l e ration w ould a cc om pa n y ' a more ra pid down n ozz l e deflection. Th e effe ct o f ro ta ting noz zles from t he f o rwar d t o af t pos i tion is sh o wn in fi gure 4 7 . The aircra ft s ink s'mom e n taril y , abou t -0 .1 g, as t h e lif tin g fo rce due " _ : " ..

_ ; to t hru s t is r emove a , --th e n p i tc hes up, in cr eas es speed, a n d c l imbs . With e x ception of t he i niti al ! " - -.in / cr e ase d s in k ra te due to t he li f t ch an g e s , the res pon se i s si milar to a power i n cr e ase i n a c onven - : . _ , t i o n al air c raft. Fi gur ,, "48 i llust r ates si m i la r e ff e _. ts f o r s maller chan ges in n oz zl e posi t ion. _ -: - _ : Fo r w ar d r o tati on of nozz l es p r oved e f fec ti ve as a means o f decreas in g speed and fli gll t pa t h . for : !

I gli de .s l o p e in te r cep tign . On ce cn t he app ro ac h, m odu l a t ion o f t he no zzl e s in t h e r an ge 7 0° to 90 * , _ . _ J: provided a ver y ef f ec ti ve a ir s peed c o n trol, ta ki ng the p l a c e of c o n ven tio n al t h rott l es. Th e initi al _ te n de u cy f o r th e a i rcra ft t o s ink on aft vec t o r ing of th e noz zles , however, rest ri cted th e i r u se as a _ " means of c ontro l li ng fla re i n pr oxi m i ty t o the g r ou n d , an d the pil ot s p re fer r e d to r e vert t o th rottle _ a n d eleva t or co ntr ol for touchdo wn . Cau ti on was a ls o req uit ed i n vecto ri ng n ozz l es a f t on low i '_ _ al t i tude waveoffs to avo i d sinking , al thou gh it was sub se q u e n t l y found that i ncreasing pitch ' attitud e _ as no zzles mov e d aft effectively count ere d th e prob l em. • , ' \ _" i "_; A i rcraft r es ponse to cha n ge:, in thro ttle setti n g is co n ve hti on al in cruise con figu ra ti on w i th i nozzles a f t, but is unconven t ion al , as e x pec t ed, i n l anding c onfi gu ra ti on wi th flaps and nozz l es !

down. Figures 49 and 5 0 illus tr at e response to a step inc r ease ' and a step decrease i n power, ; r e spectively , in the l anding configura ti on. The " pilot is con ' trolling e l evator to maintain constant _ . _ pitch a t titud e . Inc r easing thrust p ro duc e s a l ifting force instead of an a x ia l force causing the ai rc raf t !

t o heav e upward. Figure 49 sho w s a peak of 0 . 11 g for a step from" about 92.5 t o 98. 5 per ce nt r . v m .

The da t a s h ow no in crease i n l o n g i tudina l acce l era t ion A X . F l i g ht path shall ow s in i ti all y , b ut the s ubsequent decrea s e in sp _ d tends to , w as h ' out th e ch an ge in fli ght pa th an gle. Similarly, th e . i _ sho r t -t erm effe c t of re duc i ng power (fig. 5 0) wi tl i a t ti t ude he l d co ns tan t is a decrease in l oad facto r i " ' and a st e ep en ed fligh t pa th angle ; bu t a s speed in c reases the fli gh t pa th tends to shall ow. With thrust _• _ , as primary flight pat h co n tro l this , . adverse s p eed -path coupling m ade it n e c e ssary for p i lots t o _ . continually mo ni t o r b ot h att i t u de a nd thrustto achi e v e desir e dcorr e ctionsin fl ight pa th and spe e d. I : Wh ere re quir e d flight path correctio ns w e r e l arge, pilots had to r e v e rt to nozz le m odu l ation, or to a _ ; co m b in ation of nozzle and th ru st control to achieve the desired response. _ +, A - 5 418 15 + + ) La i eral and dire c tional cont rol - Lateral control power wa s mea s ured by conducting lateral control reversals. . T h e whe el w as a pp fi ed in one direction a nd the n r a pidly rev e rsed a nd h eld in t h e • " opposite direction wit h" rudder p e d al neutral. Maximum acceleration w as measured as roll rate passed th r ough zero. A typical rever s al at 65 ° flaps with SA S off is presented in figure 5 ! arid wit h SA S on in fi gu r e 5 2. R e sults are compared Withpr e dic ti ons (r e f . 19) . in figur e 53 .

In th e approach configu i ation (flaps 65 °), . at approximat e ly 40,O001b (18 , 200 kg) gr o ss : weight, maximum roll accele r .ttion availabl e is about 0. 6 7 ra d / s e c 2 at 69 knots, sli gh tly above t h e _- predicted value andrepresen ti ng a rolling moment coeffici e nt of about 0.1 6 . Extrapola ti ng t h e data, _i " and assuming control power varies with dy n amic pressu re , we find t h at control power at 6 0 knots is i _ about 0. 5 1, well in excess of t h e d e sign crit e rion of 0. 4 rad / s e c 2 . t : : T he contribu ti ons of aileron, spoiler, and chok e to t h e rolling _ homent are illustrated in - ! fi gure 54 . Data ar e s h own for SA S on and off conditio ns . SAS h ad some eff ec t on maximum _ i control power because of t h e control defl e ctions due to SA S . For t h e maximum control condition .- , ! s h own i n figure 5 4, ail e rons a lone prod uce about 3 5 perc en t of t h e moment, t he spo ile rs produ ce perc e nt, and t h e c h oke about 3 8 percent. Lateral con tr ol sensitivity (roll acceleration per u nit * : wheel displac e m e nt) is about 0 . 1 ! ra d / sec 2 / in., co m p are d t o p re dictions of about 0.0 95 . The sensi- \ i tivity w as co nsider e d sat is factory by th e pilo ts .

Fi gu re 55 pr e s e nts roll accel e ration da ta m e asur e d at a gross wei gh t of 4 5 ,000 lb, flaps 6 5 °. : ; : Maximu m acc e l era tion available is about 0. 53 rad / sec 2 at 69 kno ts at t h is h igher w e i ght be c ause of ' the in c r e ased in e rtia with fuel in t he wings.

In _figu res 5 1 and 5 2, it can b e se e n t ha t with SAS on or off a t STO L airspeeds, yaw accelera- " _ : ti o n due t o lat e ral co ntr o l d e flection is v e ry s m a l l.

Fi gu r e 56 pres e nts roll a cc e le ra ti on data T o rthe 3 0 ° and 5 . 6 ° fl a p defl e ction co nfigu ra ti on s at gro ss w e i ghts of 44 , 000 to 45 ,000 lb (20,000 to 20, 5 00 kg). With 3 0° flaps, m aximu m rolling _ a cce le ra tion at 78 kno ts is abo ut 0. 53 r a d / s ec 2 , rep re s e nting a rolling m o m ent c o e ffici e nt ( C R) of _ ab o ut 0.14, whic h allow ing f o r t he h i ghe r w e ight is abov e th e pr e dict e d valu e . With flaps up, th e m easur e d maxi m um rolling a cce leration of 1. 2 7 rad / se c2 at 1 6 6 knots ( C £ = 0.07 5 ) is - mil c h h igher _ t h an p re di c t e d, due in par t to the flap d e flection at nomi nal flaps up b ei ng 5 . 6 ° in s t e ad o f 0 °, a n d b ec au se t he predictions did not c o ns id e r the lower s u rfa ce o f t he c h o kes ac ting as a il e rons.

Maximum roll rates ach i e v e d in th e wh ee l r e ver s a l man e uv ers , flaps 65 °, ar e giv e n in fi gu re 5 7 for SAS _ on and off c onditions at about 40,000 lb g ross w e i gh t. Max im um rate availabl e , SAS on, is abov e 2 3 ° / s e c at 70 kno ts . Minimum Crit e rionat 60 kno ts is 20° / sec .

M a xi m um co n trol s urfa ce rat e s ac hie v e d in r e ve rs al ma ne uv ers w e r e : , , w ; _ 20 0 ° / se c , _ 8 Ax L > 5 o° / C ., .

6 . s p 120° I se c

• • +i , 6 CH _ 9 5 p e rc e nt / sec " ., .

Full lat e ral control from n e utral position can b e achi e v e d i n l ess than 0. 5 s e c. _ : 1 6 A- 54 18 :_.

J ,= , . • . _ k . "

[

Dir ec ti o na l co n tro l p o we r w as e v aluated b y c on d u c ting rudde r reversal s w ith w h e el at neu [ ral.

fla p con di t i ons w i t h SAS on a nd off a re p r es en t e d in fi gu re 58. A t ypical d irecti on a l reversal wi t h z M axim u m yaw accele r ati o n w as measu r e d as ya w r ate p assed throu g h ze r o. Resu l ts for 6 5 ° a n d 3 0 ° , _: . SA S off i s shown in figu re 5 9 a n d wit h SA S o n i n figu re 6 0 . At 6 5° fl a p s an d an a i r sp ee d o f _ . 7 0 k n o t s , a ya w a c c e l e rati o n of 0. 3 ra d / se c 2 was mea s ure d a t 70 p er c e n t ru dd er d eflecti o n e q ui- _ v al ent to 0. 22 ra d / se c 2 w he n e x t ra po la t ed t o 6 0 kn o ts. This is high e r t han the p redicte d value o f about 0.26 tad / see 2 at 70 k n ots with maximum 25 ° ru dder defle_ction and is well above the crite- ri on o f 0 .1 5 ra d / sec sp eci f ied for a_lequate maneuvering at 60 knots. At 33 ° flaps and 80 knots, a _" y a w in g a cc elera t i o n of 0.33 ra d / sec 2 was a chieve d at 70 p erce n t ru d der. It is e x p ecte d t hat at • _ " ru dde r deflect i on s abov e 7 0 p ercent, t he accel e ra t i o n cur v e ben d s o v er fairl y q uic k l y as shown by : _ , t h e dott e d lin e s in fi g ur e 58. D urin g the r eversal maneuver, yaw acceleration tends to reach maxi- , m um before t h e ru dd er h as ha d su ffic ien t time to come t o it s c o mma nded p b s i t i on b ec a us e t ile rate / _ i _ is b uil d in g u p ra p i d ly. A t 70 _n ot s , m aximum r a te o f ru dd er m ov eme nt a pp ears t o b e .,b o ut 38 ° / see _ 5 , w ith f u l l ped al i n p u t . R o ll accelerati o n du e t o dir ecti o n al c o ntr o l , SAS o ,1, is s n iall ( fi g. 60 ) . • _.

, M axi m u m y a w rates achieved d uri ng re v ersa l maneuvers are s h ow n in fig ure 61 At 65" fla p s : _ a n d 69 k not s, a m aximum rate o f 12.5 ° / see w a s me a sure d with SAS off, an d a bo u t 13 ° / see with • ' , _ S AS o n. I n b oth case s , c ont r o l s w er e rem ov e d b e f ore maxi m um rates were achie v ed.

_ " Lat era l di rectiona l stati c st abili ty -, Thesestabili t y c haracteri s t i c_ were assessed by performing st ea dy sidesli p maneu v ers. Te s t r e s u lts are 'p re_sen t ed f o r t ak eo ff , a pp r o a c h , a nd c ruise c o n fig u ra - _ t i o ns in figure s 6 2 t hr o ugh 66. The 'aircraft exh i b i ts p o siti v e stabili t y about b o tl', la t eral an d d irec- • : tio n al axis for all configurations tested.

i " Figures 62 and 63 p resent data for the 65 ° fl ap condition. Figure 62 shows v ariation of rudder . d e fl ecti on , wheel d eflec t i o n, b a nk an gl e, an d ele v ator p osition w i th s idesli p ; fi gure 63 sho ws t he i' same data as a function of rudder angle and indi c ates a positive'dihedral effect. The data fall within •_ : t he ran g e of as pred i c ted bv w i nd t un n el te st s , c o rrela t i n g be tt er w it h C £ _ = -0.00 4 t han with

%= 0 . %

t t - _ Rudder an d w heel d eflecti ons are r ea son abl y linear w it h s i d e s li p a ng le s out to ±15 °. A t 65 fl ap s and 65 knots, sideslip prbduced by !0 ° rudder is about 22 percent less than t he p redicted v alue , indicating a higher directio n al s tabilit y than predicted for large sideslip an g les. At s p eeds bel ow 90kn o ts , t he .aircraft exhib i ted a l ow am p litu d e d ir e ct ion al s na k in g charac t eristi c I ( 3 = :1:1- 2*) indi c ating that near 3 = 0* dir e ctional stiffn e ss may b e ver y low. Attempts to do c u- ment any n onlin e ariti e s w e r e not su c c e ssful. About hal f a v ailable rudder a n d l e s s than half a v ailabl e w h e el thr ow a r e requir e d to achiev e 15° 3 at 65 to 70 knots. (Not e that abo_,; 100 k n ots sideslip is limit e d du e to tail l o ads.) In STOL con figurations, 15° o f s i deslip is a c hie v ed with bank an g le s o f

t

abo ut 5 ° , ind icat in g l ow n e t s idefor ces. Ra ti os of ba n k angl e t o s id es l ip a n gl e i n st eady -s t at e ] sidesli p s, _ / _, were a p p ro xim a t e l y i as f o llow s :

[

. , 6 5 d 65 93 0 .25 : |

30 6 75 9 0 0 .3 0

o • 5.6 6 120 P F LF 0.65 5 .6 6 .150 PF L F " 0.85 A-S'4 1 8 ' ! 7

t

,/ i " ; y The iow _' / O at l a nding a ppro a ch speed s below 65 knots res u l t s in very little lateral accele r atio n A .

with side s, ip . Th e a b sen ce of t he later a l acc ele rat io n c ue re q uires t he p ilot to us e the sidesli _ ' : indic at or i n the cockp i t mu ch m ore as sp ee d is reduced t o mai n ta i n low s ide s lip a ngles . Th e l o w i, d ir ectional stability a ggr a v a te s t h is proble m. . • " Lo n gitudi n al trim c han g e wit h sideslip i s s mal l for al l co n fig u r a tio ns , re q uiri n g less th an 1° - : elevator d e fle c tion w it h flaps at 65 ° and l e ss t h an 6° elevator deflection w it h flaps up and at 30 °.

Lateral directional dynamic s tability - Lateral directional ch ar acteristi c s were ev al uated at flap " settings of 65 °, 30°, and 5 . 6 ° wit h SA S both on a nd off ( SA S is au t o ma tic a lly off above 100 knots), a n d with c o m bi na tio n s of roll and yaw SA S on separ a tely in t h e landing configur a tio n .

: " . : Dutch-foil characteristics ar e s u m m arized in figure 6 7 a n d t a ble 3 : roll daml / iiag c har acterist i cs are : s h own i n fi gu res 68 and 69 .. " .- . A t th e landi n g app r o a c h fl a p deflec t ion of 6 5 ° . th e Dut c h -roll damping with SA S o ff is qu ite ! low with a d a mping r a tio betw e en 0.1 a nd 0.2. T h e D u tc h -roll period i s abo ut 6 see / c y c l e a nd is • close to the p re dicted v a l u e at 70 knots . Measu re d period is slig ht ly less than predic t ed a t 7 5 to 90 knots indica ti ng a trend to w ard greater directional stiffness t h an predicted at t h e h ig h er sp ee ds.

, A typical Dutch-roll, SA S off , is s h ow n i n fig u re 70 . T he os c ill a tion is primarily a y aw i n g motion : wit h I ¢ 1 / L 8 1 amplitu de ratio abou t i . 0. Wit h SA S on , Dutc h -roll d a mping ratio incre a ses t o about " ' _ 0.3. A typical SA S o n ti m e h ist o ry is "s h own i n fig u re 71 . Th e SAS o n pe ri od i s a bout 7 se c and i s less t h a n pr e dicted. With roll ax is SA S off a nd y a w a x i s SA S on it w a s dif fi c u lt t o ob tai n Du t ch-roll data beca u se t h e ; i ircr a f t exhibited a mar ked t e n dency t o " roll off " wi th con t rols fi xed; th erefore , table 3 does not show D u tch-roll da t a for t his ca se. The roll a x is SA S o ff h as li tt le effe c t on th e Du t ch roll ; w i th. roll ax i s S , I S on a nd y a w a xis S . A S off, D u t c h-roll c h a r act eris tics were simil a r to the all SA S off c as e.

A s noted in the discussion of stat i c s t abilit y , a residu a l direc t ion al sn_-ki / l gch ara c t eris t ic was , co m mo n at speeds below 90 knots even wit h basic SA S on. Pilots fo u nd this c hara cteri st ic objec- tion ab le. By using t h e VSS , / J -damping w as doubled , increasing t h e Dutc h -roll damping ratio to about 0.4 5 (fig . 67). Th e higher damping markedly i m proved t he pilo ts ' opinion of t h e directional cha ra cteri s tics at la n ding approac h speeds.

" Time histories showing t h e ai rcraft spir a l mode wit h flap s a t 6 7 ° deflectio n a nd nozzl es a t 15 ° : SA S off a n d on, are s h own in figu re 72. Spir a l s tabi l ity wit h SA S off at 65 ° flaps , nozzle s up, is a ns table as ex p ected wi th ti m e to do u ble amp l itu de of ab o u t 6 s ec at 6 0 to 75 k n ot s. W i th n o zz le L angle i ncr e a s ed to 90 °, time to do u ble amp li tude a ppear s to increa s e sli ght ly. P re cise spir al b ehavior, ; , w i th SA S off w as n ot a lw a y s di s c e r n ible d ue to sma ll lat era l offset s an d atm o s pher ic p e rturb a tio ns , • Wit h only yaw SA S on, s piral m ode w as similar to th e SA S off c ondition with a tendency to diverge more rapidly, T a being 4 to 6 sec. With only the roll SAS on , spiral stability i s positive. Ti m es to i' hal f an _ plitud e of 7 to !0 sec were noted at 6 9 knot s . Wit h all SA S on, spi ra l stability at 65° flap s is .: neutral t o s lig h tly positive. I ncrea s ing nozzle a- gl e to 9 b ° appe a r s ag a in to gi v e a s ma ll increase in st abil ity . M e as u re m e n t of pre c i s e tim es to c o n verge w ith SA S o n w a s diff i c u lt d u e to n o nlin earities ,, introd u ced by SAS inputs and / or atmosp h eric turbulence.

• _ Th e rol l damping at 6 5 ° fl aps with SAS off is low with an app are nt roll m od e tim e c o nstant ; O " A ) of about !.0 f s e c as det e rmin _ .d fro m roll r e versal maneuv e rs (fig . 6 8). SAS o _ ti me constant is _ approximat e ly 0.4 5 s e c. Th e pilots consid e r e d the roll damping with SA , . q on satisfactory.

18 A- 5 418 :_ ¢ "4 With tak e off fl a p s a t 3 0 ° , D utc h-roll d am pi n g, SA S off , is l ow a t 1 0 0 k n o ts , b ut with SA S o n, dam p in g inc r eases to a sa t is f ac tor y l e v e l of a bo ut 0 .3 i a t 79 k n ot s . SA S o ff p e riod at 100 k n o ts i s ab o u t 4. 5 sec an d SA S o n pe riod at 7 9 kn o ts i s ab o u t 6.4 sec. Th e sp ir a l mod e i s uns t ab le w i t h SAS • o ff (7 "2 _ 8 sec at 7 8 k n ot s) , an d sh o ws neut r a l to p o s i t iv e s t ab ilit y w it h SAS o n. T h e ap p a r en t roll mod e ti nle c o ns t an t i s ab o u t 1 .0 sec (fi g . 69) an d i s appa re n tl y ab o u t t he sa me SA S o n or o ff.

_ I n cr u i s e c o n fi g urat i on, D u tc h -ro l l ch a r a cteri s tic s ar e sa tis fa c t or y wi th d a mping an d pe r iod _ close to p r edicted level s. Spira l m ode is neutral t o pos it i ve , and apparent rol l m o d e time c o n stant i s ab o u t 0.8 s e c a t 1 34 kn ot s and 0.7 sec at 166kn o ts , b o th nea r p re dicted va l ues.

2 _ T ur n entry co ordin ation - T urn entries were c o nducted a t 65_, 30 ° _ and 5.6 ° flal_s t o eva l uate ai rcr a f t la tera ! contro l a nd t u rn en try ch a r a cteri st ic s. -Tests wer e carried o u t wi t h SA S both o n an d ' o ff, a nd wit h VSS syst e m i n op e r ati on wi t h se ve ral roll- ra t e - t o- ru dde r g a i n s . Tu rn c o or d i n at ion ._ cha r a c t eri s tic s , Aj 3 / A_ v enus a irspeed, are s umm a rized in figure 73. Wi th SA S on, th e aircr a ft

! [

exh ibi ts satis f a ct o ry tu r n c oo r di na tion cha r ac t eri s ti c s . Wit h SA S o ff, re sp o n se in tu r n ent r ies _ .

_ : degr a de s ma rkedly as a irspeed i s reduced below 8 0 k n o ts d u e t o a dverse si de s lip, giv i ng A_ / A ¢ r at io s _, _ : a bove t h e m ax im u m level of 0 . 3 g ener a lly consider ed accept a ble, refer e nce 20. !:_ ' _ Wi th 65 ° fla p deflect i on a t 65 -7 5 knots, SA S o f f, _ / AO rat i o s a re bet w ee n 0.4 a nd 0 .6. T he t ime hi story " of a tu rn e n tr y at 6 8 knot s ( fig . 7 4) sh ows t h e large adver s e sideslip generated on [ i n i ti a t ion of t he tu.rnentry . Y aw ra t e l ags roll r a te b y a bo ut _ s ec . T he l a rg e sideslip wit h b a nk angle _ ii " a nd t h e l a g between b a nk an g le and t urn r at e gre a t l y i nc re as e th e pilo t ' s worklo a d wh en m a ne u v er - ing laterally d u ri n g low s pe e d approac h es wi th out SA S . Wit h SA S on, tu rn c oordi na tion i s mu c h ii improved a s sh o wn in f i gu re 74 . Yaw rate n o w fol l o ws r oll r a te w i th a sma ll la g , and adver s e side s lip _ is red u ced. T h e SA S reduces A _ / A ¢ _ rat io s to a v a l u e o f a b o u t 0.3 , w h ic h t h e pilo t s con s id e r e d _ sat i sf a ctory. T h e stab il ity a ugme n tation improves t h e t urn coo r d i n a t i o n by de fl ecti n g th e ru dd e r propor t ion a l to roll r at e. Adv ers e ya w due t o la t e ra l c o ntrol is negligible i n bot h e xa mple s .

/ At 30° flap s wit h gA S off, some difficulty w as encountered in setting up initial s t eady-state _ . ,condit i on s, wit h _ v a r yin g + !° a bo u t a s tead y b i a s o f a pp r o x im a t e ly - 2 °. T h e A _ / A_ r at io d ata i n _ fi g ure 73, t h ere f ore, s h owed s c a tter wit h an aver a ge val ue o f a bo ut 0.37 a t 77 knot s . No adver s e _ _ y a w due t o la teral co n trol deflec ti on w as evident, b ut y a w r at e l a gged roll ra t e by 1 to 2 s ec d u e to t _ adve rs e sideslip, g i ving u nsatisfac t o ry t u rn coordination overall. With SA S on, adver s e sideslip an d y a w rate l a g w e re red u ced a nd t u rn coordina ti on was conside re d s a t isfac t o ry .

t

I n the fl aps u p configuratio n (w ith SA g au to m atically off a bove 100 k n ots) , _ / A¢ has a val u e of about 0.1 at 165 knots and 0.13 at 135 knots, in flaps up turn entries, no adverse side s lip w as • -e vid e nt an d yaw rat e follow e d roll rate wit h out lag, indi c ating satisfacto ry turn coordination.

Aircr a ft Handling Qu a lities eh . : . , • The aircraft h andling qualities evaluated in° t h e flig h t test progr am are di sc ussed in bot h qu a litativ e and quantitative terms. T h e quantitative values are c o m p a r e d tOestablis h ed criteria.

, Long i tud i nal stabilit y and control- T h e longitudinal stabilit y and c ontrol c h aracteristics of t h e i air c raf t durin g STOL ope ra tions were c on s id e red margin a l by th e p il ots. Once th e proper A-5418 19 . procedures had been determined, the pilots had no problem s acco m plishi n g V FR STOL takeoffs , - t ran siti o ns, a ppr oach e s, o r lan di n gs, b ut th e y co n sid e r e d the w or k loa d h ig h.

T he primary approac h pat h control is accomplis h ed by use of eit he r engine t h rust or t h rust J_ vector c h anges combined wit h t h e elevator. Large glide pat h correctio n s or low frequency respo n se t h at g i ve es s entially direct drag or th rust con t rol are best acc om plished by m ea n s of eng i ne nozzle , vec t oring, while sm al l ( h i gh f r eq u ency) co rr ection s a re mo st e asi ly ma de by t hrust v a ria t ion (e s sen- ti a lly dir ect lif t c o n trol ). T i_e e lev a tor re s po nse i n fli gh t path c o nt rol f or S T O L app ro a c h (3 , = - 7 . 5 ° ) c o n ditio n ( 6 5 ° flaps, 6 5 k n ot s) i s su m ma ri ze d i n ta ble 4 a lo n g with c ri te r ia f rom v a rio us , reference s. T h e pilots fo u nd t h at in addi ti on t o t h e s luggi sh load factor re s pon s e , t h e pitc h control se nsitivity ( 0/ 6 c ) is m argin al (0.08). The nosedo w n pitc h ing m o m en t i n ground effect effectively _ reduces the pi t ch con t rol sens it ivity below a s a ti s f a ctory leve l. .

i T h e eleva t or control sy st e m c h ar a cte ri stics a re con s idered margin al a t a irspeeds under abo ut : 65 knots. T h e decrea se d eleva t or damp in g, dynamic ' force c h ar a cteri st i cs , and e x cessive colu m n : for c es for elevator d.efle c tion s greater t ha n ab out 14° fr o m trim th a t red uc ed t h e u sable c ontr o l a u tnori t y were al l objectionable at th e l o wer airspeed s . The effect of eleva t or overb al ance on / stability , a t s peed s le ss tha n 60 kno ts t ha t re su l t ed in reduced c ontrol force s w as also objectionable.

At zero stick force, t h e elevator m as s bal a nce produces f u ll t railin g edge u p elevator at speed s up to = abou t 30 knots. These u n desirable chara c teris t i : cs (ref. 13 ) are asso c ia t ed wi t h the spring tab c on t rol sy st em and m ass balance; the spring tab control sys t em sub s equently was replaced wit h a fully powe re d h ydra u lic control sys tem . T h e elev a tor sy s tem c ha rac t e rist ic s at th e 65 knot approac h s p eed an d ref ere nce d cri teri a a re summa ri z ed i n t a bl e 5 .

; Th e e n gi n e sp eed c o n trol i s unsa ti s f act ory wh e n use d for fli ght pa t h co nt rol b e caus e th e thrust level to co n trol han dle ge a ring i n th e ap pro ach i s t oo s e ns i t ive an d has hyst er es i s. On e i n c h o f han dle movem en t gi ve s a chan ge of 4 75 rp m ( A T/W = 0 . 10 wh ere W = 3 8 , 00 0 lb ) at th e app ro a c h _ ' p o wer s etti n g. Hy st ere s is o f - +1 / 3 in . i n t h e engin e s peed c o ntrol s y st e m.wa s al s o q u it e o bjecti on - abl e , e s pec iall y with th e high sensi t ivi ty . Engin e ac c el e ration an d dece l e ra t ion tim e s in th e nom_al ' : oper a ting range ar e c on sidered acceptable f or co n t rol o n a STOL , _ pproac h , b ut t h e spooldown time . ) after th e t h rottle i s r e t arded o n to u c h down is to o lon g . Th e engi _ . _ nozzle co n tr ol s en s itivity a p d ef f ective n ess wa s s ati s f a cto ry . U s e of t h e nozzles d u ri ng flare wa s av o ided beca use o f t he i n iti al l ift los s an d t h e probability of ho t g as reingestion o n the gr oun d _ ,t large n ozz l e deflections. I f th e nozz l es ar e use d to co ntro l flig ht pa t h a t c onst a nt s peed, a c o ncu_ en t c han g e in pit ch at t itude m u st : accomp an y a n y si z able nozzle an g l e c han ge. Control of flig h t pat h by m e a ns of e n gi ne thr u st requ ir es little o r n o accomp a nyi n g a ttit u de c hang e .

_: Th e longit u dinal s tability i s satis f a cto ry o v er th e norma l o perat ing envelope exc e pt i n t h e STOL approac h speed re gi on below ab o u t 6 5 kno t s. Th e s t atic l o n gitudin a l stabili t y bec om e s ne u - _ t ra l and u n s tabl e a s s peed i s r ed u ced. Th e n eutra _ t o u n s table slope o f flig h t pat h wit h d e creasing : velo c ity is obje c tionable a n d in c r eas es the pilots workloa d on the apl :ro a ch . Ch a ra ct eri s tic s o f th e _: : lo n git u din al stability and c rit eri a ar e s umma ri zed in table 6 fo r t h e 65 k n ot : pproach case.

L at era l - di rectional s t ab ' ,lit y and control - T h e late ra l-directional stability and con t ro l c h arac- • teri s tics a r e satisfactory over th e oper a tional s peed ra n ge wit h $A8 i n the n orma l mode e x cep t for an objec t io n a ble low amplit u de directio n al s n a king at l an d in g ap p roach sp eed s a nd u n sa ti s f a cto q , :: lat e ral control forc e s. With SAS o ff in th e land in g approach (6 5 knots 65 ° flaps), turn ' f 20 A-54 ! 8 _1 i I I . i .. c oord ina tio n, Dut ch -rol l d am pi n g, r o l l da m pi n g , an d s pi ral d am pin g are u n satisfa c tory , as di s c u sse d in t h e previous section • Wit h t h e variable SA S ( VSS ) progra mm ed to increas e t h e " B e ta-dot " dam p ing, the ob j ec t ional snaking was r educed t o a sa t isf a c t ory level . Th e la t e ra l control b r eako ut _ f orce wa sab o u t 5 Ib ( 2.3 kg) ]wi t h f ri ct ion forccs of about +-1 . 5 lb ( +-0 .7 kg)l and because of the :, l ow f o rce gra d ient [ 1 0 Ib ( 4.5 kg ) at m aximu m wheel de flecti o n], the break out fe l t l i ke a r * r on g d e te n t e an d wa s objec t ionabl e . A mo d i fi cation t o corr e c t this c o n di t ion will be eval u ate d duri n g _. the next test pe ri od.

' : ._ C h ara c t e ris t ic s o f th e lat e ral an d dir e cti o na l c o n tro l s yst em s and c ri t e ria a re p rese nt ed in : , t ab le 7. T he l at e ral cont ro l po w . r an d s ensiti v it y are sa ti s f act o ry. A l t hough l ar ge w t . eel defle ct io ns .

power, i . are re quired for maximum roll con t rol 90 percen t of t h e effec t iven e ss is achieved at about ,:_ 6 0° wheel d e fl e cti on at ST OL s p - . e d s ( fi g. 54). The con fi guratio n of the l a teral control s y stem may , b e va rie d by selec ti vel y shut t ing o ff p ower t o t he spoil e r s and c h o ke s . Te s t s were condu c te d w i th ; _ the s po iler s and chokes inopera tiv e only i n the la nd ing a p pr o ach co nfi g u ra ti on . In the landing ap p r o ach_ t he la t e r al con t rol power i s sa tis factory w it h all con t r o l surf a ces o p er at ing; wi t h ei t her _ _ . t he s po iler o r choke s u rf aces ino p e r ati y e, ho w ever, i t is unsa t isf a ctory b e c a u s e of t he low se nsit ivi t y wi t h s p oiler off and nonlfneari t y wi t h' t he choke.off. Th.e low con t rol power and sensitivity wi t h i both s po iler a n d ch o ke o ff make the aircraft unacceptable for STOL ope.ration. The control power _ . an d sen s it ivit y f o r the v ariou s la t era l c ont rol con fi g o r_ t ion s ar e gi ven bel o w f o r t he la ndi ng _ . .

i a pp roach 1 65 kn ots , gr oss wei g h t 3 8 ,000 t o 40,000 Ib ( , . , ,800 to 18,200 k g) ]. T he lateral co n trol _ " q u i cken er " (fig. 17 ) , w h i ch is o p erat io nal w hen SAS is o n, w a s not sp ec i ficall y evaluat ed i n fli ght; : howeve r , pil o ts d i d no t commen t on an y no t iceab l e re d uc t io n of con t rol sens i ti v i ty or d ea d band : w h e n fl y ing wi th the S AS t u r ne d of f . , _ ¢ ' l a w , ]: max ( a w < 4 0 o) ra d / sec_ rad / s ec 2 / d e g ....... ,l , L A ll s u_ac e s op e r at ing 0.6 7 0.0129 S p oilers inop e r ative .47 .0075 Chok e s inoperative .40 { .0129 (8 w < 17°) .0088 ( 6 w > 1 7°) "- Spo ile r s an d c hoke s inop er a t iv e .22 .00 3 7 , Th e dir e c t io nal cont r ol powe r is consid e r e d s a tis fa c t o ry fo r a l l fli g ht co n ditions in cl uding • c r oss wi n d land i n gs when the cros s wind component is a s g reat as 20 knot s .

t The la te ra l and d ir ect .ional con t ro l power charac te r is ti cs wi t h refe r e n ced c rit e ria ar e sum- " _" m arizcd in tab l e 8 for t h e l a n d in g a pproac h and t ak e off conditio n s.The m a x im u m s i des l ip( _ m a x ) giv en in the table is b asedon a t a il stallt e ndency a t largesideslipa n gles. Th e de H a vill a ndCompany ,.

h as some evidence of vertica l s tab ilizer an d rudder stall b etwe e n 20 ° an d 2 5* sides l ip .

" T h e l ateral.dh c e . t ion al dy n amic sta b i l ity is sa t isfac t orywit h norr n _ ! SA S for a ll flight condi- i tions e xcept for the Io a amplitudesnakingmotio fi , which isexperien_:ed w it h the SAS operatingin tude is l o w wit h c h anges t, r yaw ang l e of onl y 2 ° o r I ncreasing t he Du t c h -ro ll' da m pi n g ra t io

• w , ,,, o o o r0. m i ., . ii

• from 0.30 to 0.45 did n o t e li mi : l a . t ¢ t he tend e ncy but reduced it to a level t h at the pilot considered A- 5 4 1 8 21 ' \ satisfact o ry. Wit h SA S off, the Du t c h -roll d ca r_, r oll damping, spiral damping, and turn coordi- n a tion :¢un s a tisfac t o r y in th e la n ding app r o a c h _:_d ta keoff. T h e later al -di r ectio v -J d v : L amic c h ar- act , _ristic s for t h e land in g approac h and takec . ff _,: ; ," , normal SA S are sum m ar i ze , . :n t a b l *. 9 w it h , crite r ia from various references.

' Ai r c,_ f t Pe r fo rm a nce ,:, ' ._ ' ) per at io nal C hara c t e ri s tics ; Thi s s ec t ion d i scu s ses the p erf. o r _ a . n ,'. : and o pe ra t io n al char a c terist i cs o f t he re search a ir - _ cr af t in th e S T O L t ake o ff a n d landin_ ! , iJ_e o f e pe ra tio n. C ruis e a n d con v e ntion al t ak e c ff and : l and in g da ta are g iv e n in refere nc e 13. T, , c p r e liminary ST O L t ak e o ff and lan d in g peffo rmeace data • ar e summ arize d i n t ab le 1 0 .

? , Takeo ff - The n o minal takeo f f confi g uration is w i th a fl ap d efle ction of, 3 0°, aileron s drooped = 1 7 "- a nd th e n oz z les fu ll a ft . Tak eo ff p ro ced ure i s t o a dv a n c e t h e t hr ot tl es to a sp ee d o f 93 to ; 96 p e rc e n t with t h e brak es l o ck ed, r e l e a s e t h e br a k es , and ad d powe r t o th e tak eoff en gi ne speed of 9 9 pe r c en t . Tak eoffs were c o nve n ti o n al f o r a n a ir cra ft w i t h a t hru st / wei ght ra t i o o f 0. 4 t o 0. 5 _ e x cept t hat ev e nts , occ urred m o r e rapid l y during th e t ake off m an euver. The e l ev at o r has suf fic i ent control power to ' start r o tation a t ab o ut 60 kno ts. Bac k stic k i s h eld u ntil lif t off, at a bou t 7 5 k n o ts, and 8 ° angle of attack , w h en t h e control is ea _ d forw ar d tc t ma in tain a constant an gle of attack until climb rate and velocity are bot h establis h ed. Bot h d ir ection al and later al c ont r ol ar e acceptable.

:!, A time hi sto ry fro m brak e r e l ea s e o f a typica l t a keoff i s- s h o w n i n ' figlJr e 7 5 f o r a l ow gr o ss .

: w eig h t of 39,2 2 0 lb (17,800 k g ). Th e d ist anc e trave l e d , S, i s c al c u lat ed from th e m ea sure d v e locit y .

( T he wind c on d i t i o n s w e r e calm.) Full th ru s t i s a c hiev e d a bou t 1 s e e aft e r brake r e l e a se . Rotation wa s initiat e d at 65 knot s , 7 se c after s tart. Lift o ff occ urr e d 8 _ e c after s tart at a v elocity of 73 knot s - and 51 0 ft ( 155 m) af t er star t; 82 0 f t ( 24 9 m) a n d 10. 4 see after s t a rt, th e aircr a f t c leare d 3 5 f t : ( 1 0 m). An an g l e o f a t tac k o f 8 ° w as reac h ed ju s t at lift off and h eld un til fu ll c l im b ra t e w a s es t a bli s h e d. Fi g ur e 76 i s a F a irchild A nal yze r C am e ra s e qu e ntial p hot og raph of anoth e r t a k e off. In th i s n o -wind t a k eo ff, th e fl ight path a n g le g radu a lly in c r e a se d u p t o 20* at about a 300-ft (90-m) a ltitud e . T h e p re v i ous f ig ure showed a st ea d y c limb rat e of ab o ut 19" by th e tim e a 60-ft (1 8 -m) altit u d e was re ac h e d. , "' Fi gu r e 76" g i v es so m e indicati o n or " th e n o se-hi g h a tt i tud e o f th e air c raft du ri n g th e climb o ut.

A n an g l e o f a tt ac k o f 2° a nd a climb ang le of 19" p l,._ ' csth e a ir c r a ft p it c h attitud e at 21" a b ove th e h o ri zon.

_ A compilation of s e v eral ta ke offs in f igu r e 77 shows th e e f fe ct o f t ake off we ight on th e : _ dista n c e _e q u ir e d t o c le ar 35 ft ( 10 m ). Th ese data were corr e ct e d for wind conditio n s b ut not _. te mp e ratu r e. Larg e vari a tions i n p e rfo r manc e a re p re s e nt , as ex p e ct e d , b u t an u pp e r bou n dary ie * fairly w ell d e fin e d . Th e data show that th e p e rformanc e is abo u t as pr e di c t e d ( r e f. ! 0 ) . Du ri ng th e s e : : t e sts no att e mpt w as mad e to d ete rmin e , th e Optimu m flap d e fl e ction for m ini m um take o ff p e rfo r- - _ mance ; 3 0° flap d e fl e ction was cho se n as n ominal t a ke off flap d e fl e ction for e n gin e consid e rations. _ , " t • Exc e pt for th e rapid rotatio n n ece ssary f or th e s e tak e offs and th e poor visibil i ty ov e r th e nos e !

of th e airc ra f t during th e i nitia l cl imb, th e pilots consid e r e d th e t ake offs comfortabl e , with l ittl e , change in control te chniqu es r e quir e d fQrST O L op e ra t ion. \ 22 A- 5 418 i J

i

(' limb and transition to cruise - Although the angle of attack during climb with takeoff flap is .__ low, from i ° to 5 ° , t he fligh t path angle of up t o abou t 2 0 ° contributes si_ ; n i fic ant l y t o a very steep _.i nose-up at t itude during climb. Visibility ove r t he nose was grea t ly restric t ed. Direc t i o nal and la t eral 1 . controls, remained acceptable, and no problems we t :e experienced during t u r bulen t and gus t y condi- _ : tions . Figure 78 shows t he measu r ed values of takeoff climb g r adien t for several flights that had i stabilized conditio n s. The climb values computed from aerodynamics data shown previously in .

figur e 25 are includ e d for comparison, but it sh o uld be no t ed tha t tile fl ag :,, _ data sh o wn were t aken ! : under nonstandard conditions, between 4C ° and 50° F, an'd have .not bee_ : corrected for this l- temp6rature difference. In general, the agreemen t is good, wi t h t he measured climb angle bei_ I _ : sligh t ly h i gher . ,_ i i_

, 1

_- Transi t ion t o cruise or flaps up climb configuration is straigh t forward as the flaps are r e t rac t ed I ' _ , 3 " i : _ . from t he t akeoff and initial climb deflec t ion of 30 ° t o full up. Some fligh t s we r e'made with the ! , ,; _ a, a t oma t ic t rim-flap in t erconnec t removed wi t h no significant change in handling quali t ies or pilo t t _ _ workload. Part of this is due to the low static margin and the very light control forces of the .

r elevator i_:its pre s ent configuration.

T ransition and landing a pproac h - Pilot ev alua ti o n of ap pr oache s a nd land in g s w i th 55 °, 6 5 °, w "_ a nd 7 2 ° fl a p d _ , f l ec t ions i nd ic at ed t ha t a fl a p d e fl ec t i o n o f 65 ° is near o p t i m um fo: t he ai r cr a f t o n i ' _ _ _" - 7.5 ° approaches . An increase in flap deflec t ion t o 72° results in very li t tle lif t increase. "l,i_ Transition to the approach configuration of 6 5° flap deflection and nozzle angle of 80° was docile, and can be accomplished in either level flight or by decelerating on the approach. Speed is reduced front cruise (150 to 160 kn,ots) to flap extend speeds (table 1) by either reduction of power , or d o wnw a rd r o ta t i o n o f the n o zzles , B o th meth o ds are satisfactory, as is a combination of the two. !

_ Al t hough a t titude t r ansien t s occur during th e configura t ion changes, no exces s ive con t rol inpu t s are required, trim is adequate, and pilot workload is not excessive.

• :_ A n experime nta l gl ide pa t h ind icat ion c alle d a Pulse C o d e d O p tica l Landing Aid (PCOLA) was - " _ used t hat i denti fied fo r the p ilo t t he sc he d uled glidesl o pe and ind ic ated angula r distance f , o m the _, o pti m um s l ope. Th e P C O LA is an o p t i c al d ev i ce l o cat e d ne x t t o t he runw ay and p r ov i des . _ v a r i_ibl e pulse width, indicative o f devi,,tion from es tablishe 4 1 g lide slop e . The p ilots ': an s id e r e d it useftl! and si mp le t o i nterp r et .

F The t echnique t h at evolved wa s to slow to 90 knot s while on a high base leg, with fla ps at 30° _ a n d n oz zl es full af t . F la ,- s we r e t hen low ere d t o 65 ° jus t prior t o or while in te rcepting t he - 7.5 ° _. . , glide s l o pe. A sp eed o f a bou t 9 0 kn ots was mai nt ai n ed a t an a n g le of att ack of about + 2 ° b: , _ - _' a d j ust ing p o w er. Sp eed wa s re duc e d at an al t i t ude of a b o u t 800 f t (240 m) spee d by movin g th e ] | _ n o zzl e s d o w n as req uired while i n creasing p ower to about 93 p e r cent a n d main t ainin g th e es ta b - lished g lidesl o l_ with nozzle deflecti o n ..t angl e of attack of 2*. Deceleration to 60 to 65 knots was us ually c o m p leted at about 500ft (150m ) and the r e mainder of the ap p roach was flo wn at co n st an t ai rspe e d . T h e use o f n ozz l e a n gl e mod ula t i o n ( v ec tor e d t hru s t ) at c o n st an t t hru st h a s . proven sat i s fac tory for sma ll or l arge g li de pa t h c o rr e c t i ons . W i t h n oz zl es, # i d epath a n g l e ch ang e s can b e m ade with l i ttle c i t an g e i n a n gle of a tt a c k o r air speed ( fi g . 23) . Wit h nozzl e fi x ed a t 80 * , the u se o f thrust for gl idepath c o n tr o l i s also a satisfactory te chn iqu e o rov i ded large gli de p at h c orre c - ti o ns a re not ne c e s s a ry . This technique per m its s m all gl i d e pat h ch a ng es w i th n early c onstant ai rs p ' ee d a nd a ttitude, but an gl e of a tt a c k chan g es ( fig . 22 ). A n i nc rea s e i n power o n approa c h at co nstant angl e of at t ac k _d ecre a ses both th e rate of des c ent a n d speed . " 1 h e us e t , f th ru s al one t o c o rr e c t for A- 541 8 23

• I

° large_ idepath angle changes is not ad e quate b e caus e of the airsp e ed chang e s, and it usually is neces- sary to r e vertto nozzl e angl e m odulation . The pilots' cons e nsus was that th e transit i on and a pproach had to b e ti ghtly controll e d to k ee p the workload acc e ptabl e . C orr e ction for larg e e xcursions from • no m inal valu e s of airsp ee d and / or gh depa th rais e d th e pilot workload to an unacc e ptabl e l e v e l b e caus e of the difficulty in controlling bo th ai r speed an d gl id e path an gle .

Most of the approac hes to a la nding w e rc m ade at a nomin al -7. 5 _ ' flight path an gle using the P COLA for guid an ce. Figu re 79 shows a tim e history of such a landing approach starting at an al titud e o f about 200 ft ( 6 0 m). Althou gh th e trace shows a gradu al incr e as e in angle of at ta ck sta rt ing be tw ee n 5 and 6 se _ b e for e touchdown , flare initiation as us e d in th e following discussion is / taken as th e time at which a _ harp change app e ars in the el e vator position, in this cas e almost 3 s e c : b e fo re touchdown. An an al titud e of a little over 100 ft (30 m) ther e app e ars to b e a gradual decr e as e in gl ideslop e an d a small d ec r e as e in velocity .

• o • T he vanaUon of approach sp ee ds with gross w e ight is shown in figu re 80 an d is re pr e s e n ta tiv e !

of obs e rv e d sp ee ds be tw ee n about 200-ft ( 6 0- m ) altitud e an d the flar e h e ight for a s e ri e s of : app ro _ ch e s on a nominal -7. 5 ° glid e path. C hang e s in the min im um ob se rved approach sp ee d with ; gross weight wer e su ch that l ift co effi cient rema in s approxi m ately constant. The p il o ts consider e d : the abilit y to make a p re cis e flare the limiting factor in choo si ng the minimum approach sp e ed. The : angl e-of-attack margin or th e airsp ee d margin from the stall we re not limi ti ng factors. The angle-of- attack margin on the appr oa ch was ov e r 1 5 ° an d was 10° to 12° in th e flar e . S in c e the load factor r e spons e to angle -of-attack chang e s ( nz/a ) is an impo rtant f ac tor t'orcontrol in the fla, e a ppropriate • v al u e s of nz/ a ar e sho wn in the figure . The band depicts the minim um approach sp e ed for acc e pt- - z bl e flight path contr o l in th_ l an ding fla re . Wh e n landing flar e s from flight paths of -7. 5 ° w e r e e x e cut e d be l'ow this s pee d, the chang e in fli gh t path that could b e produc e d in a re asonabl e tim e by ri t e ch ange in atti tu d e was in su f fi ci e nt to arr e st th e sink rat e to v al u e s l e ss than a bo ut 5 ft / s e c (I. 5 m / s e c). _ ; La , nd in g- Land in gs ar e b e st accomplish e d by l e aving th e nozzl e s at about 80° and fly in g th e l as t ! 00 ft ( 3 0 m ) with e l e vator and throttl e control• The s in k ra t e is arr e st e d by attitud e chang e , and pow e r is usually r e duc e d for th e a ctual touchdown. The po in t at whi c h fl are .i s initiat e d must b e pi c k e d pr e cis e ly for a good land in g. Th ere i s a s tro ng f ee ling that th e no se wh ee l is go in g to hit first if flar e is d e lay e d, and th e_e is a t e nd e ncy to fl oat if flar e is ac c omplish e d to o soon. The typical flar e an gle -of-at tac k chang e from a -7. 5 ° approa c h path was 10° with v e rtical acc e l e ration l e v e ls ! .2 g or l e ss. Th e load factor r e spons e tim e is also in _ portant. The lo nge r it tak es to attain th- load factor th e slow e r the flare. Bas e d on th e vertical respons e shown in f i gures 43 an d 44 an acc e l e ra- tion r e spon se tim e constant ( nz ), of approximat e ly 1. 5 s e c i s e stimat e d to b e typical of this airplan e wh e n al lowing for control t ra nspo rt and r e spons e lag e s. R e f e r e nc e 2 ! in dicat e s a r nz of 1. 5 s e c is about th e m aximum a cce ptabl e .

Fi gu r e 8 ! shows th e chang e in s ink rat e during th e fl are man e uv e r from pitch up to touchdown for a numb e r of landings and th e pr e dic t ed chang e s in sink ra t e for giv e n av e ra g e v al u e s of normal load factor nz , al though th e ac tual variation in nz with tim e is mo re n e arly tri angu lar th an r e ctangu- l ar . Th e s i0k rat e at touchdown is shown in fi gure 82. Only a f e w landin gs w ere mad e at si nk rat e_ IF e at e r than 4 ft / s e c, and th e s e w e r e consid ere d to b e e xc e ptionally " firm " by th e pilots.

_ Crossw in d la ndin gs pr e s e nt e d no difficulti e s. Th e rudd e r was v e ry pow e rful and allow e d pr e - I cis e d e crabbi ng just b e fore to u chdown. Pitch control was con si d e r e d sluggi s h at low speeds an d : i contribut es to th e lack of p re cis e attitud e con tr ol durin8 th e flar e and landing phas e s of th e I / , | 24 A- 5 418

!-

,a Z ¢ _ approach. Ex p erie n ce bu ilt s o m e conf i d e nce, but th e el e vator ce nt , el la cks th e cr is p, pr e ci se c ha r- [ : _ ac t e r ist i cs o f t h e l a t e r al c o n trol sys t e m , t ;" - I _ _ L " " M i n i m um gro u nd rol l la nding te sts were not c ond u cted b e c a u se o f t h e poor br a k e fe el s y s tem , i Ini t i al a ttem p t s to cond u ct rap id de c ele ra t i o ns on th e gro un d r esu lt e d in flat s pot s on the tir es d u e I ; _'_. to brake gra b a nd a d a nger o f b l owi n g o u t a t i r e. Ins t a ll a tio n of an i m proved br a king s y s tem wit h t _= a nt i -sk i d control is p la nned . F igure 83 sh ow s t h e es tim a ted la nding dista n ce ov e r 3 5 ft (! 0 m ) , , _ : which is felt to be represent a tiv e of the per f ormance th a t can be obtained wit h t h e pr e s e nt br a king I s ystem: t h e landing d is tance a l s o s how s a curve for a b ra ki n g coe f ficient of 0 . 38 , w h ich i s a n t _" e s tim at e of the p e r f or man ce po ss ible wit h th e a nti- s kid s y s t e m . • " i :. t - Wave o ff - W a veoff s w ere a cc o mpli sh ed eas ily , but t h e unique ground rule s of t his aircr a ft re quired non s tand a rd a ct i o ns . T h e procedu re develop ed i s to incre a se engine speed to 99 per c ent , !i rot a te t h e no zz les a ft , concurr e ntl y increa s ing the pi tc h a ttitude , a nd t h en gr a du a lly decre hs e t he flap deflection as s p eed inc re as e s . Th e t h ru s t m u s t be incre a s e d In ' st bec a us e t he tr a n si ent effect s o f _ c ha ngi n g noz z le po s ition are oppo si te to t he st e a dy- s t a te effe c t s. Deflection of the nozzl e s a ft to _ : _ a c h i eve mo re l o ngit u d i nal acc el era tion d e creases t he lift . A co n c urrent incre a se in en gi ne speed or • _ angle of att a ck i s therefore nece ssa ry to compen sat e for the mo m ent ary lo ss i n lif t as w ell as to -_ _: a rre s t t h e s ink r a t e for t h e w a veoff , i _ Engin e-Ou t Control an d Perform a nc e ; _. S i n gl e engin e fa il u r es w e re s i m u l at ed i n t he tes t p rog ram o n app ro aches an d t a k e of fs b y " _ a b ru p tly _ t a rding on e engi n e to fl i g h t i dle . T h e rel a tively long deceler a tion t i me ( 1 8 t o 2 5 s ec) " [ . peculiar to t h e engine s c a u s ed a v e ry m oderate onset o f a ny out of trim co n ditions .

L .' !. E ngine fa ;. lure o n takeoff - - F igure 84 i s a ti m e his to ry o f a s i mu lated e ngi ne f a i lu re on tak eoff _ a t a n a ircr a ft weig h t of 41,000 lb (18, 6 00 kg) wit h fla ps a t 33° a nd t h e no zz l e s vector e d a ft. T h e rig h t- h and eng i ne was c u t a t rotation speed, and t h e l e ft engin e re ma ined at 9 9.9 p e rcent rpm .

n .

i Recove ry re q u ires less t ha n 20° w h eel deflection and le ss t h a n 1 0° rui dder deflect :. S i milar . _- control input s w e re used in recove ry fro m anot h er s i mul ated f a il ure cond uc ted j us t a ftdf liftoff a t a weig h t of 44,400 Ib (20,200 kg) and a speed of 88 knots. In general, pilots d es cribed aircraft , b e h avior during th e si mu l a t e d fai l ures a s " very doci l e wR h. l ittle corrective control req u ired, " an d con s idered recovery from t h e s i m ulated failures to be no more de m and ing t h an for some c onven- tional t w in-engin e airc ra ft.

At 3 0° flaps and appro x i m at e ly takeoff t hru st, singl e -engine climb gradients we re low as predicted. M e asur e d climb gr a di e nts for se veral t a keoff weig ht s a n d flap se tting s are s h own in J" figure 8 5 , along wit h a predict e d v a lue for 4 5 ,000 lb (20, 5 00 kg) gross w e ig h t. At 44,400 lb i ( 20,200 kg) and flaps a t 32° , t he be s t single- e ngine rat e of climb a c h ieved w as about 200 ft / m in ( 6 0 m / s e e) at 88 k n ots. T h e operating engi ne w as a t 9 9 .5 p e r c e n t rpm a n d t h e o t h er a t fli g h t i d l e.

For an eme rg e ncy condition, t h e e ngine can be increased t o about 103 percent rp m , wh ic h w ould increas e t h e climb an gl e about one- hal f a degre e .

En g ine failure o n approach- Sin gle e ngin e failu re s w ere simulated on n o m inal - 7 . 5 ° landing " approac he s at altitudes bet w een 55 0 and 800 ft (167 a nd 24 3 m ). Flaps were at 70°, n ozzles were v e ctor e d v e rtically, and airsp e ed at e ngine cut w as about 70 knots. All si m ulat e d failur e s were A- 5 4 !8 2 5 _, . |t " r C • = f o llowed by a waveoff with p o w e r restored at 1 1 30- to 200-ft ( 3 0- to 6 0-m) altit u d e . No actual single engi m : l a ndings were m ade. The _ i m e histories i n figure 86 show a "'failure " fro m 9 5 percent N H " approach pow e r setting. Recovery. wa- accomplished by first increasing power on the good engine, t h en rotating nozzles a ft (re c overy technique A). Figure 87 s h ow _ a simulated failure from 95 p er- , . , c e n t Nil inwhi c hno zz l es are fi r s t ve c to re d aft, the n poweri nc re _:d on thego od engine (re covery + te c hniqu e B).

+:J - t in the i llu s t r ation of te c h n ique A ( fi g. 86). c utting the left engine increases rate of s ink from 95 0 ft ]m in to 14 5 0 ft / m in (200 m / min to 440 m / m in) as to t al vertical for c e is de c reased and airspeed tends t o in c rease. El e vator is appli e d, pit c hing the ai rc raft nos e up, and power is incr e a se d i_ on t h e good engine t o arrest si n k rate. As nozzle s are v e ¢ : c :e d aft , with airspeed at 6 5 kn o ts, s ink , r ate f u r the r in c reas e s to 18 00 ft / min ( 55 0 m / m in) u n t il t h_ aft ve c tored thrust ca u se s sp ee d to + inc re a se , p rod u c ing pos itiv e g a n d a d ecr eas e in sink rat e . Abo u t 20 sec are re quir e d to bring the ra t e of d e_cn t to z e ro. N ozzles a rc th e n m odulat e d to r ee stablish sp e ed and r e gain th e flight path.

+ Late ra lly, c utting the l e ft engin e ro l ls the air cra ft about 4° to the I _ ft due " t o the . asy mm etric hot _ . thn _ st m oment (part L dly offs e t by an opposite m oment f ro m asym me tri c .distribution of cold : + , thrust). An increase in ri ght -. hand en gi ne pow e r contribut e s further to th e left mome nt, whi c h is co rr e ct e d by applicati o n o f ! 0° to 20 ° rig h t whee l i np u t. As n o z z .*esare n o w ve cto r e d aft, t he r o ll - : _ m o m e, + t be c om e s a yawing m o m ent r equi r ing 8° of ri gh t ru d de r to c ounte r act sid e slip, and t h e : l ' w h eel de fl e c tion i s r e du ce d to level t h e win gs . Alt h ou gh control inputs are s m all and easily a cc o m - plis h ed, c onside r able coordin at ion of c ontro l is r e qui re d th r oug h out t h e re c overy procedur e . In a g en uin e e me rg e n c y with rapid asy mme tri c t h rust loss, pil o t workload w o ul d be h i gh .

in t he illustration of t e c hn iq a e B (fig. 87), sink rate incr e as e s rapidly fro m about 8 5 0 ft / min to 1400 ft ! m i n and ai rs p ee d inc re ases as t h e right-h an d engine is c ut and nozzl e s are ve c tor e d aft. Pit ch attitude is in c r e as e d and thrust o n th e opposit e e ngin e is advanc e d t o tak e off pow e r to ar re st sink _ + rate. A s s pe e d i nc r e as e s, t h e flig h t pat h s h allows. Lat e rally and dir ec ti o n al ly t he re co very pro ce dur e " invo l v es primarily ru dd e r input. As t h e e n g in e is c ut and noz z les ar e v ec t o r e d aft, t he+ momentary - ....

++ : r ol l m oment b ec o m es a yaw m o m ent, w h ic h is corre c ted wit h about 5 ° o f le ft rudder. As power ol _- t he g o od engine is increas e d, a further 5 ° of ru dder is applied in t h e _ a m e dii-e c tio0 . Tec h nique B is , .

+, , si m pl e r than A in t h at r e v e rs a! of wh _ l po s itio n i s avoid e d, but in a genu ine e 4 n e rg e n c y this pr oce du re m ay t e nd to giv e a lar ge r initial incr e as e in sink rat e du e to an initial d e cr e as e in H ft as nozzl e s ar e vector e d aft prior to an incr e as e in pow e r : Sin gle - e ngin e c li m b: , with flaps at 65 ° or : abov e we re not p oss ib l e with any te c hnique e x c ept at the very l o w gr o ss weig h ts, Waveoff from single-engine approaches requ ir ed only vectoring the thrust aft, increas in g air- ' . spe e d, a nd re tra c ting t h e flaps. The _o ntro l and p e rforman ce w e r e th e n similar to the e ngi ne- o ut _ • 'tak e off c as e . The se r eco v e ry examples ar e not n ec essarily optimum for m inimizin g altitud e lo s s or + + for r ee stablishing the approach flight path, but they do illustrat e th e pri m a ry co ntrol inputs us e d to, _ r eco ver fr o m an e ngin e l o ss. " + . . - _ _ + - ; : 26 A- 5 418 !

t RE F ER E NCES : !

. !. B r adfo rd H. Wic k a nd Rich ar d A. K_ hn, " Turbof a n S TOL Research at NASA. " Astronautic s a n d .A eJ on a ,,fi cs_ vol. 9, no : .5 , May 19 7 1.

. / 2. D, C . Whittley. " The Augmento r _ n g: A New Means of Engine Airframe Integration for ST i_ U'AJrc ra ft." _ "_ - , A I AA pap er no. 6 4- 5 74 pr ese nted at Fou r th ICAS Conference, Paris, 19 64 .

3. J.E. Middl e br oo k , H. C. T inney, an _ D. C. Whit : ley, "T he Evolutionary D e velop m ent and C u rrent Status of th e Augm e ntor Wing Conc e pt. " Pa pe r 700812 p r e s e nted at SAE N ational Aeronautics and Spac e En _ n ee ring and Manufa _ tbring M ee ting, Los Ang e les, Octob e r i 970. - • __ 4. G.S. K e ll e y an d R. P. G e r e nd, " Propul u on Systems for Commercial STOL Air _ raft. " AIAA paper no . 71 . 7 46 .

AIA A / SAE 7th pr opuhion Joint S pe cialist Conf e r e nce, Salt Lak e Chy, Jun e 1971. _ 5 . D.C. Whittl e) , , " The A ugmentor Wing R e s e arch Prog ra m : Past, Pr e s e nt, and Fu tur e . " AIAA pap e r no. 6. 7-741 ,_ - • T e nth A _,_ ) _ -American Aeronautical Conference, L _ s Angel e s, Oc tob e r 1967. . ._ .. , 6. D.G. Koe nig, V. R. Coniglia, and J , P. Mor e lli, " A e rody n amic C haracteristi cs' of a La rge Scale Model with an _ • _ i Unsw e pt Wing and Au gment e d Jet Flap. " N AS A T N 1 _4 610, 1968. - . .

7 . A.M. C ook and T. N. Ai ken, "Lo w Sp ee d A erodynami f C h a ra cteri _ ics of a Larg e Scale STOL ' -T ra nsport Mod e l with an Augm e nt e d Jet Fl ap. " N ASA TM X- 6 2,017, ! 97 !. _ , .

8. M.D. F ala rs ki and D. G. Ko e nig, " Aerodynamic Cha r act e ristic s o f a Large-Scale M ode l with aSwf , pt Win g and 'Augm e nted Jc " glap. " N AS A TM. X -62 , 029, 1971. " - 9. R H. Ashl e man at l d H. Skavdahl, _' Th e D eVe lopm e nt 6 : f an Aug me ntor Wing J e t STOL R e s e arch Air c raft , (Modified C-gA), vol. !. Su m mar Y , NAS A CR : ! 14 5 0 3, 1972 .

i0. H.C. Q uigl e y, S. R. M. Sinclair, T. C. N ark , and J . V. O 'Kee f e , " A Pr o gre s s R e port on th e D e v e lop me nt ofan Aug me ntor Wi mg 'Jet STO L R ese arch Ai r c ra ft. " SAE Pap e r 7107 5 7, N ational A e rona u tic and Spac e E _gi n ee r- , in g and Manufacturing M ee ting, Lo s An g e les , S e pt e mb e r L 971. ' - 1 , b 1I. R E. Spitz e r, P.C. Rums e y,and ! : 1. r C . Qu igley, "U se of tl _ e FlightSimulatorin th e D e signof a STOL R e s e arch Airc ra ft. " AIAAPap e r ' 72-7 6 2, A u g u st1972.

-t 2. R.C. lnnh andS. B. Ande n on _"_ o m parisonof Simulatorand Fl i g htR e sultson A u g m e ntor Wing J e t STOL : , _ R e u ! ar c l l A ir craft. "N ASASP- 3_ 0 , Pa per 2! . 1972.

•_ 13. H. Skavdaldand D. H. Patt e n . on, " Th e D e v e lopm e nt o _m Aug me ntorWingJ e t STOL R e s e ar c hAircr a ft (Modifi e d C-gA),vol. !1.Analy m of C ontractor's F lngh t T _ t, August1972.

1 4 . "J ohn A- _ o _ nway, " Th e De v e lop me ntof an Int e gralPropulsion Syst e mfor J e t STOLFli gh tR ese arch." Pa p e r , • pr ese nt e d at AGARDProp u kion attdEn e r ge tics Pan e lM ee ting o n V / STOLProp ulsio n , S e pt e mb e r 197 3 .

1 5 . D . L. I-larkon e n , C. F. Wint e rm e y e r, and F. L. Wri gh t , ' _ tati c T es tsof a 0.7 ScaleAu gme ntor Wing Flapfor • th e Modi fie dC-gAAirplan e - T e stR e sults and An alysis. " NASACR-114 3 1 5 ,1971.

Vo m ask e , **A u gme ngor Wing STOL 16. B. Cl e v e land,RichardF. and S. R. M. Sinclair , J e t R ese ar c h ' A _ rrraft DigitalSi mu lationMod e l. " N A SATM X-6 2,149 _ 1972.

A- 5 418 27 * ° • i L : o ° 17. Susan E. Post, "'Compute r Pr e grams for Estimation of STOL Tak e off, Handlin g , an d Static P e rf m manc e . " i 1 NASA TMX- 6 2,217, 1972. / _!

, " } ; 18. H.E. Hoe m e r, ,_ 'Fluid-Dynamic Drag. " Publish e d by th e author, 19 65 . :: : , I -- _ c -,-_ 19. R.E. Spitz e r, " Pr e dict ed Flight Charact e risticsof the Augm e n t or W ' mgJ e t STOL R e s e archAircraft. " N AS A " C R - 11 44 6 3 , 1972. " _, / 20. R.G. Inn / s, C. A. Holzha uae r , and H. C. Q uisley, " Airworthin e s s Com i de ra tions for STOL Ai rc ra ft. " NA2,A *- !_ .- TN !)- 55 9 4 , 1 97 0. i_ 21. Anon.: "V / S fOL H andlin g - Q uaht ie sCriteria , " AGARD R e port no. 5 77, D e c A _ mber 1 970. _ . ~ " t_ _ t 22. R.L. Allison , " D e sign Evaluation Crit e riafor Comm e rcialSTOL Tr e nspo n _. '" NASA.CR- ! t a a54:1972, i t ". - | _ 2 3 . An on,: " Military Specification - Flying Qualiti e s of Pilot ed Ai _ plaq e ,. " MIL-F-8 7 8 5 B( AS G) _ August 19 6 9.

24. Ano n.: '*T e n ta tive Airwo r thin e ss Standards for Pow ei_ Lift T ra nspor . tCat e gor _ -Airc ra ft.' __ XX , F e deral Aviation Administ ra tion, August 1970.

- | , 2 5 . Ano n.: " l _ta ry Specification - Flying Q ualiti e s of Piloted V / STOL Airc ra ft. " MIL- F -8 33 00, D e c e mb e r I , i " . .. 1970. " _.

! 26. P.M. Condit, L. G. Kinb re l, a nd R. G. Root, "lnfli gh t and G r ound-Bas e d Si m ulation of Handling Qualities of . i - Very Larg e Ai rplan e sin Landing Approach. " N A S A CI _ - 635 , 196 6 . [' I / _ t t I 28 A- 5 4 !8 : t_ J 3: r_ !

/ APPENDIX • FLIGHT TEST AIRSPEED , ALTITUDE, AND ANGLE OF A T rACK CORRECTION The nos e boom airspe ed a nd altitude pos i tion e rr or correc t ions a re pr e s e nt e d in figur e 88 for 5 .6 °, 30 ° , and 6 5 ° fl a ps. Th e p a ramet e r ( A P/qi ) i s the r a tio of the m easur ed st a tic pr e ssur e error ( _) : o th e unco rre ct e d impact pre-_u re ( qi ). The p ara m e t e r ., CL a " i s th e t rim li f t co effi ci e nt of th e aircr a ft , based on indicat e d airsp ee d.

The an gle of attack cali b ra tion i s presented in figure 8 9, also f or fl ap d e fl e ction s of 5 . 6 ° , 3 0 ° , and 6 5 °. In thi s figure, ¢ / is the m easured or vane de fl ection (indicated angle of attack) re f e re nced / fuselag e center li ne. ..

" 2 I to th e r o se boom measuring station, and ¢ F is t he corrected ang le of a t tack re f erenced to th e _ 7 _

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I t A-:;41 8 29 1, _ TABLE I.- RESEARCH AIRCRAFT CHARACTERISTICS Weights, lb (ks) Maxi m um gross 48 , 000 ( 2 1 , 800) Maximum gross (STOL takeoff) 4 5 ,000 (20, 5 00) Maximum landing (STOL landing) 43 , 000 (19 , 500).

Op e rational empty 3 2, 6 00 (14 , 800) Maxim um fuel 14 , 000 (6 , 3 5 0) " _ Areas, ft I (ml) Wing are a , total including ailerons flaps and 111 ft 1 (I 0) of fus e lag e 86 5 (80) Wi n g flap ar e a, proj e ct e d, including ail e rons aft of wing lin e 187.10 (17) " " Total ail e ron ar e a aft of hing e l i n e , including tri m tab 46 . 3 0 ( 4 ) , - Horizontal tail ar e a,tot al . - 2 33 (21) El e vator aft of hing e lin e 8 ! . 5 (7) _ : "V e rtic al tail ar e a, total .. 1 5 2 (1 4 ) " .

< Rudder aft of h in ge line s " - t For e 3 0 ( 3 ) . " , -_ Trailing • 3 0 ( 3 ) ... ; Dimensions and General Data ?

.o J Wi n g , ft ( m ) , Span 78 3 5 (2 4 ) : Root chord 12-58 ( 3 .8) Tip chord 7.7 4 (2 .3 ) " M e an a er odynamic chord 12.1 ( 3 .7) A e rofoil section :, Root NACA 6 4 3 A i r 17. 5 (MOD) : Tip - NACA 63 2A61 5 (MOD) , Sw ee pback _ at 4 0 p e rcent chord, d e g 0 _ ) " ' , Dih e dral , out e r wing only , d e 8 . -5. 0 (Note: Wingta pe r and dih e dral e ach start 17.6 ft from plan e of s ymm e try.) _ .

Asp e ct ratio , 7.2 ; Ail e rons, ft (m) " Span 11_ _ 0 " ( 3 .5) Cho r d aft of hi ng e lin e 2 . 01 (0.6) _" / Dist an c e from plan e of sym me try to ce m troidof ail e ron 33 .70 ( IO,2) A e r od ynamic balanc e , p e rc e nt 20.0 ' , Sl _ ll e n , ft (m) _ |t " i : _- Span , i 2 q ( 3 .4) _ ; _ .

" Cho rd _ i.18 (0.t) . °_ I_ . _ Position of hi nge lin e , pe rc e nt wing chord ( a verage) • 62.4 F laps , ft ( m ) l Span ( eac h l id e ) 2 3 .0 ( 7.0) . " !

!

Ch or d aft of hing e lin e 3 2 (! . 0 ) 3 0 A-$ 4 18 t ...................... _ -............. i , ?

.4: . ..

TABLE I.-- RESEARCH AIR C RAFT CHA RA CTERISTI C S - C ontinu e d H o rizon t al tail, ft ( m ) Span : 3 2 . 0 ( 9 . 7) Root c h ord _ 8.33 (2 .5 ) Mean aer _ lynan i ic c h ord " 6.2 5 (I .9) A er ofoil sec t ion Root NA C A 63A214 (MOD) t ( invert e d) Tip NA C A 63 -212 (MOD) (inverted) - ; " Sw ee p of l e ading e dg e , d e g " 4.8 Dih e dral,d eg 0.0 =.

Asp e ct ratio 4_4 Vertical tail, ft (m) Span - . 1 3.6 0 (4.1) '_ Root c h ord 1 4 .00 (4.2) Tip c h ord 8 . 3 3 (2 . 5) Mean aerodynamic c h ord ! 1.41 (3.5) - - : : Airfoil s e ction NA C A 6 3 f21 014 (MOD) " Sw ee p of l eading e d ge , de g , ,. 2_ .6 5 ) . : As p e ct ratio ! .2 Ov e rall he ight, ft ( m ) 28.7 (8.8) Ov e ralllengt h ( with n ose be _m of 16 ft), ft ( r p) 9 3 . 32 (28. 3 ) , Distan ce , wi n g MAC, I / 4C, to h o r izo n tal t ail MAC, I / 4 C , ft (m) 4 6 . 3 ( 1 4. !)

Distan ce , win g MAC, I / 4C, t o v e rti c al tail MAC, i / 4 C , ft (m) 4 3 . 4 (1 3. 2) Winginc i d e nc e angl e, d e g +2.5 Ht n izontal tail incid e nce angl e (ground adjustabl e ), d e g + ! .0 !

ControlSurface Deflections and Rates , : , Fl a ps 5 .6° downt . o72° d o wn : : ' 4 ° / s e e e xt en si on a n d retraction , Conical no zzl e s 6° to 10 4 ° (d o w n fro m af t of air c raft) _ ; . > O0 * / sec , Ailerons ± i7 ° abou t + 35* m a _ droop an g l e • _ .. 3 0* / s ee • Spoil e rs _ - 5 0° "" i ' , 10 0 ° / s e e " A ugme ntor chok e 55% c hok e g a p ar e a closure at 75° flap d e fl e ctio n _- 3 0° l s ec _, Rudd e r ± 25 ° forward segme nt : ± 2 5 ° trailing s egmen t _: -5 0° I s e c Ele vator ( u sabl e fro m trim) - 1 5 ° _ " "F I 0 ° ° A- 54 18 , 3 1 . - I j TABLE I. - RES E ARCH AIRCRAFT CHA RA CTERISTICS - Concluded Maximum D e sign Speeds, knots , a Dive s pee d (F D ) 180 , , cm_ speed ( r e ) !6 0 , 30 ° Flaps down speed (F_ 0 120 > 5 0 ° Fla ps down s pe ed (F f _5 0o) 90 TABLE 2.- AIRC RA FT SHOR T P E RIOD AND PHUGOID C HARACTERISTICS ` Short pe riod Phugoid 8 f, U , F E , Weighi, wn ' P' C O n' P' deg d e g I_ nots 'Ib kg Cj rad /s e c _ " _ " s ec rad / sec seo 67 9 62 41, 34 0 18 , 800 -- - 1.00 0.72 9.1 0.17 0.09 3 7.8 66 4 0 6 4 3 8,800 17, 6 00 0. 3 8 .87 .9 0 16.5 .19 .09 33 .1 67 77 6 5 37, 6 00 17,100 .49 1.0 3 .90 14.1 .19 .10 33 . 6 33 14 78 40 , 000 18,200 .22 1.20 .86 10.l .18 0 3 1.9 / f 3 2 A- 5 418 i;l

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• TABLE 4.-ELEVATOR RESPON S E AND FLIGHT PATH CONTROL FOR TH E _ STOL APPROACH CONDITION. • _ ' ( AWJSRA Cri t erio n ( re f.) 11 Elevatorcontrol pow e r l!

(Fco I -- 40 lbs), rad / sec 2 0.33, -0.34 > 0.0 5 toO.20 i 21) Pitch control se nsitivity ( 0/ 6 c ) , rad / se c _ / inch 0.08 > 0.08 to 0.12 (21) • Stick forc e p e r g, lb / g 3 0 < 20 to 40 , > 3 (21) • t .

Effe ctiv e vertical sp ee d crossov e r tim e (t _ ), s e c 0.9 _ 0.8 (22) Load factor r e spons e time constant ( r nz), se c ~ 1.5 _ 1. 5 (21) Load fact o r p e r un i t control d e fl e cti o n (nz / 6 col) , g / inch 0.1 5 ~ O.! (22) - i _ Load factor p e r unit angle of t ! attack (nz la ), g / rad 1. 5 _ > 2. 3 (2 3 ) " _" Flar e control l o ad factor, g > !.1 - 1 .! 5 > 1 . 20 (22) Pitch angl e after ! sec , .t d e g ( F co ! se 40#) 2.2° > 2 to 4 (21) - / TABLE 5 .-ELEVATOR SYSTEM C HA _ CTERISTICS AT STOL AP PROACH SPEED. ..

AWJSRA Crit e rion (ref . ) Elevato r d e fl e c tion from trim @4 0# Fco1, d e 8 -1 3 , +!4 - Control c olumn d e fl e ction, th. ' +7. 5 , - 5 > ± 4, < ± 6. 5 (21) Elev ator t o column - ,: g e mn8, d e g / ln. 3. 0 - , Forc e gradi e nt( / _ e <l 0 ° ) lb / in . 4.8 _>2, < 5 (21 ) " i C o lumn forc e at TEU ; , ! maximum uu bl e 8 e , ib 4 0 _ 7 5 (24) 34 A- 5 418 • . ......................... • ........ I =) TAB L E 6 .- , LON G I T U D IN AL ST A B I LITY C H ARACTERISTICS A T " ;TOL APPR OA CH S P EE D.

• AWJSRA Crit e rion (r e f.)

Stick-fixed stability • ( A6e / A V ) , deg / knot 0. I > 0 ( 2 4) Sti ck .fr ee stability " (A Fc/ A F), I b / knot 0 > 0 ( 24 ) Chang e of flight path with _ , elocity(A T / A V ). d e , kno t " 0 < _ 0.0 6 ( 2 3 ) Short-periodnatural frequ e ncy ( t on ) , rad / s ec 0 .9 > 0. 4 7 (2 5 ) ii L Short-p e riod d amping ratio ( _ ') 1 . 2 > i .0 6 (2 5 ) : ° Phu g o id n a tu r a l f re qu ency ( to p ), rad / s e c 0.2 - / , Phugoid damping ratio ( _ ') 0.1 > 0 (2 5 ) Changeof 0 with v e locity • (A O/tX lO, deg / knot , -0 . 8 < 0 ( 22 ) | ,,, ' 7 s • 1¢ • ?

• 3 A- 5 418 _5 f _" _ .... -' _ ....... _ ...... I i_ J _ i " ,j TABLE 7.- LATERAL AND DIRECTIONAL CONTROL CHARACTERISTICS.

AWJSRA C rit e rion(r e f.)

3 t Late ral C ontrol wh ee l trav e lat max 0 , d e g ± 77 . "± 6 0 (20) Control trav e l, in. (m) + 11.2 ( ± 0 . 3 ) - Breakout forc e , lb (kg) ± 5 ( + 2. 3 ) 1 / 2 to 4 (2 5) .

Forc e gradi e nt ( _ w < 20 ° )' Ib / in. (kg / m) 0.8 1 to 3 (21) Maximum forc e at max. 0 , lb (kg) ' 10 (4. 5 ) < 20 (20) • Control r e sponse (full control), sec 0. 3 < 0. 3 s e c. for 63 % _ m a x (21) Dir e ction al I Rudd e rp e dal trav e l, in. (n) ' ± 3 . 9 (: _ 0.1) 2. 5 to 4. 5 (21) _" Forc e gradi e nt( ± 2 5 perc e nt Rudder pedal), ib / in. (kg / m) 45 (800) 10 to 3 5 (21) Maximum force, lb (kg) 98 ( 44 . 5 ) < 130 (24) Control r e s po nse (full contr o l), se c 0 . 7 < 0 .3 f or 63 % _ max (21) p -

t

I'

¢- t •

, !

3 6 A. or 418 , Ii t _J / • !

T A B L E . 8. - LATE R AL AN D D IRE C TIO NAL C O N TRO L P O WER CHAR AC '_' . _iS T ICS " FO R LAND I N G APPRO AC H A ND T AKEOF F C @ N D I TI ONS. ; A W JS RA C ri t erion ( ref. ) ;

-

'_" l a n d ing T akeof f :< S A S S A S S A S _.

_ On Of f , On _ La tcra l _ # ) m a x, rad / sec :' 0. 6 7 0.7 0 0.60 > 0 ,4 ( 20) .:' • C t / S w d e g ] in.* i.0 1.46 1.25 >0.8 ( ,_6) , , _ T ime t o 30 ° _ , s e c* 2.4 i.75 1.9 _, 2. 4 ( 20 ) ' , _ _ ¢ / 6 w , rad / s ec 2 / in., _: _ 7 ( 5 w < 40 °) 0.10 6 0 .1 I 0 0.0 9 5 >0.0 7 ( 26) _ - 7 Dir e c tion a l ? ° _ _[ ¢ , max, rad / se c2 0. 3 7 0.37 0.40 >0. 1 5 ( 2 1 ) _.

' ,2 _ _m ax , d e g 2 0 2 0 20 > . 25 ( 22) _ 2 I Decra b ( A_2. o ), d e g 21 23 --- 10 t o 1 5 ( 22) _ - ' ": D ec r ab ti me ( A_ = I 5° ) , sec 1.6 1. 5 - - - < 2 .0 (2 I) " i _i S ens itivity ( S p < ! i n), , ¢.

ra d / s e ¢ 2 / in. 0.1 2 0 .1 2 0.13 > 0 .05 (2 1) - - *B as ed on a ramp w h e e l input (r am p tim e = 0.5 s e c. ) . , _i_

i /

A- 5 41 8 37 I . _ TAELE 9.-LATERAL-DIRE C IIONAL DYNAM I C CHARACTERISTICS FOIt " LANDING A _ EROACH AND TAKEOFF CONDITIONS.

; AWJSRA Criterion (re£) ' r_ Land in g Takeoff - , Dutch roll damping _( Od), I /s e t : 0 . 30* 0.16 0.36 > 0.087 (20) ;: Du tch-roB " 0.98 !. 0 6 - I.06 > 0.5 2 3 (2 0) - - frequ en cy ( _ d)' i /se e - .

Roll timecon s tant0 r R),sec 0.4 5 " i.1 1.0 < 1.4(23) _i Spiraltimecon s tant ' _ ( r s ). s ec _ O -0 . Z Z "-0 >--o.o35 (20) ; Tu rn coordination (_ 1 _ ) 0.25 0. 65 0. 15 < 0. 3 (20) - Dihe draleff e ct (d&w / d _ ') i.2 1.2 !.4 > 0 (21) Dih e dral e ff ec t ( F wi _ max) , lb 8 8 8 < 10 (21) - i -, win e Du tch-roll dampingparameter(g'C _ d) is incr e a sedto 0.4 5 . with the VSS e ngag e d in th e landing _ : app r oach confi guration.

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f i 3 8 A-$418, - . - i k t¢ e_ TABLE 10.- ST O L TAKEOFF AND LANDIN G PERFORI _J 'ANCE. :- i G r oss weight (max). lb 45 , 00 0 (20,500 kg) Gross weight (max), Ib 43 , 000 (19. 5 00 kg) i Takeoff Landing Engine thrust , p e rc e nt 99 ' Engine thrust , p e rc e nt 9 3 / F lap d e fl e ction , d e g 3 0 Fl ap d e fl e ction, d e g 65 4 : ' - j , t Aile ron droop, d e g 17 Ai l e ron droop, d e g 3 0 "- Nozzl e position , d e g 6 Nozzl e position, de g x 80 Rotation sp ee d, knots 65 . Approachsp ee d, kn ots 65 -" Lift-off sp ee d, kno ts 7 5 Touchdown sp ee d, kno ts t 60 C limb s p ee d, kn o ts 86 G r ound ro b distanc e ( e_ t.), ft 8 4 0 (2 5 6 m) • , • * t C l imb angle, deg 16 . . Total . distance from . t 35f t ( e st.), ft ' - 1200 ( 365 m ) Ground roll distan ce , ft 700 (21 3 m) L Total distance to 35 I t, ft I 100 (3 3 4 m) x / , _| # 7 4 O ,y ." , 1 ' e. • f _ • ° , ° _ q

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A' 5 418 / " • 39 , s . ° . =

• Bypass air Conical nozzles

_ . cross duct

g

: _ _' X Bypass air

_ Main Rear

' _ engine mount mount assy rear duc t , _ _ , °

Firewall Firewall

Airbleed exit Nozzle : ?

! engine bypass actuator _

Figur e 4. - N ace ll e i n s t a ll a tion . ! • ; Z 8 -- -1 ........ _ -

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: N T I , (rpm / , / _ ) " :

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(a) Primary mas s flow and gro = (ho _ J : - : _ st. i Ft&ur _ S. - Nominal SPEY 801-SF engin _ operat l nl param e ters.

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BLC nozzles

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.... High pressure engin e rotor speed, %NH .

I F igure 7.- Bypass flow distribution.

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l Rudder a ngle, deg ; :i [ Figure13. - Directionalcontrol ch aracter i st ic s,pedal f orce and gearing. , 2 -, J -. . _ _ ' 4 K I i i , " ) i ' 'f I • ' + P+ ' : ,, i m_--' m '_r+ "+ .... ; - - , ' I P + ,. + + + , , + , _ + f 50 , .

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60 80 I 00 1 2 0 140 1 60 180 : VE, knots : I :i g ur¢,+0 . - Maneuv e ring stabilitysummary , ' . _ .

T.E . u p

20 _ GW = 43.800 Ib

i Q , _ ,, (I9,900 kg)

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' Fi B ur¢ 4 1.- Wind-up turn, flaps 6 7 °, 77 knots. !

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'i i 1974014524

I r I t !

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_ • I6 i 1,7 J 3 8,6 0 0 (,7,500) on

• I61 I,O j 39, . _ KX) (m,,O0) y es

i

n I ¢ l ee l 40 , 400 ( m , _m o) o f f • ii _

. J3Z I 15 J 44,800 (20,300) off

_ I S' _ I S' Z I 4,,ooo (re. c oo) o f f

e i SZl 9 I 40,000 (,8,200) off _

: I 33 I 15 ! 43,800 (19 , 900) on

" I s I " i 44,000 (20 ,000 ) --

'_ ' I 6 I,II I 39,700 (180000) -- t5 -- " '..... i.......... " ....... - T-- : Predicted, SAS on • I ' . J ; . _

. j . .... * , _ i! __

= ,o -!-; ' --- , t! !,: Predicted, S b , S off t

Ok . i , i.. i I

< , . 5 ............. _ l _.

.4 i @ _ ............... , , I

;' _ .... i..... W ........... i ---T - , - -

' F : ¢m 019 i SAS off prediction, _ i _.

, . , ,. , , , _ - "_ AS offp r ed ic t i o n, ' T • " -u.w . e/ o eg 60 knots I

• ti

I

o , , .... _ i i I ,,, " 40 60 80 I00 120 140 160 180 Air s peed, VE, knot s f " F igur e 67..-- Dutch roll chara c t e ristics s umm a ry.

Sy m _f ' V E

deg • • 67 69 lib 67 70 ' &A 67 70-77 • 67 69 , I _ .67 70 ....

:, 5 0 68 70 1 " 1 ' _ Wt - 37- 41,O00: b S 116,800 - 18,600) i • " " n I ; I - _ . .

il I "

_I s A S on _

'_ = • • "

: - e. <> ' ; : _ S ASoff _ _: " _" O l ( I . _ , ."

: Wt • 45 300 Ib m 120,400 kg) / ql_ . - .i

_ 2 ............. - -- - _ --

_ : BI • , , Ol Ol SAS on _ • , i! _ I ; " "

_ -, ' 0 20 : 40 60 80 : 2

;_ : 1 9 Wheel, dig / . _. ' Fi g ure 68.- Rolldam p in 8 charact o fi s tic _ - flaps 6 5 °. ',

i

: Wt' = 44,200 - 45,200 Ib

" (20 , I00- 20 , 500 kg)

_ : = 5.6 °

e_

i• In _ . __ e® 0 . _ i66 knOts _ (1 . A = 0.68 sec)

_ ! , _ -- :] f n [] E] _ r_ ('rA = 0.7-

• _ I ....

( 134 knots I 0.83sec)

i ° L , t e

! 2 - l I

8 f = 3 Y Wt = 44,900 Ib 124,000 kg)

78 knots

• I - l, SASoff _ z _ i ( I rA = l.Osec)

SAS on

i

i , i i , • , ql I

0 20 40 60 80 I00

8 Wheel, d e g

Figure 6 9.- Roll dampingcharacteristics- fl aps 33 °, 5 . 6 ° .

L _ t Y

i U ) >>o o ooo 41 ._

= = - 2 -

e d 0 m _0. _" _ O . "

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