REPORT NO. AD-200 CONCEPTUAL DESIGN OF SINGLE TURBOFAN

ENGINE POWERED LIGHT AIRCRAFT CONTRACT NO. NAS2-9243

(BbSA-C6-151973) CDNCBPTUAL D E S X G N OP P N77-~30'30 SINGL3 T U a B O P A N X N G I N S POHERFC LTGBT

% F.IRCRIIPT (Cessna lircraf t Co.) 139 p HC iA37/MF A01 CSCL 0 1 C UiicLas G3/'35 28abt)

NASA CR-151978

CESSNA AIRCRAFT COMPANY PAWNEE DIVISIIQN

WICHITA, KANSAS

'".- ..., ' . . p a l

MODEL: 1603-1A REPORT NO: MI-200 _CUNCEPTUAL DESIGN OF SINGLE N l l B O Y N l

ENGINE POWEREI) LlGWl' AIRCRAFT Contrnct No. NAS2-9243

REPORT DATE: 31. March, 1977 -

PREPARED BY: Michael Newinn, C~orge Iluggina

CHECKED BY:

APPROVED BY:

f & ~ & P ~ . @ ~ ~ $ ~ $ ~ ~ , ~ ~ ~ ~ $ ~ l ~ ~ & 5 Y $ * & w ~ E { ~ ~ f J 7 ~ #

CESSNA AIRCRAFT COMPANY, PAWNEE DIVISION, W lC)ilT A, KANSAS

CONCEPTUAL DESIGN OF SINGLE TUSDUFAN E N G m POWERED LIGHT AIRCRAFT

By Michael Newan and George L. Huggia~

31 March 1977

Diotribution of this report i e provided in the lnrexest o f information exchange. Rerponeibllity for the contento raiderr

in the author or organization that prepared St.

Prepared under Con tract No. NAS2-9243 Ceusna Aircraft Co, Pawnee Division Wichita, Kansae

for

AMES RESEARCH CENTER

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

TABLE OF CONTENTS

1.0 Abatract 1 2 0 Introduction 2 3 , O Dircueaion 3

3 .1 Phase I - Market Analysis 3 3,l.l Design Requiremants 3 3 . 1 . 2 Ca&ting Methodology 11

3 . 1 . 2 . 1 Direct Operating Cost 11 3 , 1 , 2 . 2 Airplane Price 11

3.2 Phaea 11 - BaeeZine Def in i t ion 15 3 .2 .1 Configutation Layout 15

3 . 2 . 1 , l Design Conaiderations 15 3 . 2 . 1 , 2 Final Design 23

3 + 2 , 2 Liz: 2 9 3.2.2.1 Lift Methodology 2 9 3.2 .2 .2 Evaluation of Baseline 3 1 3.2.2.3 Juetfflcation 3 1

3.2.3 Drag 3 6 3.2.3.1 Drag Methodology 38 3.2.3.2 Evaluation of Baseline 38

3+2.4 Weights 42 3 , 2 . 4 . 1 Weight Estimating Methodology 4 2 3.2.4.2 Evaluation of BaJeline 43

3.2,s Tail Size 5 0 3.2.5.1 Tail Size Methodology 50 3.2,s . 2 Evaluation of Baseline 5 1

3 .2 ,6 Performance 55 3.2.6.1 Performance Methodology 5 5 3.2,6.2 Performance Justi fication-'Sake-off F i e l d 5 6

Lane th 4.0 Results 5 0

4 . 1 Garrett Design 58 4 Evaluation by Cessna Methodology and Comparison 58

to GASP Results 4.1.1.1 Lift 4.1.1.2 Drag 4.1.1.3 Weight 4.1.1.4 Performance

4 . 1 . 2 Comments on the Garrett Design 4.1.2.1 Wing, Engine, Gear Design 4.1.2.2 Tai l Design

4 . 2 Phase III - Wing Optimization for Base l ine 4.2.1 By Cessna Methodology 4.2.2 Wing Optimization by GASP

4 . 3 Phase Ilr - Engine Optimization for Baseline 4 . 3 . 0 heconc i l l i a t ion oE Methods 4 . 3 . 1 By Cessna Methodology 4.3.2 Engine Optimization by GASP

5 , 0 Phase V - Conc luaions 5 . 1 Flying Qualities of Baseline 5 . 2 Cornpariaon of Baealine t o Reciprocating Engine Powered

Light Aircraft 5 .3 Evaluation of GAqP

5.3 .1 Methodology 5.3 .1 .1 Lift 5 , 3 , 1 . 2 Drag 5 ,3 .1 .3 Weighta 5 . 3 . 1 , 4 Tail S i z e 5 . 3 . 1 . 5 Price and Direct Operating Coat 5 , 3 , 1 . 6 Fuselage Description 5.3.1*7 Performance Method6

5.3.2 Input Format and Elsa of Uae 5.3.3 Further anera1 Comrnentm end Rec~endatioae

5.0 Idenrificatiotr of Deeigu md Technology Raquitemnt~ of Sirule Turbofan Powered Linht Mrcraft

6.0 References 7.0 Appendix I - Direct Operating Coat Equationb 8.0 Errata

8.1 Engtne Price in GASP Analysis 8.2 TOFL Sizing Requirement8 8.3 Evaluation of TORL for Garrett Design

TITLE - PACE -

Garrett 4-Place U t i l i t y Deofgn 6 1 Garrett-NASA 2-Place Design Study-Aerodynamics 6 2 Garrett-NASA 4-Place Design Study-Aerodynamics 6 3 Comparison of Estimated Drag Levels f o r ~srrctt'r 64 Configurat ions. GarrettoNASA 2-Place Design Study-Performance 66 Garrett-NASA 4-Place Design Study-Performance 6 7 Tail Size Analysio 7 0 Schema t i c of Cessna ' 8 method of Determining Payload-Range 73 a t Maximum Range Speed, Wing Sizing, Cessna Merhodology 74 Wing Geometry Op t imfea t ion ,~es sna Methodology 76 Effect of Aspect Ratio on Opera t Lng Economics-Final Analysis 77 Ffnal S iz ing Plot f o r the 2-Place Trainer. From Cesena's 80 Sizing Program. Fina l Baseline Deeign. 2-Place Trainer 8 1 CG Envelope of F i n a l 2-Place Trainer 82 Drag Level f o r F i n a l Sizing Study. 2-Place Basel ine 83 Flap6 Up St ill Speed Versus Thrust and Wing Area, Plane 86 sized f o r 3.5 h r , Endurance with a Cruise Speed of 102 k t a , From the GASP Analysis . Takeoff Fie ld Length for a 90F Hot Day Versus Thrust and 87 Wing Area. Plane sized for 3,5 h r . Endurance with a Cruiee Speed of 102 kts, from the GASP Analysis, Direct Operating Cost Versus Thrust and Wing Area. Plane 88 Sized for 3.5 n r . Endurance with a Crufee Speed of 102 kts. From the GASP Analys i s . S iz ing Plot f o r the 2-Place Trainer. Cruise Speed = 102 89 knote, Enduranccm3.5 hours. From t h e GASP Analys is , Range at SW*lSO ft2, Thrust-700 ib SLS Versus Groas Weight 90 and Cruise Speed, From the GASP Analysis. Final Sizing Plot f o r the 2-Place Trainer. From the GASP 9 1 Analysis , Correct ion Fac to r s 112 The Influence of Aspect Ratio on C, for Straight Wings. 113

%Ax The influence of Aspect Ratio on C for Swept Wings + 114 Wing Weight Trend Used in GASP

41Ax 117

Corre la t ion of Horizontal T a i l Volume Coe f f i c i en t with 118 Fuselage-Wing Parametere.

LIST OF TABLES

PAGE

Garrett Design Concepts Single Turbofan Engine Powered Light: Aircraft Wing Geometry Matrix Weight Distribution Evaluation of the Weight Statement of the Garrett 2-Place Trainer Wing Geometry Optimization Reeults Comparison of the 2-Place Turbofan Powered Trainer to a Reciprocating Powered Light Aircraft, Comparison of the Components of the Direct Operating Cost Landing Gear Weights Fl ight Control Surface Actuation Systems Data

ef equiv

SYMBOLS

Aapsct ratio

*Average acceleration during takeoff ground run

Average accsleration during tromi tion f roar liftoff to rtrady rtrta climb

Conatante in weight trend equation

Wing rpan

Drag coaf f i c i e n t

Parsoiro drag coefficient

f nduced drag coef f iciene

Drag coefficient in ground ef f ecr

Equivalent friction coefficient

Lift coefficient

Mmdmum l i f t coefficient

Mean geometric chord of wing

Mean geometric chord of stabilizer

Mean gewetric chord of vertical

Sectian lift coefficient

Section maximum lift: coefficient

Oawald efficiency factor

EguivaXent f la t p l a t e area

Acceleration due to gt8vity

viii

Haximurn depth of furlrga

Coaoturtr hi weight aquation

L i f t

tloeixmt&l tdl arm (.25 cW to .25 5) Vertical tail arm (.25 rw t o .25 & ) Mach number

Derign lord factar

Reynolde number

Ground distance cwered during takeoff ground tun

Ground dim tanca covered during trrnritiorr from X i f tof f t o rteady state climb

Ground dietmce covered during climb from and of trim- e i t i o n to 50 ft. above runway

Horizontal ta i l area

Vertical tail area

W i n g area

Wetted area of cam,.l~te aircraft

Engine temperature' ratio (proportional to throttle eetr ingr)

Thrust, during takeoff, a t the average velocity for the takeoff eegmenc being calculated

Thrust, during takeoff, a t the climb rpeed, V50-l.ZVStall

Airfoil thlckneer ratio (rubacripto r-root, t-tip, svg-average)

$,:red

L i f t off rprnd

Wiad apamd

Climb out @geed over 50 ft , obrtrcls

Stall rpred

Hotieontal t a i l valumr coefficiant

Vertical t a i l v o l m c o e f f i c i ~ n t

Takeoff groem weight

Wing weight

Fuumlagt wsighc

Pmerplanf welght

Nacelle weight

Horizontal tail weight

Vertical tail weight

W n gear weight

Nose gear weight

Catrol system weight

Equipment weight

Furniehjnga weight

Inlet duct weight

Tai lp ipe weight

a-sg

BEW

DEW

DOC

Garrett

GASP

LFL

LSE

MGC

SLS or SLST

TOFL

Dry empty weight rmtnur urgiaer , propa, rviorrfie8, bat taw, a l t~nutor , and rmgulrtor

Maximum width of furrlrge Centar o f grwity range (is aft cg limit lninur fwd cg limit)

Tapar rati o

Rolling friction corf ficirrrt

Slope of mway

ASS drnrity

climb angle

Average

Basic Empty Weight - DEW plua optionaltqufpment a d all liquids except uueable fuel

Dry Empty Weight

Direct Operatin3 Cost

Refere t o ASResearch lfanufacturlng Company of Arizona, a division of The Garrest Corporation

General Aviation Synthesis Progra~~

Landing Field Length

Lighr Singla Engine

Mean Geometric Chord ( s w as yw) Sea Level Sta t ic lkruet

Takeoff Field Length

ThU rogart prrranto the rraultu ~f r atudy by Ccratrr Aircraft Cob ia which thr NA6A drvm2oped knrrrl Aviation Byntharir Frobram (GASP) w a r avrluatmd rr t o itr usetulnem II a d88im tool. Thia mvalurtlorr w u rcc- plbhad byr

1) Apprririnu tha resulfa of C6L 137944 vhich ware obtainad by the Gar- rett Corp. umiag the GASP,

2) conduct$ng a concaptual rtudy of a Cerann derigoed turbofan ria- craft wing Caamrlr .icing routiner and thr GASP and

3) comparing the GASP mathodology to the damipn ptocmdurrr now in ura by Cerrna, Thie evaluation concluded that the GASP nosdr extencive modif ica t ion6 to f ul- f i l l itr purpooe; but once there r ra audu thn proarll~ could be a wefu2 new tool for gaaarnl uviatim.

Ib a corollary t o thie evaluation $ha grser.~r;rlity of uring turbofan8 mi powerplanta for general avSrt$cm afr;te anggne aircraft war inveetig.ted, The conchaion ir that thw will be prcrctical if and when large rrductioru in tht coat end furl conrvrpcion of sm1P turbofanr are rchievad.

Thio rapor t p r reur t s the r e r u l t e of a etudy by Cerrua Aircraft Co. under NASA contract NAS2-9243 ~;ponruted by the Amer Reaearch Crater, The purparaa of the rtudy warat

1) To determine ths v a l i d i t y of t h e r a e u l t e obtained by AiUeeearch Manufacturing Company of Ari~ona (Divillon of Garrat t Corp. ) ueing the GenrraX Aviation Syuthesie Program (GASP) i n i.he derigu of a l i g h t turbo- f la pomrsd grnera l av ia t ion c laer airplane. (Ref. 1)

2) To conduct a preliminary design of a r imi la r a i r c r a f t bwed on Ceorna experience eud maly~ed both by GASP and by nethodn cl;:iently i n ura by Cerrna.

3) To evaluate the uaefulnere of the GASP a8 a d a i g n tool and compare the methodology of t h i r program to the daragn procedure6 now i n use by Cerena. Further, t o r e c m a d modificationr of the GASP t o increaee i t r w a f u l n ~ s a .

4) Ta tdent i fy rcaearch tha t mumt be done before the turbofan engine c m ba utP1ir;ad e f fec t ive ly on l i g h t , mingle engine general avia t ion a i r c r a f t .

In MA-Amea RFP2-26167, (Ref. 4) which i n i t i a t e d t h i a etud,r, the con- tractor wam required to choora onry one of the a i r c r a f t type6 considered In Ref. 1 for analycri~. Hmever, i n an e f f o r t t o msrlre t h i a design exercine cunfom t o a c t u a l darign pract icer , Cemena chose t o investigate both a 2- place t r a i n e r and a 4-place u t i l i t y a i r c r a f t . The approach was t o allow minor campromiree i n the derPgn of each i n order t o maximize cnwna: i ty between them. Thie is standard practice when i n i t i a t i n g a new l i n e of a i r craf t . Otherwiee, in t roduct ion of addi t ional new model8 2e neeidleesly I expaw ive .

. .

A market analyeir wae conducted t o detcxmine the deeign spi:cificatione far both of theee planes. Additional consideratione t h a t influenced the de- r ign of ~ e e e n a ' e baealine conf iguratfon were aee thecics, craehworehinene , rafe ty of operation, allowance f o r configurerion grcwrh, and l m technica l r i r k t o minimize development t h e , e f f o r t and caet.

The turbofan engine data wae supplied by NASA ae epeci f ied by t h e con- t rac t . Several d i f f i c u l t i e s , however, were encountered. F i r e t , the i a e t a l - lation Xaeees, included i n the datn, were determined f o r a podded engine inora l l e t ion such a s ahown i n Ref, 1 (see Fig. 28). However, the Cerrena design incorporated a buried engine (see Fig. 12) , for which tlbs iuetalla- t i on loeeea were expected to be eigniffcant3: d i f f e ren t . Eaeed on a Ccssna analy8ie and with the concurrence of NASA, an addi t ional 6 percen'r, peqalty i n thrurt walr charged agaitwt the eupplied engine data to account: for chis i a r t a l l a t i o n . It is of course recognized t h a t the Garrett engine would require radeoign to mrtch the engine nozzle cha rac re r i e t i ce to the duct design.

Second, t h e Ceeena deeign methodology required a l a r g e r range of t h r o t t l e ee t r ings than were supplied i n the engine data. Further , Ceesna iound.8 need

for data to M.ch numbers rlightly above . 3 a Only for the engine eiecd for Catrett'r 2-place plane were data a t theue Mach number6 ouppliad. It w u porrible to extrapolate the 2-place engine data for the necersary throttle ratting# but i t wan not poaribla to extrapolate the 4-plaza engine data to the required Mach nrrmbsre, There data could not be eupplfed to Ceerna ia t b for thir plane t o be evaluated, Therefore, the 4-place deaign was not riead for ltr miarion, The ef fec t of dropping the 4-place deofgtl ie miniawl, In a l l plrrticulars, except size, i t would have duplicated the 2-place air- plane, The procedurra followed in the design would not have changed nor would the data trends that reeulted, When the deeign wee corapromieed to a l l w atretching the 2-place alrcrafe to the 4-place configuration, the major objective of chaoeltlg to look at both piaues war achieved*

Ths baric criterion ueed by Ceeana fo r eizing them aircraft woe mini- mum DOC. T4ir of couree required a knowledge of engine price. N A W drcirPon, this war eet at $25/lb. of SLS thruat.

No effort wee mads to evaluate the approach and landing performance of thia drrign. The engine characteri~tica supplied to '~esena were for an en- gine optimized for a different ine tollation. ' ryefore, the f l ight idle charactrrierice (ill defined evett for the engine supplied) could not relaeon- ably be ured to evaluaerr the approach flight path charactetiatice o f the Cerrae derign. There ia no reaeon to believe that the landing requirements would be coaatraining on this deeign.

Thr darign requirements, incorporated i n the Garrett deoign n tudy , l a l d down in Ref. 4 md r h m in Tabla XI are neither adequate t o def bue the aieplrum nor completely realietic. Therefore, Cesena undertook a study to dtfinei neu rat of derign requirtmanta.

ThU rtudy took the form o f a cmparieon of the parfarnuance parsmeters o f preaent day rircrafe in the Z-plrce trainer and bplace LSE cacegorier. Thue compubanr ale rhavn oh Fig. * r L and 2.

The planer rhavn in the campari~an were choren rince, from their nccept- a c e , i t ir felt that they rport ideally meet the mrket for which they are .bed,

The parameterr r h m (takeoff field length, eta11 speed, etc.) were adjudged to be the m e t important. Nowever, it 8hould be noted that they ere not a l l equally corurtrainiag on the deefgn of the aircraft,

Bud on these graphs upd on the experience af the thein8 deeigu pereon- mT, ths requioementr in Table 2 wets derived.

The range requirementr for these aircraft were developed from coneider- atioar of eadurmce and cruba mpesde It la felt that 3.5 hrs i e a comfortable flight tirat fo r tha 2 - p h airctsf t deelgrrad far the training role but ueed In rhe trampartation made. For the 4-place u t i l i t y aitcreft four houra f l ight time w u aonuidsred comfortable. The design range for each aircraft h darb;dned for there f l i ~ h t timer at the cruise speed for saximum rnge .

Cruise npasd ir not chosen but i r r one of the deoign parameters that are . varied to muiaPite the range while mhimtzing DOC*

The cruira altitude w u r a t higher than the typical reciprocatlag engine powered aircraft to take artvantage of the jet1# lower power lapse rate with altitude and the Inherently higher true: airerpeed~ available.

-.

It i a imperative tbrrt a trainer be able t o f l y and laad without f laps . Landing ie .-st alwayr the moat d i f f i c u l t lessm far the etudenr t o - . grrrp. Therefore, i t SI umential that the procedures be kept &a s imple a8 porr5ble. If the tdngr are optimized for cruiae &ad the landing perfonmaace *a obtained with a pwerful f lap eyetem wing large defl.ection6, then the Plapr becorae necemsarp for the landiq. The apeed spread with flaps becornea large md tht f lyim qurlitieu vary eSgnificent2y from f lapa up t o f laps d m . Furthermore, once ret far 1and8&g, it irr then necessary t o a t least partially retract the flap8 in order t o execute a go-around. Although student8 a r ~ trained to manage the flrapa, i t i a daoiruble that early a010

f l i M t s be made with l ittle i f any flepr and that uriamanagement not reeult i n an unrecoverable incident. Otharwioe, the pottrnr$al fo r a etudenr mkirag a mirtake i r too great to tolerate* The flapr up eta11 epeed must, therefore, be met at 8 rar8onable level. ThwD eta11 rpeed directly impacts on the wing design rather than iadirectly through the flap configuration. AEI a corol- lazy, thin phlloraphy remulca in a riarple f lap myatem, a plana that c m be configured much that i t handles well during a crorrwlad landing, and a plane that i r flexible enough to perfom adequately a variety of m i ~ o i o n a i n addition to the brric training function,

Garrett, i n Ref, 1, a c t the take-off f i e l d length a t 2000 ft. Caeena feeble that th i s i s excaeeivalg long. Many airportij, preeently open to train- ere could not be uoed by a plane with this field length. Therefore, e f i e l d lsllgth of 1790 f t. on a hot day ( 9 0 9 ) was choran.

TABLE 1

GARRETT DESIGN CONCEPTS

2-Seat 4-Seat Trainer -

U t i l i t y

Design Payload (lba) 400 Maximum Payload (lbe) 400

Design Cruise Speed (mph) 125 Design Crulee Altitude (f t) 7500 Deeign Range (6 .m. ) 400

Single Turbofan Engine

NASA GA(W)-1 Wing Airfoil Section Ful l Span Fowler Flaps

Spollezs for Lateral Control

High Aspect Ratio Wing

Conventional Materials and Structural Design

Fixed Landing Gear

Standard Equipment Weight Includes Dual Controls and IFR Insrrumente and Avionics

Firaxe 1, Cornparfsm of Perfomaace Parameters of Present Day dircrrft. 2-Place Trainers. r%

G C -

-

Figure 2, mari son of Performance Par.metera of Present Day Lllrdt. 4-P- ISE, . - -

-

e~~yicrnuc~~~,rnu OF nl Qlth!~MAlk PAGE 18 PIHI1

TABU 2

S'INGte TURBOFAN ENGINE POWEBED LIGHT AIRCRAFT

Cantract NAS2-9243

2-P'he Trainer

Payload (including op tiorul ewiwmt*) 400 lb , Endutanca (no tenerveu) 3& hre, (to be re-evaluated If t i p tanka

are neceseaty to hold the required fuel volrtprs)

C d a e Altitude 10,600 fr. Takeoff Field Length t o SO ft, (SmL.@90'F) 1700 f t , Stall Speed (bf=O) SO Ktr.

Paylond (including op tiom1 equipment*) 800 lb. Endurance (no reeerves) 4 hxe. (to be re-evaluated Sf tip tanka

are nacerrrary to hold the required fuel volume)

Cruise Altitude 10,000 ft.

Takeoff Field Length to 50 ft. (S*~m@90'P) 1700 f t e

Stall Speed (Bf=O) 55 K t s .

Service ceiling, crube apeed and landkg field length are not expected t o be

eon6traSdug on the daefgn. They w i l l , of course, be monitored to make certah

that the rerulc;iog value8 are reelistic.

* NILVCOM and Transponder

mated and rwaned with the pricer of the following item@: enginer, rvionica, battery, altsrnaloro, oad regulator8 . Airf tunc price i n emtimated uring a fomulr derived from a correlation of a l l Cerrna aingle engine airframe pricar, m d shown in Fig. 4. Engine price for th i s study b,: been dictated by NASA to $25*00 par pound of rea level eta t ic ~ h r u r t , Avionicr, battery, alternator, ant: regulator pricer are for thooc uecd on rhe Ceasnr 150 C o w mutar 11, Aviaaica cork i r an item that i r not amenable to correlation with other factorr (much a8 alrcraft price), The price i s calculared from m itmized 1 S r t o f epecific squipmcnt, taken to be that typically inr ta l l ed in aircraft eimilar t o ti.: 3eeign configuration, In the caer o f the baeelint, thir war choaen, becaure of its popularity, to be that equipment inrtalled in the Conrmuterr 11 vcreion o f the Cevaaa 150. (Early i n the study thir list w r u originally taken from the Cardinal 11, When cmpheais wae ohifted to the 2- place trainer, the weight and coet eetimatea were revired t o t c f l e c t the avion- i c r of the Commuter XZ.1 The avionics coetr are ehom below bared on 1976 retail prices.

Commuter 11 Cardinal 11 ( w / M V PAC) Baeic Avionics Kit 175 Antenna and Coupler 5 5 300 Nav Corn (360 Channel) 1895 300 Nav Corn (720 Channel) 2295 300 ADF - 300 Traneponder

Figure 3. LUrfranre Maiateaance Cost

2.04Y ~ M P Z Y WC1&HT MINUS C N 6 f N E , P m 4 &ICTAICAL SYITfiM

Figure 41 Mrframe Coet

PHASE 11 BASELINE DEFINITION

3.2.1 Caof imratton tryout

Ilr Sn a11 cwcrpcurl derigu lryoute the configuration o f ~emmn'r dauign for a light curbofrn pwmred general aviation aircraft involved aaveral con- prornirn, The firm of there war whether to try m uaconvcntlonrl dtrign in m ef for t to improvr the ganeral aviatfon rtrte-of-thr-art in aircraft dcaign or to purrua a ccnvrntionrl lryoue. Bevarrl unconvmtionrl d e r i ~ a r are r h m on Fig, 5. Thrra hnva potanrial for reducing aircraft weight and increarlng parfomnncr . Wevar , the fal lwing murt alas be coneiderrd f

1) Unueual P l y i a ~ Qual i t ier - h the cror of the Cafllerr derlgnn the lr,ndfng attitude would of necerrity be high and it wottld be difficult t o incorporate flapr; therefore, theee derigna would not be adequate for the trainer applicrtion. In fact, the f ly ing qualities of a taillerr alrcxsft are no differmc that a p i l o t trained to f l y i n one would require further treinin8 t o f l y r convsntioaal configurntion,

la the cam of the canard, tht interig+rcnce between a canard and the via8 can cauee vety poor handling qualiel~% ., empecially near the r tall, un- lerr the w i q and canard are very carefully daeiwad and tatlored to each orhsr, Further, the t i p mounted vertice~r are rubject to large yaw-roll coupling of control inputa. The complexity of eheee deeign tmks and r i ~ k r are hard to overrtate and the deeign coet can be expected to be high.

2) Development Coete - Fox each of these deeignr new technology w i l l probably be needed in the avioaiclr area. Such technology daeo not presently exiet for the msnufacru+ar of general aviat ion aircraft. Thie w i l l be re- f lected irp high develop~lent comta and increamed aircraft price,

3) Market A c c e p t ~ b i l i t y - Unlike the military or space markets, wherc parformrrnce i e everything, the general aviation marker muse cater to the desireo of the coarrumer, Am tho a1:tomobile Induetry ham proven over and over, thir marker i e not open to radical changee i n styling and deeign. Rather, i t xsrrponda bcrt to elow evolutionary changes.

4) Purpoea of Contract - The vety pUrpO8e of ehia etudy was t o compare the GASP to the Cesena declign methode. Since these methode are calibtatcd against canventionnl deuignr, they are inadequate to handle much radically d i f f erenr concep t e and the credibi l i ty of the deaired GASP critique would be questionable. Por there rcaeone, the unconventional derigna were not pur- eued.

For the conventional design (i.e. horizontal, v e r t i c a l tail8 behind the wing mounted on fueelage a+ boom, conventional cabin, c t c * ) three main de- cirionm were coaftontedr

1) Wing Locatiz - high, mid or lov, Thla decision must coneider the

clffcetr of the lncrtion m v l ~ i h i l l ~ y , PmpennnlF and aft f r ~ m r l n ~ r rlerrtun, rprrr ( m i m 0 , hn ig t t at111 a1 l&t*ltt~iuttl f t t w l i t h $11 fttnmla&m) ant1 ottfilhta l w . ' m t 1 4 ~ ~ \ l b ~ i s l routing, ou t5et design, end interf ctence with main #par design).

2) Ennine location - the decision on where t o place the engine muet ac- count for the impact on center of gravity travel, power induced trim changerr, high velocity-high temperature exhaurrt flow f islda and on marksting conridere- rtiom of r e ~ t h e t i c e ~

With rtciprocatitlg powered aircraft, the engine watght f a ured t o balance the t a i l and aft fuselage weight, This allow6 the payload t o be located around the wing-body neutral pofint which i n turn minimitee the center of gravity mhift between heavy and light lordinq. With the j e t , however, the parerplab muet b e located sf t o f the wing (or podded on the win$) t o avoid sxcemriveXy long exhauet ductr, Therefore, the payload i e the only weight left to place ahead of the wing to balance the rai l , furelage, and engine and

' there i s r large c.g. ahfi t between heavy and light payloads. This i s o f par- ticular concern in the trainar where the eeudenr f i r s t leame t o f l y with the inotruccor aboard; which means the plane in at i c s forward c.g, On the firat: solo, haprever, with the inetrurtor out, the student is faced for the first time with a= af t c.g. condition. It i e , therefore, imperative that the cngine be located as close t o the wing-body neutral point an possible in order t o minimize thie effect_,The effect; of engine placement: on c.g, t s v e l can be i l lwtrated by comparing the estimated loading of the turbofan powered trainer to that of the Cessnb 150, Thie i a shown i n Sketch (a) below, Note that the turbofan loading i e all forward of the minimum flying weight while the loading of the 150 can be either faward or a f l of ihe minimum flying weight depending an placement of the np tional equipment.

Sketch (a)

SIDE-INLET DESIGNS

Figure 6. Design Conceptr. Configuretiom Featuring S l d e Inlets.

Dwign Concepts NOSEIINLET DESIGNS

F i w e 7. Design Concepts, Configur.rSm Faaturiog Nora Xnletr .

OVERHEAD-INLET DESIGNS

Figure 8. Desia Concep tr . Coaf igurationm Featuring Overhand Inlete .

Owign Conaept~ - FINAL DESIGNS CONSlDERED

Figure 9. D s r i g n Concepts. Final Damigue Conaidered.

TABLE 3

WING GEC4fErnY umut Taper tatlo .5, ,75, 1.00 Thicheea r a t i o (root-tip) .L5-e32,.17-.13,.19-.LS k p e c t ratio 6, 9 , 12

After rrslectislg the. optimum wSng geometry, the wing area and engine thruet v .re varied t o give the minimum DOC plane. reeults of these etudiee ate found in Secttons 3.2.2 throu~h 3.2.6 and 4.2.1 through 4.2*2.

Figure 10, Baneline Deeign Of ~esena's 2-Place Trainer. PreUmhw Conflgurad~.

CESSNA 4-PLACE UTILITY DESIGN

Figure 11. Baselins Design of ~essna's 4-Place LSE. Preliminary Configuration.

FtGURE 12 ENGINE INSTALLATION FOR CESSNA'S 2-PLACE f RAWER

F ~GuRE I3 ENGINE INSTALLATION FOR CESSNA'S 4- PLACE LSE

3*2*2.1 Lift b t h 0 d 0 1 0 ~

The ving l i f t sad murhum l i f t comfficirnt are effected by taper: ratio and upret ratio and by the valua and diotribution of thm dihrdral, thick- nrrr ratio, twirt and airfoil rrctlonr,

Effortr to account for thama affrcte by purely q i r i c a l marnm have been made* However, thSI h u p r m n impractical due to the ohurr number o f porriLblr varlntiom of the demggn parrmatmrr and the mutur. intmractioar of their effectr ,

Therefore, I t i m nacrrreary to w e an axact; method, bared on theoxet.lca1 conrrid~raeiom, that can account for arch of thsae effectr. The approaclr to dacarmtriag C ha trrkaa hate, i e the cmbinatfoa of a l i f t ing l i n e thaoty to drtermins #pan lording with two-dbneional wind tunnel data for the chonen 4irf0il ractione. The l i ft ing line theory u6re Fourier: a ~ l y s i r to deraminr the eprnwier l i f t dirtrlbutiun for an arbitrary apport~onment of chord, mist, dihedral, en8 airfoil on a rtraight wing configuration.

Maximum wing l i f t l a datermined by u r h g input maximum l i f t : ar a fvtnc- tioa o f Reynolds number, for the aixfoil rectiona along the w i n g span (for eat- ampla Ice Fig, 14). A t a particular wing Reynold6 number the rection maximum l i f t diettibution acrore the entire span of the wing, baaed on the appropriate local Reyuolde numbar, is determined. Then by a convergent iteration echam the angle o f attack a t which the wing aectlou l i f t dberibution becomee tan- @ant to the ractlon maximum l i f t dietributiaa i e found. A t thir angle tha rpnuwioe loadin8 is integrated to derenine the viag C*.

The remult (eea Fig. 20 for example) of thie proceas can alao be wed t o evaluate the atall chnracteriecics. R e k . 14 suggeete aa a criterion for good rtal l charactarirticr, thmt a c l margin of at lemt .15 at 70% eeniepan be ~ l n t a i n e d .

Figure 1 4 , Variation Of 1 ' ! u n Section Lift Coefficient With Reynolds N d e r For Various Airfoils. I = 0.15; MrfoUa Smmth.

fa order to evaluate the influence of wing paranatere upon the maa~iR3um l i f t coefficient, tho aapltct ratio, tapor ratio, &rid thicknear ratio ware ryrtcmutlcalLy varied over the aurtxix outlined in Table 3 of rection 3.2.1.2. Figure 15 nhowr that i n general, with conatant wing area, us the aepect ratio lncrruan the mnx- l i f r coat ficient decree8~6, Thio i r due to the fact that tha prana chord, und therefore the Reylroldr number, ir decreasing with hcrrroriug arpect ratio, An n h m i n Fig, 14 s decreaaa in Reynold8 nutuber decrsura the rection ~IPPm ljlf t available.

The influence o f t h i c h e s r ratio on i s illuetrated k Fig. 16. The twwdimenaional trcnde for maximum l i f t coeff ic imt ahown in Fig, 17 are, re-

in three d h ~ ~ i o ~ a d tha C h decreaaea w i t h increming thickauo ratio,

WiDg taper ratio inf1ueup.e on Ch ie shown I n Pis.18. Trende IoS thla trada art not ro ear i ly delineated ae w i t h AR and thickness, The effeet.of caper ra t io on C

=lux ia -re pronounced at low mpect ratio than a t high,.and for the aurrrix ~c&died the C LMAX hau a d m u m value between taper ratio@ o f ,5 to .75.

' A t rubaonic epaedr the wing C LMAX' for high aepact rat io winge, l a dp.rectly related to the maximum l i f r of the wGg airfoi l eectioa, with pfanfom ggoaretry of recoadary importance.

The Ceesnr maximum l i f t ~ 1 ~ l t h o 3 0 1 0 ~ has been jwtified by comparing the reeultr of wind tunnel teets and f l ight test data t o value6 predicted by. the l i f t routinen, Fig's. 19 and 20 rapraent the maximum lift caefficient solu- tion for two wind tunnel modelel The; dashed line in these figures reprewnte the dietributiun of eaction maximum l i f t along t h ~ . aedepan and the sol id line repoeeente the l i f t diotrlbution. For thsee two caees the maxiawn error in the calculated reeulta is 3,672, Similar reeulte have been produced for f u l l e&ale aircraft with the maximum error ueuelly being le8r thm 5%.

17-*- 1 6 a to 12

ASPgCT RATIO Fiwre 15. Effect Of Aspect Watio Oo @ Constant W i n s Area.

Figure 16. Effect Of Thielmeoo Rat io On C h @ Comtant W i u Ate..

Figure 170 Effect Of Thicknee8 Ratio On Sectim Haximum Lift Coefficient Cl SEX

f MAXIMUM LIFT COEFf /CtENT O/S rR/BUT/ON

0 20 40 64 80 id0 PERCENT SEMISPAN

b J

WING Lwr C Q ~ W / C I E N ~ DtSrRIBUTION

.

- c~MW ca/ = 1.45

Ckmx mew = 1.43 (WINQ TUNNEL) Re= 0.733~ 10'

Figure 19. Comparieon Of Calculated And Measured Values Of C 4w(

1 1

d 20 4d 60 88 100 PERCENT SEMISPAN

Figure 20. Campariaon Of Calculated And Measured Value6 Of C Q ~

3.2.3.1 Drag Methodology

Airplma drag i m uerraaed to be divided into.two parts; equivalent para- s i t s drag and induced drag. Comprereibility drag i e aarumed t o be negligible whets cruire Mach no. ie lea8 than 0.5, The Cerena mathod for drag estlmrtioa incorporatern MTCW mnthodology (DATCOM section 4.5.3) f o r paras i te drag und urea an empirical method for estimating the W a l d efficiency factor, ''cttn

The b u i r for the p a r ~ i t e drug ea t imt ion method i e the rkln f r i c t i a n method of DATCOM. Once the akin f r ic t ion drag has been calculated it i 6 multiplied by an empirical correction factor t o get the total p a r m i t e drag of the airplane without landing gear, The gear drag eetimate i s then addnd t o get the t o t a l paraaite drag. The empirical correction factor accomto for a l l Interference coacributiom of the wing-fuselage, wing-nacelle, and a l l lnircellraeow drag item.

The empirical method fo r emtimaring the Omrld efficiency factor has been modified by Cesena to provide be t te r agreement with general a v i a t i m a i r c r a f t , Thir method, which accouutr for the varia t ion of paraaite drag with angle of attack a8 well am f o r the induced drag, meken "e" a function of the emtimated p a r m i t e drag and the arpect rat io. Typical vlrluar o f "e" are r h m i n Fig, 21.

The s iz ing program, using the input fuselage wetted area (which Is not varied by the program) and the sreaa of wings and t a l l , calculetes the air- c r a f t wetted arm. The equivalent parasite d r a ~ is then determined on the vetted .tea and the empirical correctiun factor dimcursed i n section 3.2,3.1.

For conventional a i r c r a f t r imilar to modele already i n production, the empirical correction factor i r known. l o r newer c m f i g ~ r . t i m ~ ( ~ u c h as the bameline) greater reliance is placed on the data i n Fig. 22 which ehows equivalent p a r m i t e area (flat pla te area) as a function of wettad area. On t h i r graph are rhom cleveral C ~ ~ U M producte, a Ccesna eatimate for the IJind- ecker Eagle (Included became of i t r extreme aerodynamic cleauline8e) and l i ne r of equivalent f r i c t i on coefficient (Cf ) The baeeline moot nearly

equiv matches Lhc Ceaana 177B in configuratian; i.e. fixed gear, canti lever w i x ; however there :Lo neither prop waeh nsr caoZing drag mrociated with a turbo- fan engine and therefore a l e v e l of C = .006 vaa choeea. The wetted

epulv area of the a i r c r a f t vi th the largeet wing area of the matrix t o be i a v e ~ t i - gated in tho eiziug routine (see eection 4.2.1 and 4.3.1) f a then calculated and the drag is csrimated for the chosen value o f C . Working beck,ward

fequiv frm thie, an empirical correction factor of 1.052 was obtained and wed as d i r c ~ a e d i n rectlon 3.2.3.1. Thir factor is w e d for a l l other planes ln

the m i ~ i n g u t r h a d rrrultr in higher lavala of Cf aqutv for. plrarr with

T&a Oowald ef ficieacy factor l o d e t a d n e d by the a i ~ i n g program frm the rrpect ratto md paramite drag.

Pimre 21. Effect of AB snd Parasite Drag on (llsuald'8 Efficicacy Flttor.

Figure 22. Equivalent Parasite Area

3.2.4,l Waight Emtimating Mathodology

Csrrnalr methodology for waight e r t imnt ia r i 6 bared on empirical relation- rhipa for a 13 campoaent bxeakduwn, Thia mrthod provider a r e a l i r t i c evalua- t ion o f the vatioua compmentm uming the rlveead aluminum framework coartruc- tion utiliatd i n tha febrtcation of prarent day sircraft, For banded a l d n u o r wing c o ~ ~ t r u c t i o n , da ta point8 were obtained from ~erana*s current bonded u- r d l i e r urd from aevar.1 Calranr daulgn proporale In which the bonding tech- niqw war conridered t o r mrambly.

Ai rc ra f t ueed t o datermine the equatione fo r weight amtimating rang0 from the 1600 pound Modal 150 to the 5500 pound Model 310.

3.2.4,l (a) Breakdown Of Component6 The cumpanant b r e a k d m w e d i n errtimating the dry empty weight i e am

follows : 1) Win- - Includor fuaelage carryrhru for erparo, attachment hardware,

and fairing r t r i p e . Xncludc~ u t r u t e and e t r u r f e i r inge on braced wing. 2) Funelage - Includes f u e l a g e ahe l l r r m c t u r e , floorboards, s e a t r e i l e ,

doorr, windowa, rupport bracketa and wing, t a i l , and gear attachment hardware. 3) Pawarplant h e t a l l a t $ o n - Includssr avkrything eupported by the engine

mount, Thir value is usually provided by the engine and accceaory manufacturere. 4) Nacelle - Three typee of nacelle6 were conaidered for the determination

of the nace l l e weight, The wide v a r i a t i o n i n aar.e.lle confSgurat2ons v i l l limit t h e accuracy o f the nace l l e weight pradic i t ion .

I For mingle engine a i r c r a f t the nace l l e weight conmists of t h e c w l i u g D cov1 rauueing hardware and engine mount.

I1 Bar multi-engine a i r c r a f t the nace l l e weight cans ie te of the cowling, firewall, nacelle f a i r i n g aft of the fitawall, and the engine mount.

IXI For emall jet engine a i r c r a f t t h e nace l l e weight coaaisto of the engine f a i r i n g (or i n l e t acoop) and the engine mount.

5) Vextical T a i l - Includee v e r t i c a l t a i l e t ruc tu re , attachment f i t t i n g e end hardware, trim tab, and rudder balance.

6) Horizontal T a i l - Includes h a r i ~ o n r a l t a i l e t ruc ture , attachment f i t - t ingr and hardware, t r i m devicer, and e levator balance,

7) Main Gear Aeeembliea - Includes shock absorber, s t r u t f a i r ing8 (if ap- p l i c a b l e ) , brakes, wheele, and f i t t i n g r .

8) No8a Gear hesembly (or ta i l gear aesembly) - Includes shock abeorber, f a i r ing (If applicable), rhimmy dampenillg, eteer ing linkage, wheel, and f i t - t ingr .

9) Retract ion System - fncludea ac tuatore , plumbing, emergency hand pump, se l ec to r s , valves and f l u i d ,

20) Control S y s t e ~ ~ - Includee f l i g h t and engine controls. 11) Equipment - Iucludes e l e c triaal equipment (bat tery , box, regulator,

wirlng, e t c , ) and standard a.inimum f l i g h t inatrumente. 12) Furniehinga - Includes deate, upholetery, sound proofing material, re-

rtraiat ayatem, and heat and v e n t i l a t i o n eyetem.

13) Mircrllrnewr Itmu - All i t snu not covared la tha major breakdam abwa and rp8cP.l to the chorrn d u i u n conflgurrtiba,

The dry ampty veight (DEW) ir the rum of the componantr l isted in chr ptrrviour ract ion, Tke b u i c empty weight (BeW) i r the mm o f the dry empty usight, emginr oil, rrnwrbla f u r l , and optional equipment a

The opt io ru l equipasat i u ur itemized lbt o f rpecif ic crqdynent, u ru r l ly takon t o be tha t rqu lpmnt typicr1I.y inrtallad in a i r c r a f t rimilar t o the de- r ign configuration, It l a not por r ib le t o adequately represent them irema with any trend equation, Thatafors, f o r the i n i t i a l b u e l i n e a i r c r a f t , the Cerrna Cardinal 1 I, with Nav Pac, w a r w a d for an optional equipment break- dawn tha t could be conmidared typica l of the 4-place buealina deeign, For e h p l i c i t y ' u malts t h i r waa a100 ured during the i n i t i a l work on the 2-place brrslinel

3.2.4.1 (c) Deriving Weight Trend E q u a t i m

The veight emtimating marhodo are ured to determine the b ~ i c empty weight for a range of grolre wdghtr and wiag eream, bracketing the ertimated valueo. From thir cpn be found a r e l a t i o w h i p (trend equation) bemema groer weight, wiag &*a, and empty weight. The equation ia of tk form:

BEW - k1 + k2 (W) Where kl - a1 + a2 (SW)

kt .) + a, (ly) Note tha t the equatione are lineal. Experience he.> ehown th ia t o be an ade- quate reprementation over a l imited groae weight ranae.

3.2.4.1 (d) Urrga O f The Weight Trmd Equation For t h e i n i t i a l braeline a i r r a f t 27 weight trend equation6 were derived

for the wing gaae.etry prstrix of section 3.2.1.2. Theme equatione were then used by the rieiag program t o deternine the f u e l capacity and baeic empty weight ae &+ore weight and wing area wete varied t o match the design requfre- menta. A deta i led u r a l y a b of one of there 27 cares i r ehown on the following pagea t o i l l u s t r a t e the procedure described above.

3.2.4,2 Evaluation O f Baseline

The following weight entimation of r,he major component6 of t he i n i t i a l bmerine a i r c r a f t w i l l illustrate the methodology described i n the previous eection 3.2.4.1. The I n i t i a l b a a e l h e input wed far the weight eetimate i a lie ted belowr

sw - 85.0 tt? (wine area) AR - 9.0 (aopect ratio) b.75 (taper ratio) t .17 (root wing t h i c h u s ) W - 1400 Iba. (gross weight) SV - 16.2 ft? (vcrt. t a i l area)

s~ - 26.6 f r? (horiz. t a i l area) n' 5.7 (derign load factor) SLST - 300 lbr . (sea leva1 rtatic thtua t) 3.2 ,4. 2 (a) Component Weight Earimaee

1. Wing (complete with carrythru) -6 .69 W, = 69 (B X 10 )

Where 0 - W n' SW (1,9a:BR -4 )

Note: Similar to K,L, Sandere graph in Ref. 5 . Uaed for cantilever wings only. 3 - 1400 ( 5 . 7 ) 85.0 (1.9 (9) -4) - 3 . 0 9 6 0 7 ~ 1 0 ~

1 + 11 (.17) WW - 6 9 ( 3 . 0 9 6 0 7 ) * ~ ~ m 150.5 ibs . (68.27 Kilogram)

2. Fuselage Wf us - .I1 W - .11(1400) - 154.0 Ibe. (69.85 Kg,)

3. Powerpl~nt Ins ta l la t ion W~~ - ,28713 SLST = .28713 (300) - 86.1 Ibe. (39.05 Kg.) Reference Garrett Dwg. SKP-17547

4. Nacelle (external inlet scoop and engine mount) W, - .046 StST - .046 (300) - 13.8 lbs . (6.26 Kg.)

5. Vertical Tail WV - 1.28 St,, = 1.28 (16.2) - 20,7 lbr. (9.39 Kg.)

6 . Horizontal Tail VH 1.21 W le2i - 1.1 [ 1 * 0 0 ] ~ ~ ~ ( 2 6 . 6 ) - - 26.1 iba. (11.97 Kg.)

3000 3000 7. Main Gear Assenbly

8. ~ o s e Gear Aeswbly Wng = ,006w-I-19 = ,006 (1400)+19 a 27 .4 Iba. (12 .43 Kg.

9. FieEreetion System (not applicable) 10. Control System (for l ight eingle engine aircraft)

W, 40 lbe, (18.14 Kg.)

11. Equipcnt We 3S+.OIW m 35 + 14 49.0 1b.a (22.23 & a )

12, Fumishiaga Wf = 40 (No. of occupmta) 40(2) r 80 lbe. (36.29 Kg,)

13. M i rc~lluneou~ Item Deterrained weight of materielr tor approximate phyoical ahape.

a) Englne krlat duct (~ntetnal) "id - 12 lbr, (5,44 Q.) (formed slum. ehser metal duct approx. 13" dl r , and approx. ISqC long and r ivets)

b) Engine tailpipe (Iata-1) w t ~

- 34.5 lbr. (13.65 Kg.) (atainlear eteel duct approx, 19" dia. and 80'' lo&

Dry mty Weright C Component weight - 759.0 Iba. (344.28 Kg.) 3.2.4.2 (b) Basic Empty Weight Eathate

Dry Empty Weiglre O i l (1 quart) 1.9 Unwable Fuel (1 gallon) 6.7 Typical optional Equipm&r CPioted below)

Typical Optional Equipment Breakdown Wheel, and brake fairing0 Cproe, Hor. and 51$, Light-ulap Tcw bar Vacuum ryatem Primary Ijxoup Overall paint True airepaed indicator (each) Heated p i t o t Dual conrrole Light-oourteey Static Air Aet. errurce ILT 300 Nav Pak Nav/Com

2nd Nav/Com A3F Waneponder

Total

3.2,4.2 (c ) Weight Trend Bquatlon

18.0 6.0

.2 1.6 4.2 5.9 7 *O .1 .9 4.1 1.0

5 1.8

14.7 10.2

8.4 k * 6 -

89.2 (40.46 Kg.)

The weight range relected for the weight trend equtitian was 1200 to 1f100 lbr . The erne weight eatimrrthg procedure rn ~hown in the previow section was used t o determine the follorhg basic empty weighte.

BLW - 811.6 Ibe, BEW 901 - 2 Ib.,

for 1200 Ib, airersft for 1600 l b . aircraft

.

BEW kl + k9 W 542.8 + ,224 W Thia analysie was repeated for a wing area of 100 ft2 giving:

BEW kl .). k2 W - 549.3 + m233 W

From th i s we can aolve for the "a" coefficients kl " "1 + 8 2 SW rn 505.97 + ,4333 SW k2 * 3 3 +all SW = .I73 + ,0006 S ~ J

The erne proceduti wati bead tn determine the weight trend cquatioae for the wiw geometry matrix in Table 3, aecrion 3.2.1.2. 3 , 2 , 4 . 2 (d) Weight Distribution

This eect ion is included to illustrate how the forward and aft loading pointa are selected to describe the moat extreme conditione poseible for the f l y ing aircraft,

The f i r s t point t o obtain for the weight dis tr ibut ion envelope w i l l be the lightest possible aircraft with minimu fuel and a pi lot , This ie the minimum flying weight and Pa defined by FAR 23.25(b), The minimum fuel for th is condit ion is the fuel necessary for one half hour of operation at rated maximum continuous power (FAR 23.25 (b) (4)).

The location of the components estimated i n sect ion 3.2.4.2. (a) and 3.2.1.2 (b) are determined from the three view drawing and the minimum fly- ing weight c m g m is compiled &s shown i n Table 4. Each frem of the optional equipment (section 3.2.4.2. (b)) muat be visualized on the three view and compiled to find the c.g. of the total, The fue l load c.g. is located ac- cordfng to where it would be structurally feasible to locate a wet c e l l in the wing structure,

By starting with the minimum flying weight and adding the most forward items, the foward loading points are determined u n t i l the grose weight i s reached as shown on the second page of Table 4. This same proceduie i s a l so wed for the aft loading points. These pointe are then plotted an Fig. 23 and muat fall within reasonable aerodynamic limits.

If the c m g . range ie unreasonable, the wing is usually re:.tl : ~ ; r r l f a 7 r~5Xf t the envelope. Since the t a i l arms also change with wing movere::' : . 1 2 sizing muat aleo be redone. The weight distributSon prograr r r . I., .,-!- 3 until a reasonable c . g . raa.ge is obtained. The f l i g h t enve,. -. .. s t . ': drawn to cover the loading points aa shown by the broken U.ne ' * . & . a s

I fEM WING PNCC CARRY-TMU) F u S E L ~ ~ E POWERPLANT ~ N S T L NACELLE VERTf CAL TA / I &=/~.zRY

r TABLE 4 WEf6Mr OtS7R/BUTtUN

, 6ROSswr=t4u0~8 ~ ~ m 8 5 ~ ' ARzO.0 NT

150- 5 154.0 86.1 13.8 20. 7

tuz.17 A8.75 ~ 6 C r 3~ 129- lbrbRJ6t=kS.r63.00 &X=JO** hi/ - g

/8/*56 /61.t?b 232.00 2UO*cW 32!%06

WoRlZ. 741L (sM=26.6 fi? M A W GEAR ASSEMBLY NOSE GEAR ASSEMBLY RErRACnUM SYSTEM CON~@UL SYSTEM QY/PMN~ &- SD~HEP} F UUN/SH/WS EN6fNE IN4 ET LTUCT ENGfhlE TAIL PIPE

DRY E M P r v WE~GHT

U ~ L (1 QT) UNUSABLE FUEL (GAL) MINIMUM FUEL p/LOS

M/N/MUM FL Y/NG WT

3/9-50 i84.00 68.00

/75.00 109-00 /40*0Q /92*00 278-oo

zzBc60 174-/O /74./0 140*00

1 26.4 64.6 2 7e4 N. 4-

40.0 49.0 80.0 /2.0 34-5

759. o

1.9 6.7

70.0 170.0

t00715

C 6. t X-MOMENT

8 435 1 886 / 863

7000 5 341

/ / 200 Z 304 9 591

434 I 166

I2181 23 860

X-MdMENT 2 7 325 24794

SAW

183.40

/ 75.4

4 9 975 2 760 6 728

1 b W T

fi9S

3355

759.0

1007.6

-

139 202

/ 76 789

--

TABLE 4 (CONZ) I WEf6MT OfSrR/RU710hd 1

FOR wA RO LOADING M/N/M OM FLYING WT FWD OPTIONAL EQL/PT COP/f UT 0clGGAGE FUEL TO GROSS

AF 7 LOADING M / M / M O M F L Y / N ~ W T I FW TO FULL (M~LIL) BAGGAGE COPILOT 70 GROSS

I

r W T

~ o u Z ~ 892

170.0 60.0 73.2

/ 00Z6 23SO 60, 0 93.4

X / X-MOMENT I W T

IOOZ6 /096.8 /266.8

/65.50 /4d00 / 72.00 /74. /O

/ 74.10

f76 789 12 / 76 23600 10 326 12 744

/76 789 4 / 6f 0

Z %-MUMPJT

t 7 6 789 /88 965 Z t 2 5 6 5

/ 72. 00 44C.00

4

/6 299 t68-31

175.46 17520 t

L

C.G.

175-4633-55 1 72.29

i326.8 /400.O

/OQZ6 /246..6 1366.6 /.QUO-0

16 320 /SO76

2 W f

222 885 235 629

176 789 218 399 228 7t9 24/ 795

/6Z80/292

-i

Figure 23. CG E'ave10~1.e of Prelidnary 2-Place Trainer Configuration.

3.2.5.2 T a i l Sizt Methodology

Traditionally advance design methods have e k e d the empennage baaed Ln trend equations which i n turn are based on aircraft i n the erne e ize , speed and performance category aa the plane being deetgned, The ueual correlating facrorr are: c.

L For the hor izonta l r a i l Wf l f

-

S w % For the vertical t a i l h: It

-

Ceeena has developed the fallawing trend equations baaed on general aviation airctaf t.

For the v e r t i c a l tail the correlation fac tor is as above. The data ueed i e a h m in Fig. 24 along with the teeulting curve f i t . The equation i a :

cl - hLf VV - .0025 + .435 if f o r ~ i n g l e engine aircraft

S~ b~ The vertical t a i l area does not Lrclude the dorsa l or v e n t r a l f i n areae.

The v e r t i c a l t a i l is aesumsd to end a t the hor izonta l fo r low ra i l deeigne. For mid and T-tail deelgne, the root chord of the vertical ia uaually taken t o be a line, p a r a l l e l to the fuselage reference plane, i n t e r sec t in8 the ra i l where tho trailing edge of the vertical mete the top of the fueelage.

The horizontal i e eized against the factor: 2

Wf Ax If

The trend with thia factor ie shown i n Fig. 29 for single engine aircraft wi th conventional elevator-etabilizex deeigns. The trend equation waa derived to pa88 through the top of the data in order t o be conservative.

Iucorporation of the cg range i n t o the t a i l s i z e equation i e an important eafeguard against optimizing the wing with too high an aspect ratio. High aspect r a t i o wings have emall MAC'@. The permiseable cg range for euch wings becomee very emall. (Or the t a i l become8 very large.) Such designs only be- come attractive when c d i n e d with artificfal etabilizatian system that allaw the cg range to go f u r t h e r back on the wing than is allowable with etable de- signer. Since euch Byatems are (at l e a s t a t preeent) impractical fo r general aviatioa, the wing chord muat be kept fai> ly large in order to have reaeonable cg ranges or t a i l s i z e r . The trend equation is:

m

- wf Ax If V.r ,455 + 2.325 - f o r e ingle engine aircraft with con-

h e further refinement i r a correlation of aft cg against:

which ie ahown in Fig. 26. The reaulcing equation i e :

v 2 'aft cg limit - .I938 H + .I6 .flf for v 1 4.8515 (STAB-ELEV DESX GNG) *f2 l( sw $

Sw 5

After the t a l l is sfzed againat the cg range, the aft cg limit l a calculated by thir equation. If the celculated af t cg does not match thir number, the wing i e mwed, or the cg range is increaesd r o that the actual a f t cg l iee within the limit calculated.

3.2.5.2 Rraluatitm of Baseline

These trend equatians are incorporated directly into the computer program that eiee Cesem'e aircraft. The tail atme and c.g. renge are not varied (f ,e. lHr%,$/% and h x F W do viaty) ar wing eize is changed.

The tukl az8ea thw fatlnd .re uead to eccurntely detendne the wetted area and the drag variation for the differant: wing areas considered in the s i z i n g mott ix , Thaae equatians ate a180 used in the determination of the weight trend 8quatiotm .ham in taecslon 3.2.4.1.

Figure 24. Correlatfon of Vertical Tail Volume Coefficient With fuselage-Wlng Geometrg Parameters.

1 I 0 StNGLE ENGINE A TWtN ENGINE -- GASP CURRELAt /ON - CESSNA CORRELAI/ON

d 4- 4 /

x #

I

4 0

/ d /

/ 4'

4 0

#

Figure 25. Correlaticsr. of Horizontal Tail Val-. Coefficient wit5 F-sr iage Wing Geometry Parameters.

Figure 26, Corre'atfun of Aft CG Llmit wirh Fuselage-Whg Geo~etry Parameters and Tail Volume: Coefficient.

3.2.6 Perf o r m a n s

3.2.6.1 Performance Methodology

3 . 2 . 6 . l ( a ) Takeoff Field Length The takeoff is considered as consist in^ of three ports: the ground run,

transition, and climb to 50 foot height, For preliminary design work it is convenient to assume an average acceleration so that Lhe equation for the ground run becomes

ti T 2

- V~~~ 1 1 2 ~ s 'AVG p c L c II + Where P - w - D~~

Upon integration, accounting or headwinds, the ground run becomcs

The ~ r e n a i ~ i b n region aceounts f o r the rotation of the f l i g h t path fram horizontal t a t h e clirnh gradient and acceleration t o climbout epeed. The trans1 t i on grorlnd d l e rancr 1 6 computed from

1 Which results in S = (VAVG - V W ) ( V 5 0 - VLOF)

a+,, - 2

with 'AVC 1 / 2 p S V A V C C ~ W

where = - - - 2

The rlimhout c o n s i s t s of a constant s p w d climb to 50 f t . r c s u l r i n l : i n n ;?round dis tance of

and

50 f t . Sc = ( \150 50 f t cosy) .------ -- -- V5-,SINY TANY r--T

v50 - 1/2PS(V5") - 1 i y = S I N ,

1 . w

I n tho C o s ~ n n methadologv V = 1. I V S t a l l and VS0 = 1 ' 2 VStall* 'I'he tt>ti \ l ground distance to clear a 50 f t , obstacle is:

3.2.6.1 (b) Me thodology - Range Computation The range of an a i rp lane can be ca lcula ted with conriderable accuracy

w i n g a time in tegra t ion method if the aerodynamics, m i g h t , engine character- i s t i c s and fue l are known. However, euch methoda involve coaeiderable c o w putatioaal expenme. To reduce thie expenee and eliminate the need to make a time in tegra t ion t o obtain the range f o r every payload or fuel weight deofred, a modified form of the Breguet range equation i a employed in the Ceesna uizing program t o obtain the payload range parformance. The Breguet range equation was derived with the aerunrptio~ that, 1) the climb dl1 be coneidered a6 cruiee, and 2) the rrpecd, L/D, and SFC are conetanr. Thcae reetrictive areump- tione lead t o an overeatinat ion of range. To cornpensace f o r this a correction hae been developed that reduces the i n i t i a l weight by 02 of the mfeeion f u e l , Thia correc t ion factor was obtained by cor re la t ion of range, estimated using Breguette equation, with that obtained ueing the time in tegra t ion method. I t has been found to apply t o both j e te and reciprocating enginee. This analysis was bneed on cruise a l t i t u d e 6 between 7500 and 10000 f t .

3.2.6.2 Performance J u e t i f i c a t i o n - Takeoff F i e l d Length

Cesena merhodology f o r eetimacing takeoff f i e l d length has been compared to the a c t u a l f i e l d length6 shown in the Pilotte Operating Iundbook for varioua aircraft. Fig. 27 ohms thie comparison for a .rf Ceeana mingle engine aircraft. The correlation is excellent; with variations of no more than 5Z indicated,

Figure 27, Tskeoff F i e l d Length Analysis,

4.0 - RESULTS

4elrl Evaluation By Cemma Mathodology And Conperism To GASP Rerultr

4 e l . l . l Li f t

The murimum l i f t of the Garrett derign (rec Fig, 28 and 29) war determined ueim the Comana amthod dencribad i n rection 3,2,2. Rerulta of thia method gave-a c tMAX of 1.36 and 8 full flap ( B f 35') maximum l l f f coefficient of 3.03, Thim grcdueter a a t a l l opced with fu l l flaps of 44.2 kt8 at lrea level. The GASP co.pursd perforunc. $ave a C h with full flap. o f 3.9* m d a correapondtlrg s ta l l speed of 38.9 ktr.

A cmparieoa of drags aa derived from ~ e s u n s ' a method and by Garrett uainp, the GASP 8 ~ 7 sham on Fig. ' 8 30 and 31, The Garrett drags are optid18 tic c o w pared to C ana's trtiarate, However, since Garrett d i d not urre the form factor dafeYLtr ia subroutine AERO thiu probably reflect8 their choice of in- put8 rather thaa any inherent feature in the GASP.

The Garrett valuee o f CD are converted into equivalent flat plate area 0

and plotted an Fig, 32 for campariaon t o the Ceeana eetimater. They ehow a Cf..Yiv

of .0045 - .0050, Xhia ia on the order of that level e~timaeed for the'windecker Eagle, which wae a cantilever wing, retroct8ble geat aircraft with an exceptionally clean surface skin b u i l t o f fiberglass, Conaidering that the Garrett plrrnee are fixed gear designs i t would not appear poaeible t o achieve them low velueo of C

fequiv ' For aa exaurp1,- the e f fec t of gear on drag, compare the 177B on Fig. 32

wi th the 177RG. The only difference in theoe models l e that the RG i e a rea- tractable gear version of the 1778.

The Garrett deaign wee evaluated by the equations of section 3.2.4. The reaulrs are ehown belm in Table S along wi th the GASP reeults for comperieon. For rhie pwt of the ntudy the Garrett engine s i ze , wing and t a i l areae, and groaa weight were used to determine the component weights.

Ae can be eeen the only two areaa that show a significant difference tizc the landing geat aseembly and the control Byatem. The GASP weight estimstes, i n both areas, ate very l ight compared to existing light aircraft components now in production.

4+1.1+4 Performance

The parfrrrmance of twa of the Catfete dsaignr is avaluated uaing Ceeena methodo in order t o compare reeulte directly with the GASP eolution, Fig, 'B 33 and 34 giver the renultm of th ie comparirroa. In thir illumtration when comparing adjacent bars, the bar that endm farther to the right i s cowidered moxe optlmiatic. For the case with Garrett input t o the GASP, the GASP solu- t ion csneisteutly gave more opth ie t i c vnluee than the Ceeana method. In moat caeee the difference i n the reeulte can be explained due t o the larger drag cumputed by the Ceeena method. However, the difference in takeoff dietance over 50 f t , i a more than clln be accounted for i n this manner, The Cerena advance deuign method for takeoff petfonnanca i e outlined and aubrtantiated i n aection 3.2.6.

n GARBELT' 2-PUCE 2aAINER DESIGN

Figure 28. Cure t t 2-Place T r a i n e r D e d g n

GARRETT I-PUCE UTILITY DESIGN

Figure 29. Garrett 4-Place Utility DuJga

Figure 31 AEUOOYM~MIEE

I 1

Figure 32, Cornparlean of Estimated Drag Levels for Carrert'e Canfiguratione.

ITW I

Wing

EVALUATION OF THE WEIGHT STATEMPINT OF THE GARRETT 2-PUCE TRAINER

CESSNA -

130.3

Nacelle 9.9

Vertical Tail (Projected) 13.6 1

Horizontal Tail (Projected) 13.1

Main Gear h a y s ,

N O ~ R Gear h a y 76.0

26.0

Co~~tt01 System 40.0

Equipment

Furnishings ::I:} 126.6

Empty Weight 595.1

O i l 1.9

Fuel To Gross

Groes Wef ght

GASP -

135.0

128,o

61*0*

7 *O

*Garrett eupplied value

Figure 33 PeQPoQMAdCE

Figure 3 1

POFOUMA#&E

4,l.Z Colamoatr On The Garrett Design 4.1.2,l Wing, Engine, Garr Deeign

In order to complete thicr c r i t i q u e , the following comment6 are ofll'ered, which deal with rome of the derrign featurea of the Garrett plane.

The untapcred wing plsnform may appear to offer real coat savings by reducing the number of different rib deaigna of a tapered plglrfam down t o one and allowing varlous other a impl i f ica t ionr i n f lap and aileron deeign. However, clone examination of the rpa r design will show t h a t impartant weight eavingr can be realized by using a modifled "I" shaped epar and by t ape r i ng i t aloag the epan. Since the rib6 attach t o the epar, the same r i b cnnnot be ueed along the entire apan regsrdlese of planform. Therefore, very l i t t l e real errvittga accrue and the better aerodynamic efficiency and aeethet ice of a ta~erad planform dictate its we.

The gear dorign ehown by Gar re t t would have to be excep t ioml ly a t i f f to keep from bottoming the fwelage during a hard landing, Landing with such a etiPf gear vould be both d i f f i c u l t and unpleasant,

The engine appears t o be located too f a r aft . Aa d b c w a e d in eection 3.2,l.l thir leads ta an unacceptably large cg t r a v e l between l i g h t aud heavy paylorde. In addit ion, the i n l e t i a behind the higir :ing; which maane t h a t the dieturbed flov of f the wing fuselage juncture, during a e t a l l , would probably be in~catcd by the en gin^ It eeema unlikely t h a t the engine could cantinue to run under euch circumetancee. Final ly, the exhauet of o jet alwaye e n t r a i m air into i t r e l f , S h c e thia exhauet paaeee between the V-tail units, the ent ra ined air induce8 a greater poeitive angle of s t t a c k on the t a i l as the thruet ie increared. Thiu indirect r h m t effect may be as la rge or l a rge r than the d i r e c t thruet pitching moment and w i l l be f e l t by the p i l o t as s large trim change with paver,

4,1,2.2 Tail Deaiga

The V-tail choeen by Garrett %a not uemlly ueed Ln genets1 aviation (with one notable exception), A t f i r s t i t appeare poeeible to save weight and wetted area by ueing this type t a i l . A cereful analysla of chf problem, however, shows that a h a t a l l of the aavings are i l l ueory - very l i t t l e Lf any area i e awed. In a preliminary deeipn, before wind tunnel data i e available, a good deeign practice ie to eize the V-tail to have t h e sane total area as the conventional tail it replseea, For thin reason alone G a r r e t t ' ~ t a i l appeare amell.

Secondly, the high aepect ratio wing uaed by Garre t t haa a very emall chord; and, therefore, a large cg range in percent chord. Cowequencly the p lane nee& a large tail, Since the GASP does not eetimate cg range nor m d i f y the t a i l aiae w i t h cg range, the t a i l again appears t o be small.

Usiug the Ceeena trend equations from section 5.2.5.1 t h e c a i l eize re- quirement for the Garrett plane6 was eetlmated, (An asamption wne made t h a t the enains could be moved to bring the aft cg within acceptable limits.) Theee

rreaa are r h m on Fis,r 30 and 31 of r e c t d u ~ 4.1.1.2. Schemeically, a c m pariron i a mbm i n Fig, 3s between the Garrett r iead t a i l (ahcded) and the Cerrp. drr iwd V-rail,

Obviously the Cemma V - t a l l i e 60 large and the angle between the sutfacee r o great, that I t io impracticrl. A rough e a t h a t e rhowed that the rolllng moment due to rudder, for example, would be rur large am the yawing moment,

A more remonable approach ir indicated on Fig, 35 i n the farm of a con- vmtioual horit;oncrl wirh twin verticals. Although the total vertical ta i l area of t h b dreiga i r larger than for a eltagla tail, there i e an equivalent or larger raving6 in the area o f the horirontal t a i l due to the end plating effect8 of the verticels, Thir deuiga prwidoa a more plaurible aolurion then the V-tail when, becawe of the Jet efflrsu, a conventicnul empennage i e h ptactical.

Figure 35. Tail Stsc -pin

4.2 PHASE 191 W I N O OPTMXUTIOW FOR MSULmE

The Crrmr 8 i t i u g wthod make# uae of rolatiomhips derived from tha ilc thodologima of l a c t &oar 3. 2.2 thru 3.2.5. Thela r~lationrhipr include BEW - f ( W, 4,,), tai l mica - f (wing area, winp chord, atc.) .ad drag f (uet- tad area.) Wetted area, of courrt, varier with the w i n g utd t a i l erean. Thrue oalatioarhipa are wed to evr lwta the influence of the deaign pataaetato on aitcraf t side.

The um3n s i t ing program readr i n r data r a t which conristr , in part, of a mrtrix o f denim palemterm, (such ar AR, wing area, enghc eire, etc,) Thie program then calle variaw uubroutiner that, for each poinc i n the matrix, i t e ra t e an the weight to obtain particular parformancam parameter (TOE% - 1700, VS 50 ktr. stc.) Thm perfomonce p e r ~ ~ t e r o are rcc by rhe deeiga require- ments (ree eoction 3.1.2.)

The routin-, fo r the m e t part, utilize mtrai8htforward, iterarioa and perfomanca methode, 6- explanation, however, 4s required for the oubrsutine that a l t e r tho plane for the payload-endurance,

For reciprocating powered a i r c ra f t , the maxbum crulee power i e set by conaideratiom of engine operating charactaristicrr at ih value betueeu 75 and 80 percent of the ratad horrepwer. For a given cruise altitude thie UP courso me tr the rpeed .

For turbofans no similar conatralnt exlate for aetting the engine thrwt. Bone logical rationale fo r uettitrg chruet 1iW6t, therefore, be developed. Log- ical ly , the p i lo t would lib to cruiee at the rpeed for maximu range and the decteion was made to net the t h r o t t l e to ggve t h b epeed.

In order t o accomplish thia,ona more i te ra t ion loop that ~paxbir;eo range i a requLred i n the payload-endurance ritixtg routine. Fig. 36 i l l u e trs tee the procedure wed in the Ceasae methodology to determine the a i rc raf t weight which will aacisfy the given payload-endurance. Fig ; 37 r b m a characteritstic aizing p l o t produced by the r l z iag progruin Pot the l igh t turbofan aircraft. A t e conetant eapect ratio, taper rat io , thickners rat io , and engine rating the minimum weight a i rcraf t that vi11 ratiefy all ~f the deeign requirement6 a m be determined from thir figure. The wing parornetera of u p e c t ratio, thickneee rat io , and tspar ra t io were qariad as sham i n Table 3, section 3.2.1.2 and some of the resulte are m h m in 'Zable 6. A t each of the eieed pointe the direct operating cost war calculated uring the Cerma direct operating coat uethods end a result a t oae taper ratio i r mhoun i n Fig. 38, Bared upon the resvl te of t h i r study a wing geometry uelection was mde.

I t ww found that the taper ratio had a negligible ef fec t upon both the weight and the cost and, therefore, baaed uu atabi l i ty and control require- menta (roll damping, s t a l l chetrcter i r t ics , d a m wurh gradlent, l a t e r a l eteb- i l i t y ) and aerthetic appeal tbe w i n g taper r a t io war met at 0.75.

From 8n avataga thicknam8 of .I3 to r l 5 the DOC remained vlrtwlt ly un- c-14, showing a rigm&fieauc rime only when t / ~ ) ~ ~ roma to .17. Thetafore

k was illurtrated in Pi$. 38 the Cearna methodm indicated that aa rspect satto w u dmarurd &a dlr rc t operating cort v u decrrwad in tha ranpa o f u- pact rrtior %nvrrtlgrted, Haw..rr mince tlura i s vary l i tt le variation in DOC from A b 4 to AR-8, the ammthrt5cl of thr vauclm and the interftrencr of thm cabin door vLth the wine at 1au.r u p u c t ratio@ verm conridered rod ra u p m c t ratio of 7.5 w u ralmctad.

Table 6 .ad Pig, 38 were ~enerated before m error wpr diecovered in one of Carraa'r e iz iag routinem. Such error effected only the variation of DOC with ospect ratio and the corrected CUNC i a indicated i n Fig. 39. Thia shw8 t h a t m i n h DOC (,3255 $/N Mi) occuro around AR-6.3 and the choice of an AR- 7.5 wing (DOC - .3263 $IN Mi) v u valid,

I

SET WE~GHT - SETnYTP rn * CRLCULA7E , @asin*s) RANGE *

&I

NO NO

L

Figure 36. Schmtic of Cesena's Kethod for Determining Payload-Range at Meximum Range Speed.

75 (6.97) to0 @. 29) 125 (r/.6f) /50(/3.94) I 7 5 @6-26) WiMG AREA -SN, ~~'(n73

Figure 37. Wing Sizing, Ceesna Methodologym

TABLE 6

WING GEOMETRY OPTXMIZATxm RESULTS

t l c (%I -

X -

root

Figure 38. Wing Geometry Optimization, Ces~na Methodology.

Figure 39. Effect of Aspect Ratio on OperatLng Economics- Final Analysis.

4.2.2. Wing Optimigation by GASP

When thie etudy wra origtmlly formulated, the plane called for duplicat- ing the wing geometry oprimleetion procedure done in the previow section by a similar analyeie uelug GASP.

GASP, as now written, is not capable o f directly optimizing cruise rpeed for maxinnmP raugc (eae eection 4.2.1). I t can, I n one run, analyze only one cruire epead. The proceun by which i t may be done l a outlined aad i l lustrated in section 4.3.2 where the engSne size ie optimized. Thie procedure i r ~ long and tedioue. In order to have a valid comparienn between the GASP and ~esena'o methods the ground rules muet be the came, I. . there is no poesible short cur euch ae picking a constant cruiae speed for :!le GASP anelyuis,

It was impractfcal from a time and budget standpoint t o use this procedure on the large wing geometry matrix that W 8 8 aualyeed i n eection 4 . 2 . 1 . There- fore, the GASP analysis was not done and the optimum wing geometry derived by ~easna's methodology in eettion 4.2.1 was used for the engine optimization analye is .

4.3 PHAGE XV - mCINE OPTIMIZATION FOR BAS-

4.3.0 Reconciliation af Method

In coqaring sizing renults obtained from the GASP t o thcse obtained from the Ceerna methodology, i t ahould be noted t h a t the sea l e v e l e t a t f c th rue t quoted by GASP ie an ltletalled l e v e l and that quoted by the Cesana method i e uniastel led. This difference rerul ta from the manner by which irmtallation lormes are treated by each progtm.

With the, wing geometry paramrere determined during the wing optimization, (section 4, z41) , a matrix of engine rat;inge (SLST) and wing areas vaa analy~td by the sizing program and the xesur ts were plot ted on the carpet p l o t of Fig. 40. This figure repreaentu a l l configuretiom that will ear i s fy the require- ment of 3.5 hourr endurance with 400 pound8 of payload. Crone plot ted i n th ie f igure a r e the eakeaf f field length requirement and a t a11 speed requirement and Iince of con8 tat direct operating cost, The canf igura t ion which ~ ~ t i r f iea 411 of the design requirement@ and hat the minimum d i r e c t operating cost l a represented by the in te reec t ion of the TOFL line and the VS line a t a weight of 1740 lb., a wing area of 137.3 ft2, and engine r a t i n g of 415 lb, The direct: opetmting coat a t th ie point is .332 $/am. Thie configuration, ah '1 i n Fig. 4 1 incorporatee ths optimum s ized wirlg and engine i n order to satie~y the derign requirements epeciffed i n Table 2 of sec t ion 3.1.1 a t the minimum direct operating coat.

Fig. 42 ehwe the loading envelope for this f ina l s ized airplane. The aft cg is defined by the minimum f ly ing weight and the entire useful load i e forward of thin. The loading envelope is defined between the forward cg of 17.6% and the af t cg of 32.3%.

The design payload o f 400 lb* is capable of being loaded with f u l l fuel (337 lb).

The drag levels ueed in this sizing etudy are shorn on Fig. 43.

*fncludea NAVCOH and transponder

AT THE INTERS E CTiQN OF THE WtNG AREA - 7HRUS T L /NES P A Y L O A D = J O O L B ( / ~ ~ ~ ), &NbUR4NCE=3.5 HR

97 R= xs, A=.75* $g)r=8 *# 3 a.

Figure 40. Final Sizing Plot for the 2-Place Trainer. Frm Cessna's Sizing Program,

jL MAC

Fipxre 42. IX; Envelope cf F i r s 1 2-FI;;ze Trainer,

Figure 43. Drag Level for Finel Slz ing Study. 2-Place Baseline

4.3.2 Engine Optimization by CASP

When generating the inputs into the GASP, certain fundemeatal decieioas were made, Primarily i t w a s decided that the purpose of t h b rtudy war not t o derive ia new s e t of inputs that would duplicate the reeulta of Ceeena's methode, but rather to evaluate the CASP ae i t standu and recommend changes i n the basic methodology of the program t o bring il; into agreement with at& ard des fgn practices,

In particular, Cessna did not try t o make the drag match the rerulte af out plethods. Rather the default values r ~ f the form factore in rubroutine AERO were ueed.

Since, ae was previously explained, it wa6 neceseary to expand the thruae mape t o aecomodate ~eeena'a analysis, these expanded reblee were also incorpo- rated into GASP.

As diecueeed earlier, i n ~lection 4.2 .2 , GASP i e not designed to determine the best cruise speed for nuximum range. In order to do this, and thereby match the capabil i ty of the CASP sized airplane to the Cessna sized airplane, a matrix of wing loadings and thruate was establiehed for a s e r i e s of cruise opeeds. The wing loadings were 10, 14, and 20 PSF. The thrust ratings were 400, 500, 600, 700, 800, and 900 lbs, SLST, the matrix being either 3x3 or 3x4 a t each speed and the matrix of thrust levele increasing with increneing speed. The range woe specified ae that which would give endurances of 2.8, 3 .5 and 4.2 hours (i.e, deeired endurance 2 20%). This was done for MACH numbers of ,16, .18, -20 and -23 (102, 115, 128, 147 kts respectively). Tile reau l t s were p lot tcd aa "carpet" p lo t s of s ta l l speed, Ton, DOC and gross wei&ht, vs. wing area and thrust. Examples of each of those, a t a cruise speed of 102 k t s (M=,16), are shown in Fig, 44, 4 5 , 46 and 47 respectively.

The reoults of the computer analysis d i d not y ie ld the wing areas shown. Theee lines are the results of incerpolntions between the w i n g areas that the program generated from the input wing loadings and the resultant gross weights.

From the plot of stall speed, thrust-wing area combfaations y i e ld ing 50 knore were found and this line plotted onto the gross weight plo t . (Fig. 47). Similarly, lines correspnnding t o a TOFL = 1700 f t and to constant levels of DOC were a160 added to tl;: -toss weight plat.

Nowhere, however, from them graphs is i t poseible co determine which speed yie lds the maximum range, It is neceseary t o ?.xiow t h i s in order t o compare the reeulte of GASP to those of the previous section.

T h i s is found by making, from carpete l i k e these, for each of the 9 com- binations of thrust = 500, 600, 700 lb and wing area = 100, 150, 200 f c2, a new carpet p l o t of range vs. gross weight and cruise speed. An example of chi8 at 700 lb thrust and 150 sq f t wing area is sham in Fig. 48. The dasrhed

line i o the 3.5 hour endurancm lina. The dotted line rhwa tha locua o f max- Imm rrnge valuer for each grorr woight. Tha intrreectioa o f theoe 2 l ine8 i m the ''complete rolutiar" pime (i.a. 3.5 hr endurrace, ~ X s p r u m range) a t that thrwt and wan8 area, Each of thaoe "complate rolution'' planer can then be raplottad .r new carpet of groar vaight vrraur thrust and w i n g area, Thia new plot, rrhm i n Fig, 49 differ6 from Fig. 47 in that the @peed ia varying from point to point, By judiciour aroma plot t ingD i t was porrible to add liner o f Vs - SO ktr, TOPL = 1700 b~ mi linar o f conatant DOC, The numbers intrackrtr by the interractian of the l iner of the catpat mrr the cruire cpeed in h a t @ . Aa can br racn, the mdnimum MX: point l ier below the 500 l b , thrurt l ine md the cruirr rperd 50 brlw the laweat MBch number actually run (M - ,I6), Tine and money had been exhamtad before the lower r idt of thir ~ t r i % could be f i l l e d in. Hemoar, became of the characterirtice peculiar to carpet plot6 i t ir porriblo to axtrapolate the rerultm with conoiderabla conf idrace,

Tke f h a l rolut ion plane i6 read Prom the graph sr 8 W - 1940 lb sw m 200 ftZ Doe - .40 $/d (@ depreciation period of '8 yeare) Thrwt - 435 Ib (SU) (463 l b uninutalled) Cnrlre Speed - 91 RlA8

The GASP war w e d to verify thi ixtrapolation. The verifying run i u r h m on pager 92 thru 103, Note that the engine price has been adjusted t o account for the proper hmta3lation loroee re dircwred i n section 4.3.0 and the depreciation perlCod ham been nwrdified from 8 years to 7.5 yeare for coneir- tency vith the Ceerna method. The DOC, u expected, i a higher than the ex- ttapolation would indicate due to rharc changea in angine coet and depreciation,

Figure 44. Flrpr up Still Speed V-tm mrcur -5 Wk-11 &em. Plawr Stimd for 3,s hr. Endur6mcc with 8 Crui.s S p r d of 102 kt., PKa the GASP larlpsm.

Figure 45. Takeoff P i d d Length for a 90' Elot D.p Vetam Th-t Ind Ufrrg Area. P l u m Sized for 3-5 br &durance wfth a Cruiee Speed of 102 kes. From tbe GASP Analyeis,

Figure 46, Direct Operathg Cort Vanum Thrrut and Wiog Arm, Plane S i t a d for 3.5 hr Endumnca with 8 Cruira Spaad of 102 l t t m . F ~ E Q the GASP Analymir.

l imn 47. Sir;- Plot for th. 2-Place Trsiorr, Cruira S p a d - IOd Imou, Gadurure 3 r 5 hourr, From the CABP Ilnrrlydr.

Figure 49. F i n d Sizing Plot for the 2-Place Trainer. From the W P Analyeis.

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GROSS WEIGHT a 1940r PASSENGERS 1. PLUS CREW OF I FUS EL bGE h !IF:" E LC

WEYT O A R E A DELVE P 4ZPECT R A Y I 0 A FA SPAN

T I P THtCKNESS W I U G LPADlNG W I Y G FUFl. VOL UMF

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PSF 45:3 CUFt

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F Y t l N r S NUYRFR = I. TYPE= 7 S W T Y WEIGHT= 1174. LBS . I(&X,C~?U~SF SPEED= 91. UYOTS CrlYSUllFR P 9 I t F = 37663, DnL. S ~ S I C PRICE= 3 O t d l r OOL-

ADD- EWlPlrEhsf COST= 7002- WLr

SUB-TOT1t MAUJFhCTU41 W CPST DEALER COST R A S t C PPICE

II F Z Z 5 l VQnGQAM 1NTpRQWT PLO PSY I S 071DOOOOB2125RCb PEG1 STEP CDNT AINEO 78BlCD5000800000

R hN GE= 289. N-4. R L M K F M L = 376, LBS BLOCK T I C = 3-456 tWS- FUEL P a T E = 16.2 GPH. TBF= 3D?H)- Has- HOUPS/INSP-= 390- HRSe V A Q l l P L F t O 5 T t0CM.flPI FIXED CUST ( 0 0 L / Y R I

FUEL +O l l 11-69 STORIGE 0 . INSP-+*LTYI 3 -33 INSURAWF 1626- (HULL 3.0PCT) OVEQHAUL DFS. l.Cb ~ F P R E C ~ ~ ~ I O Y 4011. 1 B~IP-ZU.PCTI OTCrfR i?- 0 DtHEU

c n ~ w $ : ~ O V E ~ E I O 3.~~7; FbA f a x 25-

16-29 TnT4L 5464. TOT4L UTtL tZATICIY IHQS/VQ 1 100- 200- 300 - 400, 500 r 800. TOTAL nPR .CfiFf lQnLIH91 70.93 43-61 31-53 79-95 27-22 23- t 2

S o 1 HLYIHG QUALZTIEB OB BASELINE Thm f iml drrign of the Csroar ba~s l ina i s , excapt Ear the powmrplant ,

a convantional altcraft. There ie, for mxrmple, no r e a m to incorporate full rput Fmler flap8 rince the wing ham been rized t o al .10~ landing with the flap0 up. Neither attr rpoilrtr naaded rince the flap8 do not occupy the antire wing trailing edge, nor $0 the wing o f uuch high aspect ratio that ailarono cannot be wad becatme of the large toreaonal wing wrpleats they produce.

In l ight of the above, normal drrSga proceduree can be expected t o yie ld ratisfactory f ly ing qualitiao, A detailed aaalyei6 of handling char- rctrrirtlcr then would not add materially to the value o f thin etudy and the time available was allocated to other t aka that appeared to offer Snfor- mtion vrre pertinent to the purpoee of the rtudy,

The tail uurfacar are r i t r d accotdiag to the trend epuationr o f Section 3.2,s.

The valuer wed in these equeriona aret

1, - 138 In x a f t cg - 32,29 X MAC

Bared on the trend equst$orur thie yieldag

x aft cg iimit - 32 ,46 x MAC

Table 7 clhwr a cmpariem between the chsracrarirticrr of the Cesuna base l ine a6 derigned by Ceaena'e method and by the GASP. There character iat ica are aleo compared t~ the Commuter 11 version of the 1976 Cerenr 150. (The av ion ic s package of t h e Comuter I1 l o s very popular opt ion and as much was choeen for the baealina.)

The gror8 weight of the GASP earlusate i r r r u b a t a n t i a l l y Iaraer than t h a t estimated by Ceeana'cr method. This ir +ue %n p a r t t o the higher f u e l conamption of the larger engine, (376 va, 33: pounde or a AW - 39) and i n part to the higher empty weight. (1174 ve, 1003 paundr or ti AW - 171). Of th i s latter f i g u r e 48 pounds can be a t t r i b u t e d t o the h igher groer weight and 3 pounds to the larger engine, The reminder (171-48-3-120) ir artrib- uted to the di f ferenct in methode,

The difference i n wing loading and engine s i z e between the Ceeane nized a i rp l ane and that sized by the GASP i a bel ieved t o be due, i n p a r t , t o the di f fe rence8 i n drag between the two methode. Comparison of Fig. 40 t o Fig, 49 ehowe a l x g e difference i n the slope of the TOFL - 1700 f t l i n e that is also ahif t ine the minimum DOC po in t (ae defined by the GASP) t o lower wing loadingn and lower thrust levele,

The b a ~ e l i n e plane (as estimated by ~esana'e method) is heavier than the Cesana 150 by 140 pounds. Yet due r o the lighter engine the baeeline enjoys a e l g a i f i c a n t l y lowar empty weight (1003 vs. 1122 poundr). Although the baseline carriee 47 pounds more payload, the difference i n grow weight ie moetly due to the higher f u e l load required (337 ve, 125 pounds).

Compared to the Ceesna 150 the trainer is a t a cons iderable economic disadvantage. The price i s from 38% t o 101% above the 1 5 0 ' ~ (depending on the method uaed to eetimtlte it). The DOC shms a 76X t o 119% i nc rea re over the 150, There i e no reason from t h i s comparison to believe tha t the turbo- fan powered plane could compete i n todloy'r market againat che t r a i n e r s cur- rently ava i l ab l e .

A breakdown of the DOC (aa ca l cu la t ed by ~eesna's s +.hod) is shown i n Table 8. This indicate8 that the turbofan is a p p r a x b t , l y equal t o the 150 in many areaa. However, i n the areae of engine periodic maintensnca, f u e l cost, and deprec ia t ion the turbofan aircraft i s considerably more expensive to operate (the fuel coat f o r the 150 was calculated for a f u e l price of $.75/gal; 1.e. 5~ per gal higher than the price of j e t fuel) , The reeul t , as mentioned above, is a DOC 76% higher than the 15Qvo.

Price $ Direct Operating Qost $/MI Weight (Iba)

Groes emps Fuel Pay load

COMPARISON OF THE 2-PLACE TURBOFAN POWERED TRAINER TO A RECIPROCATING POWERED LZGHT AIRCRAFT

Size Length Overall (it) Wing Area (f t 2 ) Tail Area (ft*)

Horizontal Var tical

Perf a m n c s Cruiea Speed (kts) Cruise Ale (ft) Endurance (hrr 1 Range (n.mi) Ton (ft) Stall Speed (kte) (flap8 up)

Engine (Unins t a l l e d )

2-Place Turbofan Trainer A6 Optimized by As Optimized Caer