Document
NA SA TECHNICAL NOTE
NASA TN D-2966
--
c'/
A FLIGHT AND SIMULATOR STUDY OF
THE HANDLING QUALITIES OF A
DEFLECTED SLIPSTREAM STOL SEAPLANE
HAVING FOUR PROPELLERS
AND BOUNDARY-LAYER CONTROL
by Cnrt A. Holzhanser, Robert C. Innis,
and Richard F. Vomuske
Ames Reseurch Center
Moffett Field, CaZ$
N A T I O N A L AERONAUTICS AND SPACE A D M I N I S T R A T I O N WASHINGTON, D. C. SEPTEMBER 1 9 6 5 TECH LIBRARY KAFB, NM I Illill llllllllll l l l l l 1 1 1 1 1 1 1 1 1 1 l l l l l 1 1 1 1 1 1 1 1 ..
0079967 NASA T N D-2966 A FLIGHT AND SIMULATOR STUDY O F THE HANDLING QUALITIES OF A DEFLECTED SLIPSTREAM STOL SEAPLANE HAVING FOURPROPELLERSANDBOUNDARY-LAYERCONTROL By C u r t A. Holzhauser, Robert C. Innis, and R i c h a r d F. Vomaske A m e s R e s e a r c h C e n t e r Moffett Field, Calif.
N A T I O N A L AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $2.00 A FLIGHT AND SlMLTLATOR STUDY O F THE HANDLING QUALITIES O F A DEFLECmD SLIPSTl33A.M STOL S E A € " HAVING F O U R PROPELLERS AND BOUNI1ARY-LA.YE3 CONTROL By C u r t A. Holzhauser, Robert C. Innis, and Richard F. Vomaske Ames Research Center F l i g h t and simulator t e s t s were made t o study low-speed handling q u a l i t i e s , p o t e n t i a l STOL problem areas, and causes of deficiencies and t h e i r solutions. Tests of t h e STOL seaplane were made i n t h e 50- t o 60-knot speed range with Automatic S t a b i l i z a t i o n Equipment (ASE) engaged and disengaged.
During t h e simulation, s e v e r a l s t a b i l i t y and damping d e r i v a t i v e s w e r e varied and evaluated.
During t h e f l i g h t t e s t s , take-offs and landings were made from water a t 5 0 knots, corresponding t o a l i f t c o e f f i c i e n t of about 4. With t h e ASE engaged, t h e handling c h a r a c t e r i s t i c s of t h e a i r c r a f t were s a t i s f a c t o r y . The ASE provided roll and p i t c h a t t i t u d e s t a b i l i z a t i o n and increased r a t e damping about these axes. With t h e ASE o f f , t h e handling c h a r a c t e r i s t i c s were unsat- i s f a c t o r y because of low s t a t i c l o n g i t u d i n a l s t a b i l i t y , a very unstable s p i r a l mode, and l a r g e s i d e s l i p excursions during t u r n e n t r i e s . Response t o c o n t r o l w a s s a t i s f a c t o r y about t h e roll and p i t c h axes, b u t t h e l i k e r o t a t i o n inputs p r o p e l l e r s reduced t h e d i r e c t i o n a l c o n t r o l t o an u n s a t i s f a c t o r y l e v e l .
The simulator t e s t s were u s e f u l i n providing a preliminary evaluation and Good c o r r e l a t i o n i n studying t h e causes of d e f i c i e n c i e s and t h e i r solutions.
w a s obtained between t h e simulator and f l i g h t r e s u l t s with t h e exception t h a t t h e s i d e s l i p excursions during manuevering were l a r g e r i n f l i g h t than on t h e simulator .
INTRODUCTION An STOL seaplane has been developed by t h e Shin M e i w a Industry Company, Ltd. (SMIC) t o provide t h e Japanese Maritime Self Defense Force (JMSDF) with a t e s t bed f o r exploring t h e p o t e n t i a l of a seaplane with low landing and take-off speeds, This a i r c r a f t , designated t h e UF-XS, u t i l i z e s Boundary- Layer Control (BLC) and t h e p r o p e l l e r slipstream t o f l y a t low speeds; Auto- matic S t a b i l i t y Equipment (ASE) f o r improved handling a t low speeds; and a h u l l with good hydrodynamic c h a r a c t e r i s t i c s . Preliminary Japanese f l i g h t tests indicated t h a t t h e s e features allowed t h e a i r c r a f t t o operate i n seas with 6-foot high waxes.
To further document and evaluate the STOL seaplane, a brief flight test program was performed by a U. S. Navy and NASA team. Prior to this flight pro- gram, a simulator study of the handling qualities in the STOL regime was made at Ames Research Center (using preliminary flight results and estimates of the UF-XS characteristics). This simulation was used for obtaining a preliminary evaluation,for studying potential problem areas and their solutions, and for investigating areas beyond the normal flight envelope of the UF-XS.
Previous flight experience with STOL aircraft reported by NASA in refer- ences 1 and 2 demonstrated that when good high-lift systems are used the pro- peller slipstream can be utilized to augment wing lift and reduce flight speeds. However, various deficiencies in the handling qualities resulted at these reduced speeds. Consequently, the primary emphasis in this report w i l l be on handling qualities at low speed rather than on performance. ?"ne results of the flight tests will be compared with the simulator tests. Further, since the aerodynamic characteristics of the UF-XS are similar to those of other STOL vehicles (refs. 1 and 2), comparisons with these aircraft will be included. Also, previously unpublished results of simulator studies using the characteristics of the Breguet 941 will be presented since the basic character- istics were generally similar to those of the UF-XS and a larger range of test conditions was covered in some areas.
Additional aerodynamic results and complete hydrodynamic characteristics have been reported by the U. S. Navy members of this evaluation team (ref. 3).
The flight tests reported herein were performed in cooperation with the Co., Ltd., of Japanese Maritime Self Defense Force and Shin Meiwa Industry Japan.
NOTATION drag coefficient, including thrust CD W lift coefficient, '=- CL qs yawing -moment coefficient Cn acceleration of gravity moments of inertia, slug -ft2 Ixx,Iyy,Izz product of inertia, slug-ft2 Ix 2 w1X.x , l/sec roll due to yaw rate, Lr ar aL/Ixx , l/sec2 d i h e d r a l e f f e c t , LP a P hL/IXx rad/sec2 maximum r o l l c o n t r o l power, Lga6%ax
a 6a
a 6r
roll produced by d i r e c t i o n a l control,
aNIIzz
a 6 r
W
a i r c r a f t m a s s , - slugs
g’ a M / h , l / s e c 2 damping i n p i t c h ,
as
p i t c h i n g moment due t o t h r u s t change,
a T/W
a M / I speed s t a b i l i t y , l / s e c 2 / f t / s e c a V a M / I
angle-of -attack s t a b i l i t y , d , l / s e c 2
aa
~MIsy
, l/sec2 p i t c h i n g moment due t o angle-of-attack change,
a&
aM1S.y
gemax, rad/sec2 maximum l o n g i t u d i n a l c o n t r o l power, a6e & N / I z z
, l / s e c
yaw due t o roll r a t e , &P W I Z Z , l / s e c d i r e c t i o n a l damping, &r a N / I z z
, l / s e c 2
d i r e c t i o n a l s t a b i l i t y , a p aN/Izz l / s e c damping due t o r a t e of s i d e s l i p ,
a (dP/dt ) ’
adverse yaw due t o a i l e r o n , aN/Izz , rad/sec2
aga aN/Izz rad/sec2 m a x i m u m d i r e c t i o n a l c o n t r o l power, ‘rmax,
a 6r
r o l l angular v e l o c i t y ( r i g h t roll, p o s i t i v e ) , rad/sec PR p i l o t r a t i n g p i t c h angular v e l o c i t y (nose up, p o s i t i v e ) , rad/sec f r e e -stream dynamic pressure, l b / f t 2 yaw angular v e l o c i t y (nose r i g h t , p o s i t i v e ) , rad/sec r r a t e of climb, ft/rr,in S wing area, f t 2 SHJ? s h a f t hors ep ower T p r o p e l l e r t h r u s t , l b T
t h r u s t c o e f f i c i e n t , -
Tc
is
t r u e airspeed, knots
v
W gross weight, l b p r o p e l l e r s i d e force, r o P , Y Prop
side-force s t a b i l i t y , m, l / s e c
YP a p w, l / s e c s i d e f o r c e due t o d i r e c t i o n a l control,
a%
a angle of a t t a c k , deg
angle of s i d e s l i p , deg
P reference l a t e r a l c o n t r o l surface d e f l e c t i o n ( r i g h t a i l e r o n E a down, p o s i t i v e ) deg lateral c o n t r o l wheel p o s i t i o n , deg e l e v a t o r angle ( t r a i l i n g edge down, p o s i t i v e ) , deg l o n g i t u d i n a l c o n t r o l p o s i t i o n (forward, p o s i t i v e ) , in.
t r a i l i n g - e d g e f l a p deflection, deg rudder d e f l e c t i o n ( t r a i l i n g edge l e f t , p o s i t i v e ) , deg rudder pedal position, in.
e p i t c h a t t i t u d e (nose up, p o s i t i v e ) , deg
(5 d e n s i t y r a t i o bank angle ( r i g h t wing down, p o s i t i v e ) , deg cp undamped n a t u r a l frequency, rad/sec wn Subs c r i p t s i indicated or uncorrected X l o n g i t u d i n a l a x i s lateral a x i s Y z v e r t i c a l a x i s max maximum DESCRIPTION OF AIRPLANE AND SIMULATOR Airplane The UF-XS eaplane ( f i g s . 1 and 2 ) has a high wing, four engine nd boundary-layer c o n t r o l (BLC) on t h e f l a p s and on a l l c o n t r o l surfaces. This a i r c r a f t , o r i g i n a l l y a twin-engined Grumman UF-1 amphibian, was extensively modified by SMIC t o lower t h e landing and take-off speeds and improve hydro- dynamic c h a r a c t e r i s t i c s . Table I contains a d d i t i o n a l geometric d a t a f o r t h i s a i r c r a f t .
H i g h - l i f t devices.- The wing w a s r e b u i l t f o r l a r g e r span f l a p s with f l a p - sur- type blowing and f o r a i l e r o n s with shroud-type blowing over t h e q p e r faces. A f i x e d leading-edge slat extended from t h e inboard n a c e l l e t o t h e wing t i p .
Controls.- For t h e landing and take-off configuration, lateral c o n t r o l w a s provided by a i l e r o n s , s p o i l e r s , and d i f f e r e n t i a l d e f l e c t i o n of t h e mid- span p o r t i o n of t h e f l a p . A rudder with BLC provided d i r e c t i o n a l c o n t r o l and t h e elevator, which had BLC on t h e lower surface,provided l o n g i t u d i n a l con- t r o l . A slat, i n a n inverted p o s i t i o n , was attached t o t h e fixed-incidence h o r i z o n t a l tail. The r e l a t i o n between maximum surface d e f l e c t i o n and p i l o t ' s c o n t r o l movement i n t h e STOL configuration i s given i n t a b l e 11. A gear changer was used t o reduce t h e maximum surface d e f l e c t i o n s i n t h e c r u i s e con- f i g u r a t i o n . All of t h e surfaces were actuated by a n i r r e v e r s i b l e , f u l l y powered hydraulic system.
The M E from an s-58 helicopter was incorporated t o provide a t t i t u d e s t a b i l i z a t i o n and rate damping about t h e r o l l and p i t c h axes, and t o displace t h e rudder when t h e ASE commanded an a i l e r o n deflection. The reference a t t i - tude f o r the.ASE was wings l e v e l and t h e h u l l inclined 6 ' nose up s o t h a t t h e a f t portion of t h e h u l l was p a r a l l e l t o t h e water. The maximum surface deflec- t i o n s and equations approximating t h e ASE output a r e given i n t a b l e 111.
Propulsion and BLC- -systems. - The four p r o p e l l e r s had t h e same r o t a t i o n , counterclockwise when viewed from the f r o n t . The inboard propellers were 1 1 f e e t i n diameter and were driven by Wright R-1820 reciprocating engines with a take-off r a t i n g of 1425 horsepower; t h e outboard propellers were 9.3 f e e t i n diameter and were driven by Pratt-Whitney R-1340 reciprocating engines with a take-off r a t i n g of 600 horsepower. The BLC a i r t o t h e f l a p s , ailerons, was supplied by two compressors (constructed by elevator, and rudder Ishikawajima Harima Heavy Industry) and driven by two General E l e c t r i c T-58 engines ( r a t e d a t about 1000 hp). Each compressor delivered 24 pounds of a i r p e r second a t a pressure r a t i o of 1.5.
Hull design.- The o r i g i n a l h u l l was lengthened f o r e and a f t and provided with a "T" t a i l . Incorporated i n the h u l l were spray suppressors, shown i n cross section i n f i g u r e 2, t h a t extended from the fuselage nose p a s t t h e pro- p e l l e r p lane.
Instrumentation. - An oscillograph and photopanel were used t o measure
f l i g h t and engine conditions and surface and c o n t r o l deflections. The air- speed, angle of a t t a c k , and angle o f s i d e s l i p were measured on a v e r t i c a l The angle of a t t a c k and angle s t r u t between t h e cockpit and fuselage nose.
of s i d e s l i p were not displayed i n the cockpit. Based on a l t i m e t e r and a t t i - tude values, it appears t h a t t h e corrected angle of a t t a c k (with respect t o t h e waterline) should be indicated angle of a t t a c k minus 6'. A t angles of a t t a c k near t h e stall, the indicated values appear t o be unreliable, probably because of t h e flow being a f f e c t e d by t h e h u l l . Comparison of t h e d i r e c t i o n a l s t a b i l i t y computed from s t a t i c and dynamic data i n f e r s t h a t t h e indicated s i d e s l i p i s g r e a t e r than t h e t r u e value (indicated s i d e s l i p may be of t h e order of 3/2 times t h e t r u e value). Based on wing-tip boom measurements, it appears t h a t the indicated airspeed measured a t t h e s t r u t is close t o t h e c o r r e c t airspeed.
SIMULATOR The simulator used i n t h i s t e s t w a s t h e Ames Moving Base Transport It w a s equipped with Simulator which has limited movement i n p i t c h and r o l l .
instrument displays and f l i g h t controls similar t o those i n t h e UF-XS a i r - c r a f t . A Dalto v i s u a l simulator, a closed-circuit t e l e v i s i o n system with t h e camera servo-driven over a model runway, projected t h e approach l i g h t i n g and runway as would be seen i n hazy, 1/2-mile v i s i b i l i t y . A l l simulated landings were on t h i s runway since t h e equipment could not simulate water and sea con- d i t i o n s . Figure 3 i s a p i c t o r i a l block diagram of t h e simulation. The cock- p i t had a maximum r o l l angle c a p a b i l i t y of 9' and was programed s o t h a t t h i s corresponded t o a commanded bank angle of 1 3 ' . The p i t c h movement was 14' up and 6 down and w a s programed t o correspond d i r e c t l y t o t h e commanded a t t i - tude. Six-degrees -of -freedom equations of motion were programed on t h e analog computer.
TEST PROCEDURES AND CONDITIONS F l i g h t Tests The f l i g h t tests were conducted from Omura Naval A i r Base a t Omura, Japan, under VRF f l i g h t conditions. The a i r p l a n e w a s flown a t an average gross weight of 32,000 pounds ( W / S = 38 p s f ) , and t h e center of g r a v i t y w a s about 22-percent mean aerodynamic chord aft of t h e wing leading edge. I n t h e landing and take-off configuration t h e inboard f l a p was a t 5 5 O , t h e midflap a t 30°, and t h e a i l e r o n s undrooped. This configuration had been chosen by t h e JMSDF p i l o t s t o have adequate a c c e l e r a t i o n i n t h e water during take-off and adequate wave-off c a p a b i l i t y during landing. Higher f l a p deflections were U. S. team, and lesser f l a p d e f l e c t i o n s were t e s t e d only t o not t e s t e d by t h e a s c e r t a i n t h e trim change. Some data were a v a i l a b l e a t other f l a p d e f l e c t i o n s tests performed by JMSDF.
from previous The a i r c r a f t ' s s t a b i l i t y , c o n t r o l and damping c h a r a c t e r i s t i c s were meas- ured a t a n a l t i t u d e of 4,000 t o 6,000 f e e t with t h e PSE on and off a t speeds higher than 50 knots. Tests could not be made with ASE o f f below 50 knots or with an engine out because t h e s e conditions were beyond t h e normal operational envelope prescribed by SMIC. All landings and take-offs were made a t sea l e v e l with t h e ASE on. The majority of tests a t a l t i t u d e were performed with power set f o r l e v e l o r s l i g h t l y descending f l i g h t .
Simulator T e s t s P i l o t e d simulator s t u d i e s of handling q u a l i t i e s were made f o r t h e landing The simulated character- approach configuration a t speeds of 45 -to 60 knots.
i s t i c s of t h e UF-XS were f o r a weight of 29,500 pounds, compared t o t h e nor- m a l f l i g h t value of 32,000 pounds. The p i l o t ' s t a s k was t o make landing approaches i n i t i a t e d under IFR conditions a t 500 f e e t a l t i t u d e on a 3 O g l i d e slope 2 miles from t h e end of t h e runway. The p i l o t flew on conventional instruments down t o 200-feet a l t i t u d e where a c l o s e d - c i r c u i t t e l e v i s i o n run- way display came i n t o view. Also included i n t h e t a s k f o r some runs were l a t e r a l side-step maneuvers t o acquire t h e ILS l o c a l i z e r p a t h which was o f f - s e t 170 f e e t a t t h e i n i t i a t i o n of t h e run, and l a t e r , t o c o r r e c t back t o t h e it came i n t o view. Control of lateral displacement runway center l i n e when was a l s o studied by maneuvering from s i d e t o s i d e of t h e runway (200 feet 20 and 50 f e e t . During some runs t h e e f f e c t s wide) a t an a l t i t u d e of between of crosswinds of 10 knots and turbulence of 1.6 r o o t mean square f e e t p e r second v e l o c i t y were studied. Runs were terminated a t contact with t h e simu- l a t e d runway or ground. ?"ne l o n g i t u d i n a l handling q u a l i t i e s during t h e f l a r e and touchdown p o r t i o n of t h e landing w e r e not evaluated because of t h e poor q u a l i t y of a l t i t u d e information presented t o t h e p i l o t . However, handling c h a r a c t e r i s t i c s during t h e flare d i d give some i n s i g h t i n t o t h e e x i s t i n g c o n t r o l and s t a b i l i t y . Main propulsion and BLC engine f a i l u r e s were simulated f o r s e l e c t e d configurations i n both t h e instrument and v i s u a l portions of approach.
RESULTS AND DISCUSSION The parameters presented i n t h e following s e c t i o n s were varied i n t h e simulation and t h e p i l o t ' s evaluation is discussed and compared with those obtained during t h e f l i g h t program. I n some cases r e s u l t s from t h e UF-XS simulation are combined with those from a similar unpublished study of t h e Breguet 941 on t h e same simulator using t h e same p i l o t s . These r e s u l t s are combined because a g r e a t e r range of parameters w a s s t u d i e d during some of t h e Breguet t e s t s , and t h e s t a b i l i t y d e r i v a t i v e s of t h e s e two vehicles a r e simi- lar. The d e r i v a t i v e s f o r t h e "basic" UF-XS were based on l i m i t e d f l i g h t t e s t data a v a i l a b l e before t h e U. S. Navy-NASA f l i g h t evaluation and on t h e o r e t i c a l and semiempirical calculations. The d e r i v a t i v e s f o r t h e "basic" Breguet 941 were based on extensive f l i g h t t e s t data. Table I V l i s t s these derivatives and a l s o gives t h e ranges of parameters t e s t e d , and t h e values measured during t h e f l i g h t tests reported herein.
Low -Speed Envelope Figure 4 shows t h e low-speed envelope f o r t h e UF-XS a t 5,000 f e e t a l t i t u d e i n t h e STOL configuration of 5 5 O inboard f l a p , 3 0 ° midspan f l a p , a i l e r o n s undrooped, and BLC on. The corresponding l i f t , drag, and t h r u s t c o e f f i c i e n t c h a r a c t e r i s t i c s a r e given i n f i g u r e 5. It should be noted t h a t t h e s e data are based on indicated angle of a t t a c k and speeds, and t h a t t h e 100-percent power curve i s an extrapolated curve. Th? majority of t h e t e s t s a t a l t i t u d e were a t 55 knots with power f o r l e v e l f l i g h t o r for. s l i g h t l y The s t a l l speeds were 54 knots a t l ~ w power ( t h i s low descending f l i g h t .
power, 34 percent of take-off power, w a s g r e a t e r than i d l e power) and 45 knots a t high power, corresponding t o maximum l i f t c o e f f i c i e n t s of 4 t o 6,respec- t i v e l y . At sea l e v e l , approaches and take-offs with adequate s t a l l margins were made a t and below 50 knots, speeds t h a t corresponded t o s t a l l e d f l i g h t a t a l t i t u d e . Such an expansion of envelope between operation a t a l € i t u d e and sea l e v e l w a s a l s o noted and documented i n more d e t a i l i n reference 1, which contained f l i g h t tests of a very similar BLC equipped STOL a i r c r a f t .
I n s u f f i c i e n t time w a s a v a i l a b l e t o t h e U. S. Navy-NASA team t o examine optimum take-off and approach speeds, descent r a t e s , and l i m i t i n g conditions.
The power-on s t a l l of t h e a i r c r a f t was very m i l d and w a s preceded by a mild buffeting. A slow uncontrollable r o l l - o f f occurred a t t h e stall; however, s a t i s f a c t o r y recovery w a s made by applying nose-down control.
The l i f t and drag c h a r a c t e r i s t i c s of t h e UF-XS w e r e very similar t o those of t h e NC-130 with BLC and t h e Breguet 941 (refs. 1 and 2 ) .
F l i g h t Control C h a r a c t e r i s t i c s Forces.- The f o r c e c h a r a c t e r i s t i c s of each c o n t r o l a r e given i n f i g u r e 6.
The l a t e r a l forces were r a t e d s a t i s f a c t o r y
- a p i l o t r a t i n g of 3.5 ( t a b l e V
describes the p i l o t ratings ( P R ) ) . The wheel throw of 1 0 0 ' was too great for STOL operation, and a wheel throw of 60° t o 7 0 ' would have been preferred.
The longitudinal f o r c e s were s a t i s f a c t o r y (PR = 3 ) . The rudder force gradient was considered t o be too l i g h t ( P R = 4 ) ; a 50-percent increase i n the gradient would have been preferred.
Response.- The following t a b l e presents t h e a i r c r a f t response character- i s t i c s measured a t 55 knots with ASE off and it a l s o gives the corresponding i s i n terms of i n i t i a l a c c e l e r a t i o n with f u l l p i l o t r a t i n g . The c o n t r o l power d e f l e c t i o n of t h e c o n t r o l from t h e trim p o s i t i o n . (When t e s t s were performed with p a r t i a l d e f l e c t i o n s only,data w e r e l i n e a r l y extrapolated t o f u l l surface deflection.) The damping i s i n terms of damping moment divided by moment of i n e r t i a . The response a f t e r 1 second w a s computed f o r a c o n t r o l input t h a t takes 0.2 second t o complete.
- _- ~ . ~ _ _
Control power, Damping, Response a f t e r 1 P i l o t I
Axes
rad/sec2 7 l / s e c --
r a t i n g -- - _ _ - L a t e r a l -0.8 +O. 5 Directional .O7 nose r i g h t - 0 3 .27 nose l e f t
-. 3
. 10 Longitudinal -1.1 .55 nose up -1.1 .21 nose down . - _ _ _ _ _ - The d i f f e r e n t responses i n t h e directionalmode r e s u l t e d from a l a r g e trim required t o compensate f o r l i k e - r o t a t i o n p r o p e l l e r s and maintain 0 ' bank angle; t h i s w i l l be discussed later. The e f f e c t of l o s i n g an engine w a s not considered i n t h i s r a t i n g . The d i f f e r e n t l o n g i t u d i n a l responses a r e due p r i - marily t o a d i f f e r e n t range of elevator deflections ( s e e t a b l e I) although some trim i s included. The t r i m requirement a t speeds below 55 knots f u r t h e r reduced t h e nose-down p i t c h c a p a b i l i t y ; however, even though t h e t r a i l i n g - edge down l i m i t w a s approached a t t h e stall, t h e p i l o t f e l t t h a t t h e combina- t i o n of pitchdown a t t h e s t a l l and t h e a v a i l a b l e p i t c h i n g c a p a b i l i t y was s u f f i c i e n t t o e f f e c t a s a t i s f a c t o r y recovery.
With t h e ASE on, t h e i n i t i a l a i r c r a f t response t o a c o n t r o l input w a s not After a f f e c t e d by t h e ASE 'because t h e l a g of t h e ASE w a s about 1 second.
1 second, t h e response w a s reduced because of increased damping and a c o n t r o l input t o r e s t o r e a t t i t u d e .
Longitudinal S t a b i l i t y Simulator.- Figure 7 shows how t h e p i l o t r a t i n g i s a f f e c t e d by angle-of- a t t a c k s t a b i l i t y ( M a ) and speed s t a b i l i t y (Mv) with M E o f f and a t a speed of A s would be expected, reducing angle-of -attack s t a b i l i t y about 50 knots.
caused t h e handling t o d e t e r i o r a t e . It can be noted t h a t t h e p i l o t t o l e r a t e d low s t a b i l i t y (even s t a t i c angle-of-attack i n s t a b i l i t y , +&). H i s tolerance of such low s t a b i l i t y can probably be a t t r i b u t e d p a r t i a l l y t o high damping about -l), and t o t h e f a c t t h a t t h e a i r p l a n e i s not e a s i l y disturbed
( M & + Mq
by gusts. I n c o n t r a s t t o what might have been expected, increasing t h e speed s t a b i l i t y (+Mv) made t h e handling worse because t h e p i l o t w a s l e s s aware of t h e improved s t a b i l i t y than i n the corresponding increased t r i m change with t h r u s t or power; t h a t is, t h e speed s t a b i l i t y and t r i m change with t h r u s t a r e i n t e r r e l a t e d .
where t h e These adverse handling e f f e c t s of trim due t o power may have been aggravated by t h e s t a t i c i n s t a b i l i t y i n p i t c h used i n t h e t e s t s . One way t o reduce t h i s detrimental e f f e c t of t h r u s t on pitching moment is t o couple t h e elevator t o t h e t h r o t t l e . F l i g h t t e s t s of such an interconnect a r e reported i n r e f e r e n e e l , where t h e l a r g e moment change obtained during a wave-off w a s eliminated. I n such a case it would be expected t h a t the handling would improve with increased values of MV i n c o n t r a s t t o the r e s u l t s shown i n f i g u r e 7 ( b ) .
The t e s t s with ASE on were made using an Ma. of 0.25 and an MT? of -0.5 f o r which conditions the a i r c r a f t w a s s t a t i c a l l y unstable with ASE o f f .
Engaging the ASE improved t h e longitudinal c h a r a c t e r i s t i c s from a PR of 6 t o 2-1/2. The ASE s a t i s f a c t o r i l y modified t h e a i r c r a f t c h a r a c t e r i s t i c s s o t h a t a given a t t i t u d e w a s maintained during g l i d e path changes and t h e airplane responded as though it were s t a b l e (an e f f e c t i v e i'k of -0.4 compared t o an of +O.25, ASE o f f ) , with 50-percent increase i n p i t c h damping.
M, F l i g h t . - Tne longitudinal c h a r a c t e r i s t i c s of t h e a i r c r a f t with ASE off a r e indicated by t h e v a r i a t i o n of t h e elevator angle with speed a t constant power ( f i g . 8 ) . These curves indicated t h a t a t 55 knots t h e s t a t i c s t a b i l i t y of t h e airplane w a s n e u t r a l . From t h r o t t l e s t e p data %I w a s found t o be -0.1, and then w a s estimated as + O . l 5 . The corresponding undamped n a t - u r a l frequency w a s calculated t o be about 0.2 rad/sec. This low s t a b i l i t y i s f u r t h e r v e r i f i e d i n f i g u r e 9 by t h e close correspondence of t h e calculated response (using M, = 0 ) t o an elevator step with t h e measured response f o r t h e case with ASE o f f . It is a l s o seen t h a t t h e pitching v e l o c i t y increased u n t i l corrective a c t i o n was taken. Generally, t h e phugoid motion could not be excited; on the occasion t h a t it was obtained, t h e period was about 30 sec- onds, and it was l i g h t l y damped. The p i l o t reported t h a t t h e a i r c r a f t was characterized by a heavily damped, mildly divergent motion; t h i s motion appeared t o be divergent because t h e s h o r t period was s o long corrective a c t i o n had t o be taken before a tendency t o r e t u r n could be noted ( s e e fig.9).
Low longitudinal s t a b i l i t y and large pitching-moment changes with power changes have been observed with other STOL a i r c r a f t ( r e f s . 1 and 2 ) and these c h a r a c t e r i s t i c s require moderate p i l o t e f f o r t t o c o r r e c t and maintain f l i g h t path. The p i l o t considered t h e s e c h a r a c t e r i s t i c s u n s a t i s f a c t o r y on t h e TJF-XS with t h e M E o f f and r a t e d them 5-1/2, although no a c t u a l approaches and land- ings were made with t h i s configuration. The corresponding values of s t a b i l i t y and trim change with power and t h e r a t i n g are included i n f i g u r e 7. It i s seen t h a t good c o r r e l a t i o n was obtained with t h e simulator r e s u l t s and t h e p i l o t remarked t h a t it f e l t l i k e t h e corresponding configuration "flown" on t h e simulator.
With t h e ASE on, t h e a i r c r a f t responded as i f it were s t a t i c a l l y s t a b l e ; it returned t o i t s o r i g i n a l t r i m a t t i t u d e i n l e s s than 10 seconds after being disturbed. Although t h e p i l o t considered t h e a i r c r a f t overdaqed, t h e longi- t u d i n a l c h a r a c t e r i s t i c s of t h e UF-XS with t h e ASE on w e r e s a t i s f a c t o r y (PR = 3 - l / 2 ) . This improvement with ASE on was not q u i t e as l a r g e as obtained on t h e simulator. A comparison of response t o an e l e v a t o r pulse i n f l i g h t with M E on and off is shown i n f i g u r e 10. I d e n t i c a l response was obtained i n i - t i a l l y because t h e ASE introduced a l a g of about 0.5 t o 1.0 second andincluded a term t o avoid reducing t h e c o n t r o l effectiveness (see t a b l e 111). This l a g w a s not included i n t h e simulation program.
S t a t i c Lateral-Directional C h a r a c t e r i s t i c s F l i g h t . - The s t e a d y - s t a t e s i d e s l i p data measured i n f l i g h t a r e given i n f i g u r e 11. Similar c h a r a c t e r i s t i c s were used during t h e simulator t e s t s . The s i d e s l i p vane on t h e a i r c r a f t had l i m i t e d motion ( + l 5 ' ) , and t h e values p r e - a t higher angles i n f i g u r e 1 1 w e r e based on t h e assumption t h a t s i d e - sented s l i p developed w a s l i n e a r l y r e l a t e d t o t h e rudder d e f l e c t i o n . Figure 12 shows t h e side-force c o e f f i c i e n t caused by t h e l i k e - r o t a t i o n p r o p e l l e r s . The d a t a (55-30-0) showed higher side-force values than t h e from t h e current t e s t s previous d a t a used f o r t h e simulator tests. A reference l i n e showing t h e r e l a t i o n of side-force c o e f f i c i e n t equivalent t o a 3 ' r i g h t bank angle and zero s i d e s l i p i s included. This l i n e shows t h a t 3 ' r i g h t bank would be required t o t r i m t h e side-force a t zero s i d e s l i p . I n f i g u r e 11, it i s seen t h a t l i t t l e c o n t r o l d e f l e c t i o n i s required t o produce a 3 ' bank angle; however, maintaining 3 ' during t h e approach w a s found t o be uncomfortable and s i d e s l i p - ping was preferred. The indicated steady-state s i d e s l i p angle vas 1 1 ' f o r zero bank angle. It was noted e a r l i e r t h a t t h e indicated s i d e s l i p angle may be 3/2 t r u e s i d e s l i p angle. This f l i g h t condition required considerable rud- der d e f l e c t i o n and g r e a t l y reduced t h e d i r e c t i o n a l c o n t r o l power a v a i l a b l e (see e a r l i e r s e c t i o n ) s o t h a t concern was expressed t h a t a n engine f a i l u r e might be catastrophic. N o engine-out d a t a were obtained i n f l i g h t . An engine failure (at wave-off power) was approximated on t h e simulator, and t h e han- d l i n g q u a l i t i e s were found t o be acceptable f o r VFR, b u t unacceptable under IFR conditions.
Wind-tunnel t e s t s by SMIC indicated t h a t t h e s i d e f o r c e r e s u l t e d l a r g e l y from asymmetric pressures on t h e nacelles. Smaller nacelles, such as used on turboprop configurations would s i g n i f i c a n t l y reduce t h e s i d e force. This is somewhat s u b s t a n t i a t e d by t h e smaller side-force values f o r t h e BLC-130 shown 1 1 i n f i g u r e 12. Side force was not a concern with t h a t a i r c r a f t . A b e t t e r solution, of course, would be t o have opposite r o t a t i o n p r o p e l l e r s l i k e those on t h e a i r c r a f t of reference 1.
Dynamic Lateral-Directional S t a b i l i t y Previous tests with STOL a i r c r a f t have shown t h a t when t h e a i r c r a f t i s banked i n t o a t u r n , a t u r n r a t e i n t h e desired d i r e c t i o n does not develop f o r several seconds, and l a r g e excursions i n s i d e s l i p angle result. For t h e s e STOL a i r c r a f t , t h e d i r e c t i o n a l period was moderately long, damping was low, and t h e r e frequently was cross-coupling; therefore, it w a s d i f f i c u l t t o coor- d i n a t e t h e t u r n with t h e rudder. These c h a r a c t e r i s t i c s have been highlighted i n a s t e p -bank maneuver with rudder f b e d . A t y p i - (See refs. 1, 2, and 4. ) c a l response t o a 15' s t e p bank with rudder f i x e d i s shown i n f i g u r e 13.
Large s i d e s l i p excursions can be noted. The following discussion and tests will p e r t a i n t o t h e f a c t o r s a f f e c t i n g t h e s e c h a r a c t e r i s t i c s .
Simulator.- The e f f e c t of d i f f e r e n t l e v e l s of s t a t i c d i r e c t i o n a l s t a b i l - i t y and d i r e c t i o n a l damping on p i l o t opinion i s shown i n f i g u r e s 14(a) and ( b ) , respectively. A t low values of d i r e c t i o n a l s t a b i l i t y , t h e periods become longer and l a r g e r s i d e s l i p excursions a r e produced by a given c o n t r o l input or by a disturbance; therefore, more of t h e p i l o t ' s a t t e n t i o n i s required t o c o n t r o l s i d e s l i p . With high s t a b i l i t y , t h e a i r c r a f t i s t o o s e n s i - t i v e t o gusts and has i n s u f f i c i e n t damping, when N r i s maintained constant.
A t low values of damping t h e a i r c r a f t i s t o o s e n s i t i v e t o , g u s t s and a t high values it becomes too sluggish. This l a t t e r condition could be somewhat improved with a d d i t i o n a l d i r e c t i o n a l c o n t r o l power. These data are combined i n f i g u r e 1.5 t o show t h e desired l e v e l s of s a t i s f a c t o r y (PR = 3.5) and unsat- i s f a c t o r y ( P R = 6.5) d i r e c t i o n a l s t a b i l i t y and damping f o r STOL a i r c r a f t with approach speeds of 50 t o 60 knots, d i r e c t i o n a l c o n t r o l power of 0.2 rad/sec2, and l i t t l e cross coupling. F l i g h t values f o r t h e NC-l30B, t h e BR 941, and t h e UF-XS without s t a b i l i t y augmentation a r e included and these d a t a i n d i c a t e good c o r r e l a t i o n with t h e simulator r e s u l t s . The UF-XS f l i g h t values a r e discussed i n more d e t a i l i n a l a t e r s e c t i o n . Reference 5 r e p o r t s s i m i l a r boundaries obtained with a v a r i a b l e s t a b i l i t y h e l i c o p t e r a t 45 knots and IFR conditions.
Figure 16(a) shows t h e e f f e c t of roll due t o yaw rate, L, without sta- b i l i t y augmentation. A t high values of L r l a r g e s p i r a l i n s t a b i l i t y r e s u l t e d , and t h e lateral motion was s i m i l a r t o t h a t of an a i r c r a f t with a l a r g e nega- t i v e d i h e d r a l e f f e c t ; t h e bank angle w a s doubled i n about 3 seconds. Fig- ure 1 6 ( b ) shows how t h e handling is a f f e c t e d by d i f f e r e n t values of t h e d i h e d r a l e f f e c t , Lp. Separate curves derived from t h e UF-XS and Breguet 941 simulation are shown. The marked difference w a s due t o a l a r g e r value of L r f o r t h e "basic" UF-XS than f o r t h e "basic" Breguet 941 (as noted on t h e curves). For both cases, near zero d i h e d r a l e f f e c t w a s desired. With zero t h e simulated "basic" UF-XS was l a t e r a l l y unstable, with t h e bank angle Lp doubling i n 10 seconds; t h e r a t i n g was 4. With a p o s i t i v e dihedral e f f e c t poor Dutch r o l l c h a r a c t e r i s t i c s were obtained and t h e d i r e c t i o n a l damping w a s Negative d i h e d r a l e f f e c t increased t h e lateral i n s t a b i l i t y .
reduced.
Varying yaw due t o lateral c o n t r o l (N8a/L& over a l a r g e range had a s u r p r i s i n g l y l i t t l e e f f e c t on t h e p i l o t r a t i n g ; t h a t is, when t h i s r a t i o w a s increased from 0 t o -0.4, t h e r a t i n g changed only by 1 / 2 a u n i t . Similar r e s u l t s w e r e reported i n t h e simulation t e s t s of reference 4.
Since much of t h e d i f f i c u l t y i n t h e l a t e r a l - d i r e c t i o n a l mode w a s r e l a t e d t o a lack of t u r n coordination which caused s i d e s l i p excursions, t e s t s were performed using Ni o r Np t o improve t h e handling q u a l i t i e s . The e f f e c t of t h e s e are shown i n f i g u r e 17. Reference 4 discussed, i n considerable d e t a i l , t h e a b i l i t y of NB t o reduce s i d e s l i p excursions during maneuvering. Refer- ence 6 a l s o discusses recent f l i g h t r e s u l t s with d i f f e r e n t methods of NB augmentation. This reference a l s o shows how Np can be used f o r t u r n coor- dination. Further, it shows t h a t t h e op~imumvalue of Np corresponds t o t h e r a t i o of g/V. Figure 17 shows t h a t Nb o r Np can g r e a t l y inrprove t h e han- d l i n g of these STOL a i r c r a f t . It should be f u r t h e r noted t h a t a i r c r a f t gener- a l l y have zero NP and s m a l l values of J . T $ ; therefore, augmentation equipment w i l l be required f o r producing t h e desired l e v e l s of these terms.
Tests were a l s o performed simulating t h e ASE of t h e UF-XS. The ASE provided a t t i t u d e s t a b i l i z a t i o n and increased damping i n t h e r o l l a x i s , and deflected t h e rudder when t h e ASE deflected t h e a i l e r o n ( t a b l e 111). The p i l o t considered t h e handling g r e a t l y improved, and r a t e d it 2-1/2.
F l i g h t . - The d i r e c t i o n a l s t a b i l i t y and damping c h a r a c t e r i s t i c s of t h e UF-XS with ASE o f f a r e given i n f i g u r e 15. The d i r e c t i o n a l period w a s 6-1/2 seconds and f a i r l y well damped (damping r a t i o of 0 . 3 ) , and t h e p i l o t r a t e d t h e s e c h a r a c t e r i s t i c s 3-l/2. The r a t i n g of 3-1/2 compares w e l l with t h e gen- e r a l i z e d simulator derived boundary. The corresponding dimensional character- i s t i c s a r e an N p of 0.8 and an N r of -0.3. I n c o n t r a s t , t h e steady-state N p of 0.5 which should have produced a period of s i d e s l i p d a t a indicated an 9 seconds (values t h a t corresponded t o t h e "basic" UF-XS simulated). The reason f o r t h i s discrepancy i s not understood, b u t it seems reasonable t h a t a p o r t i o n of it could be caused by erroneous s i d e s l i p measurements due t o t h e it i s seen t h a t flow f i e l d a t t h e m a s t on t h e fuselage.* From f i g u r e 14, e i t h e r of these values of would have r e s u l t e d i n similar p i l o t r a t i n g s N p of t h e s t a b i l i t y and damping.
The s p i r a l i n s t a b i l i t y with t h e ASE off i s seen i n t h e time h i s t o r y i n f i g u r e le; t h e bank angle i s doubled i n about 4 seconds. This high i n s t a b i l - was caused by t h e l a r g e value of r o l l due t o yaw r a t e ( L r of about 0.5) i t y and near zero d i h e d r a l e f f e c t . Because of t h i s i n s t a b i l i t y t h e p i l o t ' s con- s t a n t a t t e n t i o n w a s required t o maintain t h e desired bank angle and heading, and t h i s c h a r a c t e r i s t i c w a s considered u n s a t i s f a c t o r y (PR of 4-1/2). The measured L r was g r e a t e r than used f o r t h e 'lbasic" UF-XS simulated; when t h e c o r r e c t L r w a s used, good c o r r e l a t i o n was obtained between t h e simulator and f l i g h t r e s u l t s ( f i g . 16).
- _ _ . .. - .
_. .
*All o t h e r d e r i v a t i v e s a r e based on t h e assumption t h a t s i d e s l i p w a s a c c u r a t e l y measured.
The p i l o t r a t e d t h e o v e r - a l l l a t e r a l - d i r e c t i o n a l c h a r a c t e r i s t i c s of t h e UF-XS i n f l i g h t with ASE off as 5 -1/2 t o 6; whereas t h e r a t i n g was 4 for t h e b a s i c UF-XS simulated. The c h a r a c t e r i s t i c s include t h e lateral i n s t a b i l i t y , ease of making a coordinated turn, and t h e response t o c o n t r o l inputs. Closer agreement i s obtained when t h e derivatives used f o r t h e "basicfr UF-XS i n t h e simulator a r e corrected t o those measured i n f l i g h t . Even then a difference of about 1 r a t i n g value remains. The reason f o r t h i s discrepancy i s not c l e a r ; however, t h e r e a r e s e v e r a l contributing f a c t o r s . The primary objection reported by t h e p i l o t was "high adverse yaw," which t o him appeared t o be higher f o r t h e a i r c r a f t than f o r t h e simulated vehicle. This reported "adverse yaw" includes not only adverse yaw due t o a i l e r o n s (NGa/L?ja), but a l s o t h e terms t h a t cause s i d e s l i p (because of low s t a b i l i t y or cross coupling of aerodynamic or i n e r t i a terms), making it d i f f i c u l t t o reduce t h e s i d e s l i p excursions i n a t u r n entry. I n f l i g h t , t h e adverse yaw due t o a i l e r o n s w a s smaller and the d i r e c t i o n a l s t a b i l i t y was g r e a t e r than t h e values used f o r t h e "basic" simulated vehicle, s o these f a c t o r s did not contribute t o t h e discrepancy. However, t h e simulator lacked yaw motion, and t h e product o f i n e r t i a w a s not included i n t h e equations of motion. I n f l i g h t , t h e p i l o t did not have a s i d e s l i p indicator, and t h e t r i m s i d e s l i p w a s g r e a t e r than t h a t on t h e simulator because of t h e l a r g e r p r o p e l l e r s i d e force. Further, t h i s l a r g e trim s i d e s l i p angle of 1 1 ' made it d i f f i c u l t t o have a good reference during maneuvering so t h a t the t u r n coordination could be evaluated.
With t h e ASE on, t h e a i l e r o n s a r e deflected t o maintain wings l e v e l and t o increase t h e r o l l damping; the rudder i s a l s o deflected proportional t o t h e a i l e r o n commanded by t h e ASE ( s e e t a b l e 111). Figure 19 compares the response t o an a i l e r o n step with ASE on and o f f . With ASE on, the i n i t i a l response i s similar t o t h a t with ASE off because of a 1-second l a g i n the augmentation system and a term included t o avoid reducing t h e control e f f e c - tiveness (see t a b l e 111). After t h i s 1 second, t h e response i s reduced because of the s t a b i l i z i n g and damping input. A constant a i l e r o n defection i s
seen t o occur from time 1 second t o about 4 seconds because of t h e a u t h o r i t y
l i m i t i n the system. (The absolute value of yaw r a t e p l o t t e d i n f i g u r e 19 f o r ASE on may be i n c o r r e c t . ) The ASE eliminated t h e bad s p i r a l i n s t a b i l i t y (Fig. 18), and t h e a i r c r a f t w a s e a s i e r t o f l y and maintain on a given patn.
However, t h e d i r e c t i o n a l period w a s increased t o 8 seconds, t h e damping r a t i o decreased t o 0.2, and t h e s i d e s l i p excursions during maneuvering was not elim- inated. The l a t e r a l - d i r e c t i o n a l c h a r a c t e r i s t i c s with t h e ASE on were con- sidered t o be s a t i s f a c t o r y (PR = 3-1/2). However, t h e high a t t i t u d e s t a b i l i t y and s i d e s l i p excursions would be more objectionable on an STOL land plane where more extensive maneuvering would be required.
Operational Characteristics and P i l o t s ' Comments All take-offs and landings were made i n VFR conditions with ASE on; one f l i g h t was made from moderately rough water with 18 t o 20 knots of wind.
Acceleration during t h e take-off i s low b u t steady, with l i t t l e o r no spray s t r i k i n g the p r o p e l l e r s . Control of p i t c h a t t i t u d e i s easy and l i f t - o f f i s accomplished i n a s l i g h t nose-up a t t i t u d e between 45 and 50 knots (depending upon gross weight). Immediately a f t e r l i f t - o f f , t h e nose swings noticeably t o t h e r i g h t as t h e a i r c r a f t assumes i t s trim s i d e s l i p angle f o r a l e v e l wing a t t i t u d e . O n t h e simulator, t h i s asymmetry w a s counteracted by banking t h e a i r c r a f t s l i g h t l y t o t h e r i g h t . I n f l i g h t , however, t h i s proved inrpractical s i n c e t h e s i d e f o r c e due t o l i k e - r o t a t i o n p r o p e l l e r s w a s about twice t h e value simulated and t h e r e s u l t a n t s i d e force f e l t uncomfortable t o t h e p i l o t s . With ASE o f f , t h e combination of n e u t r a l l o n g i t u d i n a l s t a b i l i t y , s p i r a l i n s t a b i l i t y , and l a r g e s i d e s l i p excursions i n t u r n e n t r i e s rendered t h e a i r c r a f t handling q u a l i t i e s unacceptable f o r a l l b u t emergency operation. Although no a c t u a l approaches were attempted with ASE o f f , it was f e l t t h a t under v i s u a l condi- t i o n s , it would be p o s s i b l e t o land t h e a i r c r a f t with ASE o f f . Operation under instrument conditions with ASE off, however, would be quite hazardous.
The landing approaches were made i n a f l a t a t t i t u d e a t a r a t e of descent of about 500 ft/min. Sink rate was adjusted with power while p i t c h a t t i t u d e was held r e l a t i v e l y constant t o maintain t h e desired 55-knot approach speed.
A t about 200 t o 300 f e e t , power was adjusted t o reduce t h e rate of descent t o about 300 ft/min and t h e . a i r s p e e d t o about 50 knots.
J u s t before t h e a i r p l a n e contacted t h e water, i t s nose was r a i s e d s l i g h t l y t o t h e landing a t t i t u d e . A s it entered t h e "ground e f f e c t " i t s nose swung abruptly t o t h e l e f t as t h e air- c r a f t a l i n e d i t s e l f i n t h e d i r e c t i o n of f l i g h t , without any a c t i o n being taken by t h e p i l o t . This w a s not considered objectionable f o r t h e seaplane b u t would probably c r e a t e a problem f o r an a i r c r a f t landing on a runway. The s l i g h t f l a r e and favorable ground e f f e c t reduce t h e sink r a t e somewhat before touchdown. This, i n combination with t h e deep h u l l design,provided a s o f t l i t t l e tendency t o bounce.
landing with The ASE does a reasonable job of s t a b i l i z i n g t h e a i r p l a n e and makes t h e approach and landing t a s k r e l a t i v e l y easy. The objectionable f e a t u r e s of t h e ASE were t h a t a lateral c o n t r o l f o r c e had t o be maintained during steady-state t u r n s (ASE attempting t o r e t u r n t h e a i r p l a n e t o wings l e v e l ) and t h a t a s l i g h t longitudinal o s c i l l a t i o n occurred during p i t c h maneuvers (due t o l a g i n t h e ASE). This did not compromise t h e seaplane approach and landing since very l i t t l e maneuvering i s required. This type of s t a b i l i t y system would be much more objectionable i n a land plane, however, where m a n y more c o n s t r a i n t s a r e placed on t h e landing a r e a .
CONCLUDING REMARKS The following remarks p e r t a i n t o t h e seaplane c h a r a c t e r i s t i c s i n t h e 50- t o 6 0 - h o t speed regime.
The a i r c r a f t could e a s i l y be operated from t h e water a t take-off and landing speeds of 50 knots, which corresponded t o a l i f t c o e f f i c i e n t of about 4. With t h e Automatic S t a b i l i z a t i o n Equipment ( M E ) engaged t h e handling c h a r a c t e r i s t i c s were s a t i s f a c t o r y f o r t h e seaplane mission. The ASE stabilized t h e roll and p i t c h a t t i t u d e s and increased t h e rate damping about t h e s e axes.
Without t h e ASE, t h e following d e f i c i e n c i e s r e s u l t e d i n an u n s a t i s f a c t o r y a i r c r a f t ; low s t a t i c l o n g i t u d i n a l s t a b i l i t y , a very unstable s p i r a l mode, and l a r g e s i d e s l i p excursions during t u r n e n t r i e s . S a t i s f a c t o r y response t o c o n t r o l inputs w a s obtained about t h e roll and p i t c h axes, b u t t h e l i k e - r o t a t i o n p r o p e l l e r s reduced t h e d i r e c t i o n a l c o n t r o l t o an unsatisfactory l e v e l .
The simulator t e s t s performed p r i o r t o f l i g h t tests w e r e u s e f u l i n pro- viding a preliminary evaluation, showing p o t e n t i a l problem areas and some causes and s o l u t i o n s t o these problem areas. Good c o r r e l a t i o n was obtained between t h e simulator and f l i g h t results with t h e exception t h a t i n maneuvers t h e s i d e s l i p excursions were l a r g e r i n f l i g h t than on t h e simulator.
Ames Research Center National Aeronautics and Space Administration Moffett Field, C a l i f . , June 7, 1965 REFERENCES 1. Quigley, Hervey C.; Innis, Robert C.; and Holzhauser, C u r t A . : A F l i g h t I n v e s t i g a t i o n of t h e Performance, Handling Qualities, and Operational C h a r a c t e r i s t i c s of a Deflected Slipstream STOL Transport Airplane Having Four Interconnected Propellers. N A S A TN D-2231, 1964.
2. Quigley, Hervey C.; and Innis, Robert C . : Handling Q u a l i t i e s and Opera- t i o n a l Problems of a Large Four-Propeller STOL Transport Airplane.
N A S A TN D-1647, 1963.
3. Vagianos, LCDR Nicholas J., USN; and Rooney, Eugene C . : F i n a l Report F l i g h t T e s t Evaluation of t h e UF-XS Japanese STOL SeaplaEe.
Tech. Rep.
FT 212-031R-64, Naval Air Test Center.
4. Quigley, Hervey C. : and Lawson, Herbert F., Jr. : Simulator Study of t h e a Large Four-Propellered STOL Lateral-Directional Handling Q u a l i t i e s of Transport Airplane. N A S A TN D-1773, 1963.
5 . Garren, John F., Jr.; Kelly, James R.; and Reeder, John P. : Effects of Gross Changes i n S t a t i c Directional S t a b i l i t y on V/STOL Handling Char- a c t e r i s t i c s Based on a F l i g h t Investigation. N A S A TN D-2477, 1964.
6. Anderson, Seth B.; Quigley, Hervey C.; and Innis, Robert C.: S t a b i l i t y and Control Considerations f o r STOL A i r c r a f t . Presented a t t h e Advisory Group f o r Aeronautical Research and Development Panel Meeting on F l i g h t Mechanics, P a r i s , France, June 9-11, 1965.
TABLE I . . GEOMETRIC DATA O F AIRJ?LA.NE
. .
~ . . .
Wing . . . . . . . . . . . . . . . . . . . . . . . . . .
Total area, s q f t 835
Span. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 c
Mean aerodynamic chord . . . . . . . . . . . . . . . . . . . . 10 ft. 9 i n .
Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
0.5
A s p e c t r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.65
Incidence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.0c
A i r f o i l s e c t i o n
ROO% . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA 23017
Tip . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA23012
Dihedral (lower surface) . . . . . . . . . . . . . . . . . . . . . . 2010'
Flap Span (percent wing span)
Inboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Outboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
Chord (percent wing chord) . . . . . . . . . . . . . . . . . . . . . . 25
Deflection (maximum)
Inboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 800
Outboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60°
Aileron
span (percent wing span) . . . . . . . . . . . . . . . . . . . . . . . 28
Chord (percent wing chord) . . . . . . . . . . . . . . . . . . . . . . 25
Horizontal t a i l
Area. s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200
Span. f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.5
A i r f o i l section . . . . . . . . . . . . . . . . . . . . Inverted NACA 2414
Elevator area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . . 60
Vertical t a i l Area. s q f t . . . . . . . . . . . . . . . . . . .
Span. f t . . . . . . . . . . . . . . . . . . . . .
A i r f o i l section NACA 0014
Rudder area. sq f t . . . . . . . . . . . . . . . . . . . . . . . 41.2
. . . . . . . .
TABU 11.- GEOMETRIC CWLSACTERISTICS OF CONTROLS IN TAKE-OFF AND LANDING CONFIGURATION Maximum c o n t r o l Maximum surface d e f l e c t i o n d e f l e c t i o n
-
8.2 i n . a f t Elevator - 4 0 ° (up)
Longitudinal 4.0 i n . forward Elevator +22O (down) Lateral* 1 0 5 ' r i g h t Right a i l e r o n 25' up Right outboard f l a p 5' down
Right s p o i l e r 5 7 ' up
L e f t a i l e r o n 1 8 ' down
L e f t outboard f l a p 4 5 ' down
Left s p o i l e r 0 ' 1 0 0 ' l e f t Vice versa Directional 2.4 i n . r i g h t Right rudder 36'
2.5 i n . Left rudder 4 4 '
l e f t *The r i g h t a i l e r o n d e f l e c t i o n i s used throughout t h e r e p o r t as a reference for t h e lateral c o n t r o l input; t h e s p o i l e r does not start deflecting u n t i l about 50-percent con- t r o l input.
1 8 TABU3 111.- EQUATIONS FOR AUTOMATIC STABILIZATION EQUIPMENT (ASE) '~ASE = '~ASE ',total = '~ASE + 'rpilot are provided s o t h a t the ASE where 0, is 6 ' and &8epilot and Ka6 a p i l o t does not reduce t h e apparent control effectiveness. These equations and con- s t a n t s were provided by SMIC. The constants, applied t o the a t t i t u d e and rate portions of the equation, approximate the f l i g h t system with i t s 1-second time lag. The values for Ke and K a were not given b u t t h e i r magnitude w a s between 0.5 and 1 . 5 . The a u t h o r i t y of t h e ASE w a s 20 percent of the m a x i " surface deflection.
TABLE 1 V . - VALUES OF DERIVATIVES USED I N SIMULATION AND MEASURED I N F L I G H T -~ .
F l i g h t
Der ivat ivc U F -xs
Breguet 941
UF -xs
V a s i c V a r i a t i o n "Basic" ASE o f f - 0.25 0 . 2 5 t o - 0 . 2 5 -0.09 0 . 1 5
-0 - 47
-0.22 t o -0.91 -0.41 -0.1 --- -0.40 -0.43 -0.4
---
-0.67 -1.02 -0.7 0.61 1.05 --- 0 . 5 5 -0.28 -0 * 7 5 -0.21 --- 0.46 0.54 0 . 8 -0.30 .Os2 t o -0.4 -0.33 -0.3 -0.16 0 . 5 t o -0.5 - 0 . 0 5 -0.1 --- 0 --- 1.07 nose r i g ?
---
t 0 . 2 0 k0.19 ) . 2 7 nose left 0.2 to -0.2 -0.32 - 0 . 0 3 ~ 0.25 0 . 2 5 t o 0.60 0.14 0 . 5
---
-0.67 -0.82 -0.8
---
Io. 50 k0.42 +o. 5 -0.20 0 t o -0.4 -0.01 -0.06 --- -0.10 0.23 -0.2 -0.13 0.08 t o -0.13 -0.10 -0.14l --- 0.03 0 . 0 5 --- --- -0.008 -0.018 ;1 knots 30 knots 55 knots 4.0 3.7 3.7
916 1 1 9 6
-64 , 000
2 2 8 , 0 0 0 1 8 4 , 0 0 0 '
1-36 , 000
.557200 173, 5o02
194,000 +17 , coo
329,000' 12,300 13,8002 10- 7 5 12.15 1 0 . 7 5 80.0 76.1 80.0 ~ ~ .. .- _.:. .
- - 'Based on i n d i c a t e d s i d e s l i p reading, see t e x t .
'Estimated from v a l u e s p r o v i d e d a t lower g r o s s weights.
TABLE V.- PILOT OPINION RATING SYSTEM FOR UNIVERSAL USE Primary Numerical Ad j ec tive Can be Description mission rating rating landed accomplished ~~~ ~ 1 Excellent, includes optimum Yes Yes 2 Good, pleasant to fly Yes Yes Satisfactory, but with some mildly unpleasant characteristics Yes Yes Acceptable, but with unpleasant
' r t
characteristics Yes Yes Eknergency Unsatisfactory Unacceptable for normal operation Doubtful Yes operation Acceptable for emergency condition o n l y ' Doubtful Yes Unacceptable even for emergency Doubtful No condition' Unacceptable
operation 8 Unacceptable - dangerous
No
Unacceptable - uncontrollable No
'Failure of a stability augmenter.
I l l I I I I l l I I l l 1 I
6- a - :n.
A-33534 Iu w Figure 1.- The UF-XS in STOL configuration.
*
t 80.0' (a) Top and front v i e w Figure 2. - Sketch of airplane.
73.5' Spray suppressor ( b ) Side view and hull cross s e c t i o n Figure 2. - Concluded.
d A-34701 Figure 3 . - Diagram of Ames Moving Base Transport Simulator.
c .
O/O T.O. Power
- 100 (4050 hp)I
I I
/
E 0
u- L n: \ / -400 u.
-\ Stall
.- 34-
-800
- I200
40 44 48 52 56 60 64 68 72 76
Vi F, knots
Figure 4.- Operating envelope i n STOL configuration a t 5,000 feet; 6f = 55/30/0, BLC on.
- 1 3 u
I
I I I
I
( I (
i I
-0.4 0.8 0 0.4 0 0 I 2 Tcl c D Figure 5.- Approximate lift and drag characteristics in STOL configuration at 5,000 feet; 6f = 55/30/0, BLC on.
If! 20
-
Q)
I
\ R d -10 -8 -6 -4 -2 0 2 4 6
PUl I Column center travel , in- Push
(b)Longitudinal -P 20 ,.
Q)
E
9 0
-20 -40 -3 -2 - I 0 I 2 3
L e f t Pedal travel, i n . Right
(c) Directional Figure 6.- Force characteristics of control system.
. ..
4'
0 Flight -0. I
I / = Simulator -0.5
[L a
I
0.4 0.2 0 -0.2 - 0 . 4 -0.6 M a I I/sec2 I I I I 0 0.4 0.6 0.8
wn , I/sec
( a ) Angle of attack stability Ma
0 Flight 0.15 1
, 4 5 7 Simulator 0.25
I
Figure 7.- Effect of l o n g i t u d i n a l s t a b i l i t y on p i l o t r a t i n g ; ASE o f f .
Max . . .
- I O o / o TOO. Power
- 62
- 34
I O '/////7 Max 44 4% 52 56 60 64 68 72 76
V i F , knots
Figure 8.- Variation of elevator angle w i t h speed of constant power.
Q) TI ..
Q) do IO
s
g
7 Flight
---- Computed with M,=O
-
- 1 1 1 I -0.2
t , sec
Figure 9.- Response t o elevator step; ASE o f f .
I P u , W 9 IO 0 I 2 3 4 5 6 t , sec Figure 10.- Comparison of response t o s t i c k rap w i t h ASE on and ASE off.
P t % 30 -I IO a l T1 ..
k o
-10 + 2 - 2 0 E -30 -40 Left Pi 9 deg Right Figure 11.- Steady-state sideslip characteristics.
O 53-40-0
} U F - X S
0 55-30-0
0 70-30 BLC-130 ref. 2
CL -0.4 -0.3 -0.2 - 0 . 1 0 Figure 12.- Side-force c h a r a c t e r i s t i c s with l i k e - r o t a t i o n propellers.
111 11111111 I I I I I I1 IT Q) -0 -10
&
-20 - -
-
0.1 Q) In \ v) c .- P O -0
-
L -0.1 - IO a
u o
-10 0 2 4 6 8 I O t , s e c Figure 13.- Time history of typical lateral-directional oscillation following lateral control input for STOL aircraft; V - 60 knots.
LL Q 0 0.4 0.8 1.2 I .6 2.0
NP , I/sec2
( a ) Directional stability, Nr=-0.3/sec
I
Sensitive Sluggish to gusts E a 0 -0.2 -0.4 -0.6 -0.8 -1.0 N r , I / s e c ( b ) Directional damping, NP= 0 . 5 / s e c 2 Figure 14.- E f f e c t of d i r e c t i o n a l s t a b i l i t y and damping; V = 50 t o 60 knots 9 N€+'rmx = 0.2 rad/sec2, Lp = 0 t o -0.3/sec2, Y p = -O.l/sec, ~p = -O.l/sec.
I U F - X S
1 BR941 (PR= 3i)
- 0 . 4 NC- 130B (PR= 4) 0 0.2 0.4 0.6 0.8 I .o 1.2
NB , I /sec2
Figure 1 5 . - Directional stability and damping boundaries for
unaugmented aircraft with low cross coupling; v = 50 to
60 knots, NgrGrmax = 0.2 rad/sec2, Lp = 0 to -0.3/sec2, Y p = -~.l/sec, ~p = -O.l/sec.
LP
a Flight -0.03
Simulator 0 to -0.3 U LL -
I spiral instability
I I
.- - - . .
0 0.4 0.8 I .2 I .6 2.0 2.4
Lr , I /sec
( a 1 Roll due to yaw rate L r , U F - X S 0.25 -
/ BR941 “‘I4\
a Q
rol I -7-7
spiral instability and reduced damping
-
0.8 0.4 0 -0.4 -0.8 -1.2 - I .6 Lp , I /sec2 ( b ) Dihedral e f f e c t Figure 16. - E f f e c t of l a t e r a l parameters on p i l o t opinion; ASE o f f , V = 50 t o 60 knots, Np = 0.5/sec2, N , = 0.3/sec.
S l u g g i s h = ~ 0 0 . 4 0.8 1.2 1.6 2.0 NB, I/sec ( a ) Damping due to sideslip rate
-0.4 - 0.2 0 0.2 0 . 4 0 . 6
N , I h e c
P ( b ) Yaw due to roll rate Figure 17.- Augmenta%ion to improve turn coordination; V = 50 to 60 ~ 6 ~ 6 ~ ~ = 0.2 rad/sec2, knots, ~p = 0.5/sec2, N~ = -0.3/sec 2 ~p = o to -0.3/sec2, ~p = -O.l/sec, ~p = -O.l/sec.
1 6 I 8- 2 4 6 a IO 12 1 4 t ,sec Figure 18.- F l i g h t time h i s t o r y of bank angle change with controls trimmed f o r 3 ’ bank.
I ASE off -e- IO
- 20
CT a , U I -10 m -20 -30 0.3 Q L m l . .
Q.
0 0 0.2 U .-
-0 CT a , - -20 O q 1 c-r e---
-
L 0’ L m , /’ -----.---,, --#/4 -400 -~ -0.1 2 3 4 5 0 I 2 3 4 5 I t , sec t, sec Figure 19.- Comparison of response due t o a i l e r o n s t e p with ASE on and o f f , i n f l i g h t .
NASA-Langley, 1965 A -2082
... . . .. . . . - - . . .. . . .._-. .. ... . - **
“The aeronautical aiid space activities of the United States shall be so as t o contribute . . . to the expaizsion of human h o w l - conducted edge of phenomena in the atmosphere and space. The Administration shall provide for the widest practicable and appropriate dissemination of information concerning its activities and the reszilts thereof .” -NATIONAL AERONAUTICS AND SPACE ACT OF 1958
NASA SCIENTIFIC A N D TECHNICAL PUBLICATIONS
TECHNICAL REPORTS: Scientific and technical information considered important, complete, and a lasting contribution to existing knowledge.
TECHNICAL NOTES: Information less broad in scope but nevertheless of importance as a contribution to existing knowledge.
TECHNICAL MEMORANDUMS: Information receiving limited distri- bution’because of preliminary data, security classification, or other reasons.
CONTRACTOR REPORTS: Technical information generated in con- nection with a NASA contract or grant and released under NASA auspices.
TECHNICAL TRANSLATIONS: Information published in a foreign language considered to merit NASA distribution in English.
TECHNICAL REPRINTS: Information derived from NASA activities and initially published in the form of journal articles.
SPECIAL PUBLICATIONS: Information derived from or of value to NASA activities but not necessarily reporting the results .of individual NASA-programmed scientific efforts. Publications include conference proceedings, monographs, data compilations, handbooks, sourcebooks, and special bibliographies.
Details on the availability of these publications may be obtained from: SCIENTIFIC AND TECHNICAL INFORMATION DIVISION
N AT I 0 N A L A E R 0 N A U T I CS A N D SPACE A DM I N I STRATI 0 N
Washington, D.C. PO546