Economic evaluation of flying-qualities design criteria for a transport configured with relaxed static stability

NASA Technical Paper 1760

LOAn C W Y A W L TECH KIRTLAND I= W r: m JI .r = I E .

Economic Evaluation of

j '#,' , Flying-Qualities Design Criteria . :

for . a .,Transport Configured

With. Relaxed StaticStability

Steven M . Sliwa DECEMBER 1980 -~ - TECH LIBRARY KAFB, NM

NASA Technical Paper 1760

EconomicEvaluation of

Flying-QualitiesDesignCriteria

for a TransportConfigured

With Relaxed Static Stability

Steven M. Sliwa Larzgle~t Research Center Hampton, Virgim'a National Aeronautics and Space Administration Scientific and Technical Information Branch I SUMMARY An optimal design program with constrained parameter optimization has been shown to be useful in evaluating the impact of certain flying-qualities design assumptions and in determining the sensitivity to several related parameter variations. Transports optimally configured with relaxed static stability showed a potential savings in direct operating cost of 1 . 4 percent when compared with transports with conventional static margins. This corresponded to a fuel savings of 4 . 2 percent for the medium-range mission considered. Savings of nearly 1 percent in direct operating cost were also possible from utilizing half the nominal center-of-gravity range of travel and from allowing the landing gear to be structurally dislocated from the wing. Requiring transports to be able to take off with the stabiliz.er trimmed in the most adverse position was shown to penalize the aircraft over 4 percent in direct operating cost.

During the course of this study, it became obvious that there is a need for developing design criteria for the minimum flying qualities that are necessary for specifying the inherent stability and control characteristics of augmented transports. Most existing criteria did not have useful parameters for defining handling qualities of inherently unstable transports which rely upon augmentation systems. Furthermore, few flying-qualities data were avail- able in terms of factors that would be useful for developing appropriate inherent longitudinal-handling-qualities design criteria for transports con- figured to take maximum advantage of relaxed-static-stability augmentation systems.

INTRODUCTION Active controls technology in aircraft design involves the application of automatic control systems which augment either the rigid o r flexible body dynamics of the aircraft. This is done to enhance either performance, struc- tural efficiency, airframe lifetime, ride quality, or some other measure.

Relaxed-static-stability augmentation systems (RSSAS) constitute an active controls concept that has already been successfully applied to fighter and supersonic transport configurations. Similarly, significant benefits are anticipated for subsonic commercial transports if the RSSAS concept could be applied (refs. 1 and 2 ) . Utilization of RSSAS permits a more aft center-of- gravity position, which typically attenuates the required tail lift for trim, thereby reducing the induced drag of the tail. Stability reductions also per- mit smaller tail surfaces, which reduce wetted area drag and weight.

The benefits of applying RSSAS to transports were initially hypothesized by conceptually retrofitting current configurations (refs. 3 and 4 ) . The per- formance gains were estimated by adjusting center of gravity and tail size to minimize weight and tail drag while satisfying flying-qualities and control- power requirements. This scheme provided rough approximations of the potential benefits, but fell short of documenting the full benefits which would be possible by applying RSSAS early in the design process. Currently, a program is underway to study the implications of minor configuration alterations at th preliminary design level which could enhance the application of active contro technology (ref. 5).

With respect to RSSAS, there are two major obstacles to hinder the reali- zation of the maximum performance improvements. First, syntheses of the config- urations under study are being heavily influenced by current hardware and, in reality, have only minor degrees of freedom in geometry. The designs are being optimized by classical engineering methods, which include intuition in achieving the proper balance between weight savings and performance improvements. It is assumed that the operating cost of the overall vehicle will then be optimized.

Secondly, the first assumption required when designing a transport with an RSSAS is the level of the aerodynamic or inherent stability contribution toward the fully augmented flying qualities. If it can be assumed that it is always possible to augment an airplane to the desired level of flying qualit the unaugmented flying qualities impact the design principally through failure mode considerations. It is still unresolved, even philosophically, what level of flying qualities a transport should have in the event of control system fa ures (refs. 6 to 11 ). Design philosophies range from requiring excellent fly- ing qualities to having marginally safe handling qualities for landing and eve to allowing loss of the aircraft (requiring fail-safe reliability in the auto- matic control system).

The study reported herein utilized a direct constrained optimization pro- cedure for the preliminary optimal design of transport aircraft for the purpos of identifying the full benefits of RSSAS. The aircraft geometry was optimally sized to yield the maximum obtainable improvements from RSSAS in terms of mini- mum direct operating cost per block hour. The flying-qualities and related con- straints were systematically varied to identify the configuration sensitivity to these assumptions. This information (which was presented in condensed form in ref. 1 2 ) should allow both designers and those concerned with flight safety to appreciate the impact of choosing appropria-te unaugmented longitudinal- flying-qualities criteria upon the design of transports configured with R S S A S .

SYMBOLS AND ABBREVIATIONS AR aspect ratio CAS airplane cost, 1 9 7 9 dollars D drag coefficient, total drag coefficient at zero lift ‘D,o 1 979 dollars fuel cost per block hour, CFS c.g. center of gravity L lift coefficient, CL

gs

design lift coefficient of the airfoil section representing the center CL, 0 of the drag bucket maintenance cost per block hour, 1979 dollars CMS D drag, N Doc direct operating cost per block hour, 1979 dollars acceleration due to gravity, 9.8 m/sec2 g FARE income per seat-kilometer required to generate a 15-percent ROI, 1979 dollars IOC indirect operating cost per block hour, 1979 dollars L lift, N aerodynamic efficiency, CL/CD L/D fuselage length, m MAC mean aerodynamic chord, m MDOC modified direct operating cost per block hour, 1979 dollars modified direct operating cost per block hour for baseline config- uration, 1979 dollars steady-state normal acceleration change per unit change in angle of attack for an incremental longitudinal control deflection at constant airspeed, gravity units/radian OPDOT computer program, Optimal Preliminary Design of a Transport P savings in augmented direct operating cost, percent PR pilot rating free-stream dynamic pressure, N/m2 q rms root mean square with respect to mean a n n u a lr e t u r n oninvestment,percent relaxed-static-stability a u g n e n t a t i o n s y s t e m s l i f t i n g surface area, r n 2 ( ~ 2 ) i n s t a l l e d t h r u s t , N ( l b f )

weight , N

l o n g i t u d i n a l l a n d i n g g e a r p o s i t i o n , f r a c t i o n of t h e mean aerodynamic chord s h o r t - p e r i o d damping ratio s h o r t - p e r i o dn a t u r a lf r e q u e n c y , sec-l Subscripts: e m P t Y max maximum t h o r i z o n t a l t a i l t o take-off tot total W wing PROCEDURE Method of C a l c u l a t i o n The cunputerprogram used f o rp e r f o r m i n gt h e trade s t u d i e s d u r i n gt h e f l y i n g - q u a l i t i e s a n a l y s i s was OPDOT, OptimalPreliminaryDesignof a Trans- p o r t . A mre completedescriptionofthiscomputerprogram is p r e s e n t e di n r e f e r e n c e 13. The o p t i m i z a t i o ni n d i c a t e di nr e f e r e n c e 13 was performedusing a m o d i f i c a t i o no ft h es e q u e n t i a ls i m p l e xo p t i m i z e r proposed i n r e f e r e n c e s 1 4 and 15. The nonlinearprogramminglogic is shown i nf i g u r e 1 . A t r i g o n a n e t r i c f u n c t i o nt r a n s f o r m a t i o n( r e f . 1 6 ) was u t i l i z e dw h i c ha u t a n a t i c a l l y scaled t h e i n d e p e n d e n td e s i g nv a r i a b l e s iterated by t h eo p t i m i z e r and applied c o n s t r a i n t s d i r e c t l y t o t h ed e s i g n variables. Naninalvalues for theindependentdesign v a r i a b l e s( w i n g area, wing a s p e c t ratio, f u s e l a g el e n g t h ,h o r i z o n t a l - t a i l area, h o r i z o n t a l - t a i l aspect ratio, i n s t a l l e dt h r u s t ,a n dc . g .l o c a t i o n ) were assuned, and a set of d e s i g nc o n s t a n t s were i n p u ti n t ot h e data base (table I ) . These c o n s t a n t s were used to s p e c i f yt h em i s s i o n ,o p e r a t i n ge c o n c m i c s ,n o n v a r y i n g or simplyscaledgeometries,and some of thenonlinearaerodynamic terms. These i n p u t s were h e l d c o n s t a n t t h r o u g h o u t t h i s a n a l y s i s u n l e s s e x p l i c i t l y s t a t e d o t h e r w i s e .

The o p t i m i z e r is a s e c t i o n of computer code w h i c hi n t e r a c t sw i t ht h ed a t a base, t a k i n gt h ec u r r e n tv a l u e of e a c hd e s i g nv a r i a b l e as i n p u t t o g e n e r a t e a performance index. The performance index used as a figure-of-merit for t h i s s t u d y was a m o d i f i e dd i r e c to p e r a t i n g cost per blockhour (MDOC). T h i s crite- rion,which was minimized by t h eo p t i m i z e rw i t h i nt h ec o n s t r a i n tb o u n d a r i e s , i n v o l v e st h ee s t i m a t i o n of t h e cost performance from a simulatedmission. A schematic for t h el o g i c a lf l o w of t h i s s e c t i o n of computercode is shown i n f i g u r e 2.

The mission profile was a m u l t i p l e - s t e p classical (ref. 1 7 ) approximation to an optimal f u e l - e f f i c i e n tf l i g h tp a t h .T h i sp a t hb e g a nw i t h a climb to a l t i - t u d e w i t h i n t h e maximum s p e e dr e g u l a t i o n s ,f o l l o w e d by a cruise-climb a t maxi- mum cL/CD3I2 to maximize range factor, then a cruise-climb a t maximum cL/cD, and f i n a l l y a rapid d e s c e n t to l a n d i n g . The f u e lu s a g e of t h i s profile hasbeen shown to be w i t h i na b o u t 3 p e r c e n t of a c o n t i n u o u s o p t i m a l l y f u e l - e f f i c i e n t f l i g h tp a t h (ref. 1 8 ) . Although t h i s was n o ti n c l u d e di nt h e cost r e l a t i o n - s h i p s , t h e a i r c r a f t was s i z e d to c a r r y enough f u e l to s a t i s f y t h e r e s e r v e r e q u i r e m e n t s .S i n c ea b o u t 95 p e r c e n to ft h ef u e lb u r n o f f is r e a l i z e dd u r i n g t h ec r u i s e - c l i m bp a r t s of t h em i s s i o np r o f i l e ,t h ei n d e p e n d e n td e s i g nv a r i a b l e s a n do t h e rd e s i g ni n p u t s w i l l havethe most impact o nt h e s ep o r t i o n s of OPDOT's model o ft h et r a n s p o r to p e r a t i o n .

The a i r c r a f t w e i g h t was e s t i m a t e d i n a n i t e r a t i v e f a s h i o n from theequa- t i o n so fr e f e r e n c e s 1 7 , 19, and 20. Although take-off weight is t h e summation of each of theestimatedcomponentweights as well as b o t ht h ep a y l o a da n df u e l , thetake-offweight was r e q u i r e d by many o f t h e s t a t i s t i c a l r e l a t i o n su s e d t o estimate eachcomponentweight.Eachiteration of t h eo p t i m i z e r was s t a r t e d withthetake-offweight from t h ep r e v i o u s set o fi n d e p e n d e n td e s i g nv a r i a b l e s .

A n e n t i r em i s s i o n was simulated, i n c l u d i n gt h er e s e r v es e g m e n t for eachweight i t e r a t i o n . If t h ed i f f e r e n c eb e t w e e nt h e l a s t estimate of grosstake-offweight a n dt h ec a l c u l a t e dw e i g h t was g r e a t e rt h a n 0 . 2 2 n e w t o n ,a n o t h e ri t e r a t i o n was begun withanupdated estimate. The programaveragedaboutfourweight itera- t i o n sp e rp e r f o r m a n c ef u n c t i o n c a l l for an e n t i r e o p t i m i z a t i o n .

The a i r c r a f t w a s trimmed for c r u i s i n gf l i g h tu s i n g a n o n l i n e a r ,i t e r a t i v e method. The aerodynamic forces and moments were e s t i m a t e du s i n g classical aero- n a u t i c s and s t a t i s t i c a l l yn o r m a l i z e dd a t a for supercritical aerodynamics. The drag was estimated u s i n gr e f e r e n c e s 1 7 and 21 to 23. The wing was assumed to be s u p e r c r i t i c a l ,a n dt h ep i t c h i n g moment anddrag were estimated as a f u n c t i o n of wing t h i c k n e s s ratio, Reynoldsnumber, Mach number, and sweep u s i n gt h et e c h - niques from r e f e r e n c e s 1 7 and 23 to 27.

Once t h e a i r c r a f t w a s trimmed, a l l t h ec o n t r i b u t i o n s of d r a g were summed to d e t e r m i n et h er e q u i r e dt h r u s ta n d ,h e n c e ,f u e lc o n s u m p t i o n . A parabolic d r a g polar was assumed with a d e s i g n C L , ~ of 0.4 r e p r e s e n t i n gt h ec e n t e r of t h e a i r f o i l s e c t i o nd r a gb u c k e t . Drag c o n t r i b u t i o n s due to t a i l l i f t andtail/wing i n t e r f e r e n c e were e s t i m a t e du s i n gb i p l a n et h e o r y (refs. 2 and 2 8 ) . The e n g i n e performanceandweight were s c a l e d from a b a s e l i n ee n g i n e as suggested by ref- erence17. The e n g i n eo p e r a t i n gc h a r a c t e r i s t i c s were determined as a f u n c t i o n of Mach number a n d a l t i t u d e from a model developed i n r e f e r e n c e 1 8 .

The modified direct o p e r a t i n g cost MDOC usedthesummation of t h e follow- ing costs: d e p r e c i a t i o n , s u p p o r t , spares, d e l a y , i n s u r a n c e , f u e l , m a i n t e n a n c e , l a n d i n g fee, crew, a t t e n d a n t s ,f u e ls e r v i c e , and c o n t r o l . The parameter MDOC differs from i n d u s t r ys t a n d a r dm e t h o d s by t h e i n c l u s i o n of support, d e l a y , a t t e n d a n t s , f u e l - s e r v i c e , c o n t r o l , a n d l a n d i n g - f e e costs. Operating costs were estimated u s i n gt h er e l a t i o n s h i p sf o u n di nr e f e r e n c e s 29 t o 32. T h i ss t u d y was p e r f o r m e dw i t ht h ef u e l cost set a t 0.2 U . S . d o l l a r s per l i t e r ($0.75per g a l l o n ) .P a r a l l e ls t u d i e s were also c o m p l e t e dw i t hf u e l costs of up to 0 . 4 U . S .

d o l l a r s per l i t e r ($1.50 per g a l l o n ) . The d e p r e c i a t i o n costs were c a l c u l a t e d u s i n gt h ea i r p l a n ep u r c h a s e price e s t i m a t e d from reference17andassuming a r e s i d u a lo f1 2p e r c e n ta n d a d e p r e c i a t i o n period of 1 4 y e a r s . The i n c l u s i o n of a c t i v e c o n t r o l s r e s u l t e d i n a n appropriate i n c r e a s ei nt h ep u r c h a s e price and maintenance cost as i n d i c a t e d by r e f e r e n c e s 4 , 5, and 20. It was assumed t h a t t h e same l e v e l of r e l i a b i l i t y a n d d i s p a t c h a b i l i t y c o u l d be maintainedand, t h e r e f o r e , t h a t t h e i n c l u s i o n o f t h i s new t e c h n o l o g yw o u l dn o tr e s u l ti n i n c r e a s e dd e l a y s or h i g h e ri n s u r a n c e rates.

A f t e rt h ep e r f o r m a n c ef u n c t i o n h a db e e ne v a l u a t e d ,t h eo p t i m i z e rc a l l e d a s e c t i o n of computercodewhichevaluatedthe set of c o n s t r a i n t sb e i n ga p p l i e d .

The list of a v a i l a b l ei n e q u a l i t yc o n s t r a i n t s ,w h i c hh a v e upper and lower bound- aries, is shown i nt a b l e 11. The c o n s t r a i n t st h a t were s e l e c t e da n dt h e limits t h a t were imposed were t h e means by whichthedesign was s p e c i f i e d .T h e r e were b o t h o p e r a t i o n a l or d e s i g n c o n s t r a i n t s a n d f l y i n g - q u a l i t y c o n s t r a i n t s .

A modified cost f u n c t i o n was formed by a d d i n gp e n a l t y terms to t h e perfor- mance i n d e xf u n c t i o n for e a c hc o n s t r a i n tf u n c t i o nt h a tv i o l a t e d its upper or lower limit. E a c hp e n a l t y was p r o p o r t i o n a l to t h es q u a r e of t h e amount of t h e v i o l a t i o n times a l a r g ew e i g h t i n g factor. When t h eo p t i m i z e rm i n i m i z e dt h e modified cost f u n c t i o n , it f o r c e dt h ec o n s t r a i n tv i o l a t i o n st o w a r dz e r o i f t h e w e i g h t i n gf a c t o r was s u f f i c i e n t l yl a r g e .C o n s t r a i n t st h a t are on or n e a rt h e boundary are s a i d to be a c t i v e .

C o n s t r a i n tf u n c t i o n st h a ti n v o l v e dt h e a i r c r a f t o p e r a t i o nd u r i n g cruise u t i l i z e dd a t as a v e d in' t h ed a t ab a s ed u r i n gt h e cruise p o r t i o n s . The a i r c r a f t aerodynamic moments and forces were determined,and it was a l t e r n a t e l y trimmed o u t i n take-off or l a n d i n gc o n f i g u r a t i o n s a t t h e appropriate speeds to d e t e r - mine t h e p e r f o r m a n c e , s t a b i l i t y , a n d trim c h a r a c t e r i s t i c s . The nondimensional s t a b i l i t y d e r i v a t i v e s used for t h e s ea n a l y s e s were s a v e di nt h ed a t ab a s e for approachandcruiseconfigurations.These were converted t o dimensionalderiv- a t i v e s (ref. 3 3 ) , and a f o u r t h - o r d e ra n a l y s i so ft h el o n g i t u d i n a ld y n a m i c s was performedusingsystemroutines a t LangleyResearchCenter. The roots o ft h e s e were used to e v a l u a t e many of t h e f l y i n g - q u a l i t i e s param- d i f f e r e n t i a l e q u a t i o n s eters which were part of t h ec o n s t r a i n tf u n c t i o n s .

The o p t i m i z a t i o n c o n t i n u e d u n t i l s a t i s f a c t o r y c o n v e r g e n c e was o b t a i n e d .

Typically,convergencerequiredontheorder of 1500 t o 2200 calls of t h e per- formancefunction to g e n e r a t ea n optimum a u g m e n t e df u n c t i o n ,w h i c hr e s u l t e di n t h e set of i n d e p e n d e n td e s i g nv a r i a b l e sw i t ht h e minimum performanceindexthat s a t i s f i e d t h e selected c o n s t r a i n tf u n c t i o n s .T h i sr e q u i r e da b o u t 1600 seconds O f e x e c u t i o n time usingtheLangleyResearchCentercanputer facilities.

Method o f C a p a r i s o n The r e s u l t s were normalizedusing a b a s e l i n e set o fd e s i g ns p e c i f i c a t i o n s .

T h e s e s p e c i f i c a t i o n s were primarily t h e m i s s i o n i n p u t s of table I a l o n g w i t h t h e m i l i t a r y l e v e l I f l y i n g - q u a l i t i e s criteria for t r a n s p o r t aircraft (ref. 3 4 ) .

F i n a l v a l u e s for boththeindependentanddependentdesignvariableswhich resulted f r a n t h e o p t i m i z a t i o n of t h e b a s e l i n e m i s s i o n are Shawn i n table I11 a l o n gw i t h a number ofperformance indices. A s mentionedpreviously,the modi- f i e d direct o p e r a t i n g cost per b l o c k hour was t h ef i g u r e - o f - m e r i t to w h i c ht h e c o n f i g u r a t i o n was optimized.Althoughthelevel I a i r c r a f t was o b v i o u s l y much more stable t h a nc u r r e n td e s i g n s (static m a r g i no fn e a r l y 43 p e r c e n t ) it r e p r e s e n t s a good b a s e l i n e ?s i n c e it usedthe most c o n s e r v a t i v eo ft h e proposed criteria for t h eu n a u g n e n t e df l y i n gq u a l i t i e s of t r a n s p o r t s .

S i n c e modified direct o p e r a t i n g cost per b l o c k hour was theoptimized per- formanceindex, a l l r e s u l t s are shown i n terms o ft h i sq u a n t i t yn o r m a l i z e d by the b a s e l i n e airplane performance. S p e c i f i c a l l y ,p e r c e n ts a v i n g si n modified direct o p e r a t i n g cost is theparameter t h a t is p l o t t e d as a f u n c t i o n of t h e various f l y i n g - q u a l i t i e s criteria b e i n gc o n s i d e r e dh e r e i n . I t was calculated as follows : RESULTS AND DISCUSSION F l y i n g - Q u a l i t i e s S e n s i t i v i t y S t u d y Static margin, or t h e d e g r e eo fs t i c k - f i x e d s t a t i c s t a b i l i t y ? h a s b e e n t h e parameter by which most a n a l y s e so fs t a b i l i t ya u g m e n t a t i o ns y s t e m sf o rt r a n s - ports w i t ha c t i v ec o n t r o l s havebeenevaluated.Figure 3 shows thenormalized m o d i f i e d - d i r e c t - o p e r a t i n g - c o s ts a v i n g s as a f u n c t i o n o f s t a t i c marginduring l a n d i n g . The lower c u r v ei n c l u d e st h e impact of adding a f l i g h tc o n t r o l cow puter for a u g m e n t i n gt h es t a b i l i t y ?w h i l et h eu p p e rc u r v er e p r e s e n t sn o tc a r r y - i n g an a u g n e n t a t i o n canputer. The d i f f e r e n c eb e t w e e nt h e t w o c u r v e s is n e a r l y c o n s t a n ts i n c et h ec a n p u t e r is n o t scaled o nt h ed e g r e e of i n s t a b i l i t y ; t h e r e - fore? it r e p r e s e n t s an i n i t i a li n v e s t m e n tp e n a l t y for thedevelopment, certifi- c a t i o n ? andmaintenanceof a required a u t a n a t i cf l i g h tc o n t r o ls y s t e m .A l t h o u g h t h e i n c l u s i o n of a l a r g e f l i g h t c a n p u t e r is e x p e c t e d to p r o v i d e t h e c a p a b i l i t y f o rs i g n i f i c a n ti m p r o v e m e n t si no t h e r areas such as s a f e t y ,o p e r a t i n ge f f i - ciency?and cost management, i n t h i s case it is c h a r g e d e n t i r e l y to t h e RSSAS system.

The s a v i n g s i n modified direct o p e r a t i n g cost between an unaugnented stable aircraft r e p r e s e n t a t i v eo fc u r r e n tc o n f i g u r a t i o n s (20 p e r c e n t static margin)and an augmentedunstable aircraft (-1 0 p e r c e n t static margin) is approximately1.4percent.Since f u e l cost m a k e s up a b o u t 40 p e r c e n t of t h e m o d i f i e dd i r e c to p e r a t i n g cost per b l o c kh o u r ,t h ef u e ls a v i n g sb e t w e e nt h e s e t w o a i r p l a n e s is a b o u t 4.2 p e r c e n t , as c a nb es e e ni nt a b l e 111. T h i s compares f a v o r a b l yw i t ht h e 3- t o 5 - p e r c e n ts a v i n g sp r e v i o u s l ye s t i m a t e df o rt h i s class of aircraft w i t h similar r e d u c t i o n si n static s t a b i l i t y (refs. 1 a n d2 ) .I n terms of 1 9 7 9 d o l l a r s as comparedwiththeunaugmented20-percent static margin design case, t h i s is e q u i v a l e n t to a savingsofabout$86,400 per y e a r , or $1 .21

m i l l i o n overthe lifetime of t h e aircraft i n M D O C . A s f u e l cost rises over

t h e $0.20 p e r l i t e r ( $ 0 . 7 5p e rg a l l o n )u s e di nt h i ss t u d y ,t h es a v i n g sp r o j e c t e d w i l l beevengreater.

T h e s es a v i n g si n c l u d e a b e s t estimate o ft h ei n c r e m e n t a l costs fromthe e x t r ae n g i n e e r i n g ,f l i g h tt e s t i n g ,a n dq u a l i t yc o n t r o lr e q u i r e m e n t st h a ty i e l d t h es u b s e q u e n ti n c r e a s ei np u r c h a s ep r i c e .A d d i t i o n a l l y ,t h ei n c r e a s ei n main- tenance costs is also r e f l e c t e d . The f a c tt h a ts u c h a l a r g es a v i n g s is still p o s s i b l eg i v e sc r e d e n c e to t h e p o s s i b i l i t y t h a t t h e t e c h n o l o g i c a l and s a f e t y b a r r i e r s t h a t impede t h e u s e of RSSAS canbeovercome i na ne c o n o m i c a l l yf e a s i - b l ef a s h i o n .

The improvementsnotedinthiscomparison came p r i n c i p a l l yf r o mr e d u c t i o n s i n maximum take-offgrossweightandinanimprovementin maximum CL/CD. Sche- matic d i a g r a m so ft h e t w o a i r c r a f t are shown i n f i g u r e 4 , and key d e s i g nd a t a are compared i nt a b l e 111. The take-offgrossweight was reduced 22 000 N (4946 l b f ) ,w i t h 7300 N (1 641 l b f )o f it a t t r i b u t a b l e t o t h e3 2 - p e r c e n tr e d u c t i o n o f t h e t a i l area. The 3.9-percent improvement in maximum CL/CD was due p r i n - c i p a l l y to a 1 7 - p e r c e n tr e d u c t i o ni n t o t a l t a i l drag(wettedandinduced).

Anotherparameterthat is o f t e nc o n s i d e r e dd u r i n gf l y i n g - q u a l i t i e sa n a l y s e s is manuever margin, or t h ed e g r e eo f maneuver s t a b i l i t y( r e f . 4 ) . Maneuver s t a b i l i t y is p r o p o r t i o n a l to t h ee l e v a t o rd e f l e c t i o nr e q u i r e dp e ru n i tg r a v i t y n o r m a la c c e l e r a t i o n .S i n c e it can be shown t h a tt h ed i f f e r e n c eb e t w e e n s t a t i c marginandmaneuvermargin is a p p r o x i m a t e l yc o n s t a n t for a g i v e na i r p l a n e ( r e f . 21 ) , t h e f a c t t h a t s a v i n g s i n m o d i f i e d d i r e c t o p e r a t i n g cost have a similar t r e n df o rb o t h static marginandmaneuvermargin is n o ts u r p r i s i n g( f i g s . 3 and 5 ) .S i n c ep i l o t s would be unable to c o n t r o la na i r c r a f tw i t h a n e g a t i v e maneuvermargin,allowingnegativemaneuvermargin as a d e s i g n c r i t e r i o n f o r unaugmented f l y i n g q u a l i t i e s would be tantamount t o a s s u m i n gt h a tt h ea i r p l a n e would be l o s t i nt h ee v e n to f a c o n t r o ls y s t e m f a i l u r e .

Time to d o u b l et h ea m p l i t u d eo ft h el o n g i t u d i n a ld i v e r g e n td y n a m i c s is sometimes s p e c i f i e di nf l y i n g - q u a l i t i e s criteria andconsideredinhandling- qualities s t u d i e so fu n s t a b l ea i r p l a n e s( r e f .7 ) .S i n c et i m e - t o - d o u b l ev a l u e s result fromunstable root l o c a t i o n so ft h e dynamic e q u a t i o n s ,t h e y are c r i t e r i a a p p r o p r i a t ef o rs t u d y i n gt h e unaugmentedmotionsoftheairplane. The s a v i n g s i nm o d i f i e dd i r e c to p e r a t i n g cost as a f u n c t i o n of time to d o u b l ea m p l i t u d ei n approach are shown i nf i g u r e 6. I n i t i a l l y , as t h e time to double is reduced from a marginallyunstablevalueof55seconds, small d e c r e a s e s i n d i r e c t oper- a t i n g cost are o b t a i n e d . However, as t h e time to d o u b l ed e c r e a s e sf u r t h e r ,t h e s a v i n g s i n d i r e c t o p e r a t i n g cost i n c r e a s e r a p i d l y u n t i l a v a l u eo f 2 seconds for time to double anplitude. A t t h i s point,thedata show thatother con- straints becane c r i t i c a l and prevent any other improvenents to be made i n savings of modified direct operating cost fran reducing the time-to-double constraint.

Flying-Qualities Design Criteria Generally, when flying-qualities design criteria or regulationsare devel- oped, they encanpass a variety of parameters i n severalflightoonditions,for example, references 7 to 1 2 and 34. To providetheappropriate aerodynamic con- tribution to the stability, the unaugmented flying qualities must be specified.

These inherent characteristics are significant when consideration is given t o potentialfailure modes of the autanatic f l i g h t control system. Philosophi- cally, it becanes a canpromisebased on the minimun acceptable handling qualities.

One set of longitudinal-flying-qualities criteria that designers have appliedtothe design of transports configured w i t h RSSAS is the military flying-qualitiesspecifications(ref. 34). The short-period frequency require- ment i n approach is one of the activeconstraints i f thesemilitaryspecifica- tionsareutilized asdesign criteria. Short-periodfrequency is plotted i n figure 7 w i t h the narrow range of applicability, i n terms of n/a, for t h i s s t u d y shown. Assuming constant n/a, thesensitivity of optimal direct oper- ating oost toshort-period frequency constraint is shown i n figure 8.

As expected, substantial benefits were initially realized when relaxing theshort-period frequency criterionfranlevel I tolevels I1 and 111. The economicimprovenent was anticipated because it is generallyacceptedthat level I for transports is extremely harsh (ref. 35). Observing thelarge t a i l surfaces of thebaselineconfiguration(level I) i n table I11 helps to i l l u s - trate t h i s point. I n fact,reference 35 points out that modern transports do not, i n general, satisfy these criteria without augnentation, i n spite of their generallyacceptable f l y i n g qualities.

The problem w i t h the military specifications and a nunberof other - longitudinal-flying-qualities design criteria(refs. 1 0 and 1 1 ) is that they rely upon specifying modal damping ratios and frequencies. I n thecase of unstableairplanes,discussion of the dynamic longitudinal modes i n terms of damping ratios and frequenciesloses its meaning. Therefore, new longitudinal- flying-qualities design criteria are needed fortransportsconfigured w i t h a reduced s t a t i c margin.

The c r i t e r i a proposed i n references 8 and 9 have parameters which could be usefulfor imposing flying-qualitiesspecificationsforlongitudinally unstableairplanesat the preliminarydesignlevel. These criteriaare shown i n figure 9. The abscissa and ordinatearecoefficients of thecharacteristic polynanialthatresultsfran a linearanalysis of theshort-period mode, enabl- i n g easy considerationasconstraintfunctions. However, theregionindicated i n the figure by a dashed line is where unaugmented transports designed w i t h RSSAS are expected t o f a l l , which is outsidetheareacontainingtheflying- qualitiesdata. T h i s lack of appropriatedataillustrates another problem i n designing transports with RSSAS. Although the representation of flying quali- ties in figure 9 would be useful for developing longitudinal-flying-qualities design criteria, there is a need to collect simulator and flight test data respect to the minimum acceptable handling qualities of transports.

Impact of Related Design Constants Landing gear location.- Several design constants that were input for the baseline mission have a significant impact upon stability and control charac- teristics. One such factor is the main landing gear location. Current practice requires that its structure be located such that the loads are carried in the wing spar. Industry estimates that the maximum aft position that is structur- ally feasible is 65 percent of the mean aerodynamic chord (ref. 5 ) . Since a margin between the center of gravityand landing gear is necessary to insure enough nose-wheel steering traction, and since the supercritical airfoil data used in OPDOT assume large pitching moments in cruise, the main gear position usually constrains the most aft allowable center-of-gravity position for transports with reduced static stability.

Figure 1 0 shows the impact of relaxing this constraint for each of the three levels of military flying qualities. Nearly 1 percent savings in modified direct operating cost per block hour could be realized by allowing the main ge to be located off the spar, provided that the structural weight penalties asso- ciated with the relocation would be negligible. Since this corresponds to a fuel savings of about 3 percent, it suggests a possible area for further research.

Loadabi1ity.- Another factor that had a bearing upon the results reported herein was an assumption that the allowable center-of-gravity range be at leas 1.2 meters ( 4 feet). The impact of reducing the required loadability from

1 .2 meters to 0.61 meter ( 2 feet) to 0 meters is shown in figure 11 . It is

readily apparent that the biggest improvements came from the first reduction 0.61 meter. Only modest improvements were possible with further reduction, and this analysis ignores the cost of installing and operating a center-of-gravity control system that would certainly be necessary in this region of loadability.

However, since the benefits in modified direct operating cost were slightly greater than 1 percent ( 3 . 2 percent savings in fuel), it may be worthwhile to pursue schemes to allow such reductions. The new generation of transport designs already incorporate load cells in the gear with computer monitoring for the optimal placement of cargo at the gate.

Take-off stabilizer trim angle.- Manufacturers have been expected to demonstrate that their transports are capable of satisfying the nose gear unstick requirement with the horizontal stabilizer in the most adverse trim position. This constraint, which is satisfied at the forward center-of- gravity limit, was shown to be extremely harsh. In fact, savings in direct operating costof over 4 percent were indicated when the stabilizer was allowed to be trimmed to the position anticipated for climbout. It seems highly rea- sonable that for the corresponding $3.9 million that could be saved during the lifetime of the airplane, a suitable compromise between added complexity and safety could be reached to insure proper tail positioning during take-off.

1 0 Related Observations Useful information for the aircraft designer would be a list of the design constraints that tended to be active at the optimal design point. It was deter- mined from studying the program output that virtually without exception the following constraints were active at the converged solution point: ( 1 ) cruise thrust: ( 2 ) second-segment climb gradient: ( 3 ) landing field length: ( 4 ) nose- wheel steering traction; and (5) passenger volume. The first two were most sensitive to changes in thrust and wing aspect ratio, while landing field lengt was most influenced by wing area. Nose-wheel steering traction was a function of aft center of gravity, and passenger volume required a minimum fuselage length. As anticipated, the chosen flying-qualities constraint parameters were also active at the design point, and the solution was principally affected by this constraint through adjustments to the horizontal-tail area and aspect rat Reference 36 predicts that the optimum tail load, in terms of drag, would be a download for high downwash gradients. The low-tail geometry of this study was located in regions of high downwash; therefore, it was no surprise when the optimum design points for all configurations had tail lift coefficients ranging between -0.05 and -0.12. This result was not assumed in the formulation, but a model of downwash and multiple-lifting-surface interference effects was included in the performance evaluation. The optimizer adjusted the design variables, principallyin this case those which impacted tail volume and center of gravity, to obtain the minimum cost in the presence of control and stabil constraints. This result helped to validate the conclusions of reference 36.

As indicated in references 1 2 and 13, an analysis was performed to insure that (1 ) the unaugmented configuration was capable of being augmented to good flying qualities; ( 2 ) the control deflections required for augmentation would be sufficiently small to avoid significant control surface drag contributions: and ( 3 ) the control surface deflection rates commanded by the automatic control system would be sufficiently low in turbulence to be achievable. These goals were accomplished by simulating a pitch-attitude-hold/pitch-rate-command auto- pilot in heavy turbulence.

The following factors were then available as inequality constraint func- tions in cruise and approach: ( 1 ) pitch attitude feedback gain, ( 2 ) pitch rate feedback gain, ( 3 ) variance of elevator deflection in turbulence, and ( 4 ) vari- ance of elevator deflection rate in turbulence. However, except for when the unaugmented configuration was designed for extremely low time-to-double and maneuver margin, all configurations that the optimization generated satisfied these constraints. This was an indication that for the range of values con- sidered in the research, the resulting configurations could be augmented to good flying qualities.

Since the price of fuel has already matched the $0.20 per liter ($0.75 per gallon) used in this study, a series of design runs was performed with higher fuel prices. When the baseline was reconfigured to reflect the inflated fuel prices, it was observed that the same trends existed with slightly greater mag- nitudes in savings with respect to the flying-qualities parameters. This indicatesthat t h e benefits of utilizing new activecontrolstechnologiesto reduce theinherent s t a t i c s t a b i l i t y should be increasingly significant as fuel priceescalates.

CONCLUDING REMARKS A series of design runs u t i l i z i n g a computer progran fortheoptimal pre- liminarydesign of transport aircraft was used to s t u d y the impact of unaug- mented flying-qualities design c r i t e r i a and theinfluence of u t i l i z i n g relaxed- static-stability augnentation systems. Transportsoptimallyconfigured w i t h relaxed s t a t i c s t a b i l i t y showed a potential savings i n direct operating cost of 1.4 percent when canpared w i t h transports w i t h conventional s t a t i c margins.

T h i s translates into a fuel savings of 4.2 percentforthe 5600-kilaneter (3000-nautical-mile)range, 200-seat transport w i t h a cruising Mach nunber of 0.8 which is considered i n t h i s report.Similartrends of savings can be observed when evaluating s t a t i c margin, maneuver margin, ortime to double anplitudeastheconstraininghandling-qualities parameter.

It was observed that the same trends of savings i n direct operating cost i n fuelprice. It was also shown that were expected forlargevariations efforts to remove the maximun rearward position constraint on thelanding gear would be rewarded w i t h gains of nearly 1 percent i n modified directoperating cost.Additionally, a reduction of allowablecenter-of-gravity range from about 1.2 metersto about 0.6 meter couldsave nearly 1 percent i n modified directoperatingcost. A constrainttorequiretheelevatortorotatethe aircraft duringtake-off w i t h the stabilizer i n its most adverseposition was found to be very harsh i n termsofeconomic profitability, penalizing the air- craft over 4 percent i n directoperatingcost.

Constraintstoinsure enough thrust i n cruise, to satisfy second-segment climb gradients,to f u l f i l l landing field length requirments, to provide enough tractionfor nosewheel steering, and to allaw enough volmefor passengers were shown to be active at the design point along w i t h the critical flying-qualities criteria. The optimum airplane tended to f l y thecruise mission w i t h a download on the t a i l as was predictedfor an airplane w i t h thehorizontal t a i l i n the influence of a strong downwash field.

I n thecourse of the study, it was determined, through thehypothetical design and evaluation of a simple autopilot,thatthe designsconsidered were practically augmentable to good flyingqualities. The rms deflections and rates of deflection of theelevator due to heavy turbulence were acceptableas were the feedback gainsrequiredto achieve satisfactory augmentation.

Most of the flying-qualities criteria proposed for unaugnented transports proved to be inappropriate,sincetryingtospecify a modal frequency and damp i n g ratioloses its significanceforunstableairplanes. I n particular, it was shown that the military specifications, when used for unaugmented airplane flying-qualities design criteria, were particularly harsh for t h i s category of aircraft.

1 2 It is recanmended t h a ts y s t e m a t i cf l i g h ta n ds i m u l a t i o nr e s e a r c h be under- taken to provide a data b a s ef o rd e v e l o p i n g u s e f u l u n a u g n e n t e df l y i n g - q u a l i t i e s design criteria f o rt r a n s p o r t sc o n f i g u r e dw i t hr e l a x e d s t a t i c s t a b i l i t y . Fur- t h e r m o r e ,a ni n t e g r a t e de f f o r tb e t w e e nt h ed e s i g n e r and the h a n d l i n g - q u a l i t i e s specialist is required i n order ( 1 ) to e n h a n c et h ea p p l i c a b i l i t yo f new criteria to designmethodologies; (2) to m a i n t a i n s u f f i c i e n t m a r g i n s of f l i g h t s a f e t y ; and ( 3 ) to i n s u r e t h a t e c o n a n i c p r o f i t a b i l i t y is considered as any new criteria are developed.

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Active Controls Concepts for Subsonic Transport Aircraft. Volume I1 -

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4. Sizlo, T . R.; Berg, R. A.; and Gilles, D. L.: Development of a Low-Risk Augmentation System for an Energy-Efficient Transport Having Relaxed Static Stability. NASA CR-159166, 1979.

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7. Grantham, William D.; and Deal, Perry L . : Simulator Study of Minimum Acceptable Level of Longitudinal Stability €or a Representative STOL Con- figuration During Landing Approach. NASA "N D-7733, 1974.

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12. Sliwa, Steven M . : Impact of Longitudinal Flying Qualities Upon the Design of a Transport With Active Controls. AIAA-80-1570, Aug. 1980.

13. Sliwa, Steven M; and Arbuckle, P. Douglas: OPDOT: A Computer Program for the Optimum Preliminary Design of a Transport Airplane. NASA TM-81857, 1980.

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, , , .. .. - .. . . - . . . . . _". . ...... . . .. .

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T A B m 1.- KEY DESIGN ONSTANTS UTILIZED FOR DESIGN OPTIMIZATION (a) Mission

C r u i s e Mach number . . . . . . . . . . . . . . . . . . . . . . . . 0.80

Divergence Mach nunber . . . . . . . . . . . . . . . . . . . . . . 0.84

Design range. lan . . . . . . . . . . . . . . . . . . . . . . . . . 5600

Nunber of seats . . . . . . . . . . . . . . . . . . . . . . . . . 200

Cargo. N . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33 400

Maximum l i f t c o e f f i c i e n t . . . . . . . . . . . . . . . . . . . . . 3.15

L a n d i n gf i e l dr e q u i r e m e n t . m . . . . . . . . . . . . . . . . . . . 2440

Take-off f i e l d r e q u i r e m e n t . m . . . . . . . . . . . . . . . . . . 3050

( b ) Geanetr y

Wing sweep angle. deg . . . . . . . . . . . . . . . . . . . . . . 26.4

. . . . . . 0.12

Wing t h i c k n e s s ratio . . . . . . . . . . . . . . . . .

. . . . . . 0.38

Wing taper ratio . . . . . . . . . . . . . . . . . . .

Wing incidence angle. deg . . . . . . . . . . . . . . . . . . . . . 2

Wing g e a n e t r i c twist. deg . . . . . . . . . . . . . . . . . . . . . 5

T a i l t h i c k n e s s ratio . . . . . . . . . . . . . . . . . . . . . . . 0.10

. . . . . . . 3 0

T a i l sweep angle. deg . . . . . . . . . . . . . . . .

Tail taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . 0.4

Vertical- t a i l sweep.deg . . . . . . . . . . . . . . . . . . . . . . 3 5

Ratio of rudder area to v e r t i c a l - t a i l area . . . . . . . . . . . . 0.30

Ratio o fe l e v a t o rc h o r d to h o r i z o n t a l - t a i lc h o r d . . . . . . . . . 0.25

Ratio o ff l a ps p a n to wing span . . . . . . . . . . . . . . . . . 0.6

Maximun f l a pd e f l e c t i o n .d e g . . . . . . . . . . . . . . . . . . . . 4 5

F u s e l a g ed i a n e t e r . m . . . . . . . . . . . . . . . . . . . . . . . 5.08

Height of aerodynamiccenterabovec.g.,fraction MAC . . . . . . 0.08

Height of t h r u s tv e c t o ra b o v ec . g . ,f r a c t i o n MAC . . . . . . .

-0.1 2

Height of h o r i z o n t a l t a i l above c.g. . . . . . . . . . . . . . . . . 0

N u m b e r of e n g i n e s . . . . . . . . . . . . . . . . . . . . . . . . . 2

.

TABLE I . . CONCLUDED

(c) Economics

Fuel cost. $/L . . . . . . . . . . . . . . . . . . . . . . . . . . 0 . 2 0

Load factor . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.55

Passenger revenue . #/seat-km . . . . . . . . . . . . . . . . . . . 4.9

Utilization rate. hr/yr . . . . . . . . . . . . . . . . . . . . . 3200

Depreciation period. yr . . . . . . . . . . . . . . . . . . . . . . 1 4

Residual value. % . . . . . . . . . . . . . . . . . . . . . . . . . 1 2

Tax rate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.48

Year of study . . . . . . . . . . . . . . . . . . . . . . . . . . 1 9 7 9

Assuned annual inflationrate . . . . . . . . . . . . . . . . . . 0.07

Nmber of prototype aircraft . . . . . . . . . . . . . . . . . . . . 2

Aircraft fleet size . . . . . . . . . . . . . . . . . . . . . . . 250

I n i t i a l production rate. per month . . . . . . . . . . . . . . . . 0.5

F u l l production rate. per month . . . . . . . . . . . . . . . . . . 5

Engineering rate (1 974) . $/hr . . . . . . . . . . . . . . . . . 1 9 . 5 5

Tooling rate (1 9 7 4 ) . $ / h r . . . . . . . . . . . . . . . . . . . 1 4 . 0 0

Labor rate (1 974) . $/hr . . . . . . . . . . . . . . . . . . . . 10.90

Engines for test aircraft . . . . . . . . . . . . . . . . . . . . . 3

Ratio of manufacturer'sairframe weight to take-off weight . . . . 0.75

(d) Miscellaneous

Maximun dynamic pressure. N/m2 . . . . . . . . . . . . . . . . . . 5 . 1 3

. . . . . . . . . . . .

Pressurized volume. m3 . . . . . . . . . 178.2

Nunber of pilots . . . . . . . . . . . . . . . . . . . . . . . . . . 3

Number of attendants . . . . . . . . . . . . . . . . . . . . . . . . 8

Air conditioning flow rate. kg/min . . . . . . . . . . . . . . . . 200

Autopilot channels ( w i t h multiplexers) . . . . . . . . . . . . .

. . 5

. . . . . . . . . . . .

Generator capacity. kV-A . . . . . . . . . 750

Maintenance ccmplexity factor . . . . . . . . . . . . . . . . . . 1 . 6

Hydraulics volune flow rate. L/min . . . . . . . . . . . . . . . . 300

Nmber of inertial platform systems . . . . . . . . . . . . . . . . 1

engine on-time . . . .

Ratio of auxiliary-power-unit on-time to . 0.1

. . . . . . . . . . . . . 0.15

Ratio of first class to economy seating . . . . . . . . . . . .

Maximun speed. m/s . . . . . . . . . . . 248.5

Airfoil design l i f tc o e f f i c i e n t . . . . . . . . . . . . . . . . . 0 . 5

Baseline engine . . . . . . . . . . . . . . . . . . . . . . . . . CF-6

Elevator servo time constant. sec . . . . . . . . . . . . . . . . 0.1

Curved windshield Supercriticalairfoil technology Sane nonlinear aerodynamics terms 1 8 TABLE 11.- SAMPLE MEQUALITY CONSTRAINT FUNCTIONS AVAILABLE DURING DESIGN OPTIMIZATION Mission C r u i s et h r u s t climb g r a d i e n t Second-segment Missed-approach climb gradient Landing f i e l d l e n g t h Take-off f i e l d l e n g t h P a s s e n g e rv o l m e C r u i s e a l t i t u d e F u e l volume C r u i s e l i f t c o e f f i c i e n t C o n t r o l Nose g e a rs t e e r i n gt r a c t i o n Nose gearunstickduringtake-off T a i l - l i f t - c o e f f i c i e n t s t a l l margininapproach E l e v a t o rd e f l e c t i o n a S t a b i l i t y Static margina Maneuvermargina S h o r tp e r i o df r e q u e n c y a S h o r t p e r i o d dampinga Phugoidfrequencya Phugoiddampinga Mode frequency ratioa Time-to-half (double) a Vertical r e s p o n s ef a c t o r a A u t o p i l o t P i t c hf e e d b a c kg a i n a P i t c h rate feedbackgaina E l e v a t o rv a r i a n c e a E l e v a t o r rate v a r i a n c e a a A v a i l a b l e for both cruise andapproachconfigurations.

1 9 L .. . . . ._ ., TABLE 111.- CBARACTERISTICS OF SAMPLE " ."

I n d e p e n d e n td e s i g nv a r i a b l e s S m e ~ ~ . " - - Design case T, F u e l , W t t , Wte , Wtto, Lf I W k , ARIl kN kN kN kN 1 CD,O,W m . " __ . . " .. .

. . ~. .

Level Ia ( b a s e l i n e ) 199.3 52.7 103.6 6.38 338.4 145.5 44.2 10.75 677.7 1230 26 033 0.0078 L e v e lI I a 188.7 1 1 .66 84.4 5.1 6 31 3.2 149.4 34.4 657.0 1 1 92 24 235 .0079 L e v e l I I I ~ 188.6 11.65 84.1 5.15 313.2 149.1 34.3 656.4 1 1 91 24 223 .0079 S t a t i c m a r g i n = '20% 184.1 12.55 43.0 5.34 292.6 154.8 24.5 640.0 1 1 55 22 501 .0080 S t a t i c m a r g i n = -10% 183.6 5.16 294.7 12.46 49.7 153.9 25.9 641 .7 1 159 22 569 S t a t i c m a r g i n = -5% 184.6 12.45 51 .8 5.71 295.8 28.1 153.9 646.9 1167 22 709 S t a t i c m a r g i n = 0% 185.9 12.50 53.9 6.41 297.1 157.1 30.8 654.0 1 1 75 22 970 S t a t i c m a r g i n = 5% 186.3 12.28 58.2 6.33 300.2 154.8 31 .8 654.1 1177 23197 = 10% S t a t i c m a r g i n 185.5 12.10 66.4 5.06 302.5 151.2 29.8 647.9 1173 23 402 S t a t i c m a r g i n = 20% 187.4 11.90 73.4 5.61 307.6 33.5 655.2 1 1 85 23 807 .0079 S t a t i c m a r g i n = 20%b 186.8 5.55 306.7 11.94 73.3 33.2 652.3 11 81 23 716 .0080 Maneuvermargin = 0% 184.3 12.37 54.5 5.1 1 298.1 153.7 27.0 644.6 1164 22 907 Maneuvermargin = 1 0 % 185.4 12.20 60.4 5.54 300.2 152.8 29.9 648.6 1172 23198 Time to d o u b l e = 55 sec 184.5 12.19 59.1 4.63 299.3 151 .4 26.6 641 .9 1163 23 077 T h e to double = 40 sec 184.0 12.10 59.1 4.61 299.7 150.4 26.6 640.6 1162 23 080 Time t od o u b l e = 20 sec 183.7 4.88 56.8 299.9 149.8 26.9 640.1 1161 23 027 t t Time to double = 3 sec 181.6 43.8 4.95 296.8 147.3 23.6 630.2 1146 22 599 ~ ~. . - . " a M i l i t a r yl e v e ls p e c i f i c a t i o n s from MIL-F-8785B ( r e f .7 4 ) .

h o a c t i v ec o n t r o ls y s t e m si n c l u d e di n cost or w e i g h te s t i m a t e s .

dependentdesignvariables I Performance i n d i c e s _ _ _ ~. " -~ . .. ~ Fa1 I F A R E , ' D , O , t *D,o, tot CL, t cD, t % $/seat-hn

-~ " .. . . " ~~ L ~

" 0.0037 0.01 93 -0.124 I. 0045 19.1 15.87 736 163 1978 904 12.9 0.051 685 150 14.1 .0032 .0190 -.112 .0039 20.1 15.36 1904 900 .050 .0032 15.35 685 160 900 14.1 .050 .0190 -.112 .0039 20.1 1904

.0018 .0177 -. 206 .ooze 21.4 14.94 635 157 1836 897 15.2

.0020 .0179 -. 179 .0030 21.2 14.98 1844 898 15.0

-0021 -. 173 21.2 645 1851 14.9

. o l e o 15.06

1 I

.0022 . o l e 1 -. 166 21 . 2 15.16 649 158 1859 14.7

655 158 .0024 .0182 -.154 .0031 21 .o 15.19 1867 899 14.6 - O r .050

.0026 . 01 8 5 -. 135 20.6 15.14 661 159 1870 14.6

.0034 899

.0029 .0187 -. 123 .0036 20.4 15.28 673 159 1888 900 14.3

*0029 14.96 670 1 5 5 896 14.8

.0187 -. 123 .0036 20.4 1871

0022 . o l e 1 -. 1 64 21 .o 15.05 647 158 1853 898 14.9

.0031

-0024 -. 149 .0033 20.8 15.13 656 1860 14.7

.0023 -. 1 51 1 5 . 0 3 652 1856 14.9 20.8

I

- -. 150 1 15.02 652 1856

1 8r 1 - O j S .048

.0181 -. 156 .0032 15.02 651 157 1855

.0018 .0177 -. 1 99 20.9 639 1836 898 15.3

.0029 14.86 155 ~ ". .. ~ " " - . . ___

...., ,.,.,...,. , . . . I ..... I. 111.. . I . a 1 1

c . g .

BASELINE ENGINE m R U S T ETC .

ETC .

PERFORMANCE

- DATA

z L FUNCTION

' - BASE R O I c FARE I _ LID

W t CONSTRAINT +

FUEL ETC . I1 FUNCTIONS ( 5 2 ) ENGINE-OUT PERFORMANCE FIELO LENGTH FLYING QUALITIES GEOMETRY CONSTRAIKTS CONTROL POWER ETC .

c _ NO CONSTRAINED SOLUT I ON Figure 1.- Schematic diagram representingtheconstrained parameter optimizationlogic. Nunbers i n parenthesesare nunber of parameters available i n OPDOT (ref. 1 3 ) .

PERFORMANCE DOC ROI INDEX DESIGN VARlABlfS

''

COMPUTATION COST

1 1

MANUFACTURING MAINTENANCE I O C

t t

FEATURES 0 INDUSTRY STATISTICS FOR COSTS AND WEIGHTS CRUISE STEP DATC@A-TYPE S T A B l L l N AND CONTROL DERIVATIVES NON-LINEAR,ITERATIVE TRIMSOLUTION 0 MULTIPLE-STEP, SUBOPTIMALCRUISE~LIMB 0 GENERALIZED INTERFERENCE DRAG I COMPUTER GRAPHICS OUTPUT OPTION STABILITY AND CONTROL AcL, t Figure 2.- Hierarchy of logic flow forevaluatingthe performance indices.

FLIGHT CONTROL COMPUTER ON BOARD

- - - cl NO FLIGHT CONTROL COMPUTER

-

8 - 6 - S A V l N G S - I N MDOC, 4 % 2 - I I I I I I I I

'

0 - -20 -15 -10 -5 0 5 10 15 20 S T A T I C M A R G I N , % M A C F i g u r e 3.- P e r c e n t s a v i n g s i n m o d i f i e d d i r e c t o p e r a t i n g cost w i t h respect t o t h e b a s e l i n e c o n f i g u r a t i o n as a f u n c t i o no f s t a t i c m a r g i nr e q u i r e d dur ing landing .

2Oo/o STATIC M R G I N -lO"/o STAT1 C M R G I N 2 2 2 2 Sw = 183.6 m (1976 ft 1 sw = 186.8 m (2016 ft 1 ARw = 12.46 ARw = 11.94 2 2 2 2 st = 49.7 m (535 ft ) st = 73.3 m (789 ft ) T = 294.7 kN (66 x loJ Ibf) T = 306.7 kN (69 x 1 0 ' Ibf) F i g u r e 4.- Sketches of t w o o p t i m a l l yd e s i g n e d airplanes. The l e f t a i r p l a n e is representative of a t r a n s p o r t s i z e d w i t h c o n v e n t i o n a l f l y i n g ' q u a l i t i e s , a n d r i g h t t r a n s p o r t was s i z e d w i t h r e l a x e d static s t a b i l i t y a n dr e q u i r e s an RSSAS s y s t e n for a d e q u a t ef l y i n gq u a l i t i e s .

S A V I N G S IN MDOC,

L

0 L I I’ I I -15 -10 -5 0 5 10 15 M A N E U V E R M A R G IN. % M A C F i g u r e 5 . - P e r c e n t s a v i n g s i n m o d i f i e d direct o p e r a t i n g cost w i t h respect to t h e b a s e l i n e c o n f i g u r a t i o n as a f u n c t i o n of minimunmaneuvermargin required i na p p r o a c h . It s h o u l d be n o t e dt h a tu n a u g n e n t e da i r c r a f t withnegativemaneuvermarginswould be u n c o n t r o l l a b l e .

6 - S A V I N G S IN M D O C , % 4 - 2 - ~ 0 10 20 3 0 40 50 60

T I M E - T O - D O U B L E , sec ji

I 1 F i g u r e 6.- P e r c e n t s a v i n g s i n m o d i f i e d direct o p e r a t i n g cost w i t h respect to t h e b a s e l i n e c o n f i g u r a t i o n as a f u n c t i o n of minimum time to double amplitude inapproach.

c l -

I -

1. c

L

T

‘;r

ak- ANTICIPATEDRANGE OF n/a

FOR UNAUGMENTED TRANSPORT

0.1

1 -

1.0 10 100

nla, glrad

Figure 7.- Short-periodfrequencyboundariesformilitary s p e c i f i c a t i o n s ( r e f . 3 4 ) .

lo[ 8

S A V I N G S I N MDOC, ' 1 - 7

- LEVEL 'II

i % 4 , LEVEL I I I I 0 . 2 . 4 . 6 . 8 1.0 SHORT-PERIOD FREQUENCY, w sec -1 SP' F i g u r e 8.- S e n s i t i v i t y of p e r c e n ts a v i n g si nm o d i f i e d direct o p e r a t i n g cost w i t h r e s p e c t to b a s e l i n e c o n f i g u r a t i o n s to s h o r t - p e r i o d - f r e q u e n c y c o n s t r a i n t .

PR = 3.5

50 t P R Z 6 . 5 /

EMERGENCY OPERATION

\ TYPICAL RANGE OF VALUES

---,/ FOR UNAUGMENTED TRANSPORTS

,.-

\ I I I I 4 6 8 10 :..Q"-/ 2 I SHORT-PER IOD DAMP INGTERM, 2SwspI sec -1 Figure 9.- Proposed b o u n d a r i e so fl o n g i t u d i n a ls h o r t - p e r i o d c h a r a c t e r i s t i c s for s p e c i f y i n g f l y i n g q u a l i t i e s f r m r e f e r e n c e 8 .

M IL-F-8785B CR ITER I A ( REF. 34)

___O LEVEL I

""" -0 L E V E L I I ""

+' LEVEL I I I

S A V I N G S % 4 IN MDoC' .65 .70 .75 .80 .85 .90 .95 1.0 LOCATION OF L A N D I N G GEAR, X / M A C lg Figure 10.- Impact of maximun a f tl o c a t i o no fl a n d i n gg e a r upon p e r c e n t s a v i n g s i n modified direct o p e r a t i n g cost w i t h respect to b a s e l i n e c o n f i g u r a t i o n .

ALLOWABLE C. G. TRAVEL __O 1.22 METERS

"_ - U .61 METERS

" + 0 METERS

, L U I - . I I I I I I

-20 -15 -10 -5 0 5 10 15 20 S T A T I C M A R G I N . % M A C F i g u r e 11.- S e n s i t i v i t y o f p e r c e n t s a v i n g s i n m o d i f i e d direct o p e r a t i n g cost w i t h respect to b a s e l i n e c o n f i g u r a t i o n to allowable center-of- g r a v i t y t r a v e l as a f u n c t i o n o f required s t a t i c margininapproach.

1. Report No. 2. Government Accession No. 3. Recipient's C a t a l o g No.

NASA TP-1760 4. Title and Subtitle 5. Repon Date December 1980 ECONOMIC EVALUATION OF FLYING-QUALITIES D E S I G N 6. Performing Organization Code CRITERIA E U R A TRANSPORT CONFIGURED WITH RELAXED STATIC STABILITY 505-34-33-05 7. Author(s1 8. Performing Organization Report No.

L-13944 Steven M. Sliwa 10. Work Unit No.

9. Performing Organization Name and Address NASA Langley Research Center 1 1 . Contract or Grant No.

Hampton, VA 23665 13. Type of Report and Period Covered Technical Paper 2 . Sponsoring AgencyNameandAddress NationalAeronautics and Space Adninistration 14. Sponsoring Agency Code Washington, DC 20546 5. Supplementary Notes 6. Abstract Direct a m s t r a i n e d parameter optimization was used to optimally size a mediun-range transportfor minimun direct operating cost. S e v e r a ls t a b i l i t y and control con- s t r a i n t s were varied to study the sensitivity of theconfiguration to specifying the unaugnented flying qualities of transportsdesigned to take maximun advantage of relaxed-static-stabilityaugmentationsystems.Additionally, a nunber of handling-qualities-relateddesignconstants were studiedwith respect to t h e i r impact on thedesign.

7. Key Words (Suggested by Author(s)) 18. Distribution Statement

Conceptualdesignof a i r c r a f t Unclassified - Unlimited

Constrainedoptimization Nonlinearprograming Emnan ics Flying qualities Relaxed static s t a b i l i t y S t a b i l i t y augmentationsystems SubjectCategory 08

I

3. Security Classif. (of this report) 21. No. of Pages 20. Security Classif. (of this page) 22. Price A0 2 27 unclassified Unclassieied For sale by the National Technical Information Servlce. Sprinefield. Virglnia 22161 NASA-Langley, 1980