Analysis of wind tunnel test results for a 9.39-per cent scale model of a VSTOL fighter/attack aircraft. Volume 2: Evaluation of prediction methodologies

3 1176 00 166 1322 [ASA CR-152391 NASA-CR-152391-VOL-2 ANALYSIS OF WIND TUNNEL TEST RESULTS FOR A 9.39-PER CENT SCALE MODEL OF A VSTOL FIGHTER / ATTACK AIRCRAFT VOLUME II - EVALUATION OF PREDICTION METHODOLOGIES DR. J. R. LUMM S G. T. JOYCE C. D. O'MALLEY / Prepared under Contract NAS2-I0344 by General Dynamics _ "_'_ _ _., ' _,-_-:7 Fort Worth Division _ - " ' : ' "'' " __ " for Ames Research Center ;; . ._._ LiF S _: , :,,_Y, i¢ ,1 : .-_ , % NATIONAL AERONAUTICS AND SPACE ADMINISTRATION _] s _locument c ontains T ech n icalDataconsi d ere dt o I_ I ' esourceunder ASPR1-329.1(b)and DoDDirective 5400.7 and is Rot a "record" required to be releasedutider the Freedomof InformationAct .

1. R _ ort No. 2. Go _ rnmemA m m J onNo. :L R _ m r , Cata _ l No.

NASA C R-15 239 1 4. Tit l e and Su _i tle 5. S eoml Dute A nal ysi s o f W i nd T unnel Te s t R esul t s f o r a Octobe r. 1980 9.39- p e r c e n t Scal e M o del o f a VSTOL F ig h t er / A tt ack 8 . p ,_o, , . _ o_ , i .. ti o, o_ x A i rcraf t 7. Auth o r(s) 8 . Perfo r ming O rgDni U t @ o . Repo r t No.

" D r. J. R. Lumm us , G. T. Joyce , U. D . O'Ma lley I O.W. k U n . No .

9 .P e rfo r ming O.mei_kmN_. e a. d _ -- Gener al D yna m i cs / For t Wor t h D ivi s io n 11 . =o , , , = _ G , ,, No.

P. O. Box 7 48 NAS 2 -10344 FortWor t h, Te x as 7610 1 1 3. T ,_ o f R wo. _ d _ . ,o_ Co_ 12. Soons _i ng , A _Btn c:y _ l id Addr lu Con t rac t or Fi nal Re por t Sep t . 10, 1979-Fe5. 10_ 1981 NASA, AMES Res e arch Cen t e r , 1 4 . s,x , . ,= , .i ._ * g.,,m, ca= .

Moffett Fiel d , C a 9403 5 1S, S u i) _ ementary Noum A M ES R e s ear c h C e n ter Te c h n i cal Mo n i to r W. P . N e lms ( 4 15) 9 6 5-588 0 IE . _a_ Th e r e sul t s o f a s e r i es o f N AS A A MES w i nd t u nnel te s t s o f a General D yn amics v ec tore d- e n gi n e - o ver win g. Na v y V S TOL f ig h te r / a tt ackc o nf iE ura tlo n - h a ve tee n an al yzed t o ( I ) a s s e s s p redi c t io n m e t hod cap a bi l ities, (2) evaluate geo m etry varia t ionssu c h a s mul t iple c anard lo n gitu d inal locations and strake shapes, and (3) evaluate the effe c ts of c onfiguration changes asso c iate d wi t h varying the p ro p ulsivelif t sys t em fro m a j et- diffuser e j e c tor to a Remote Augmen t ationLif t Syste m ( RA LS). C onf i gura- tion m odificationand addition a ltes t ing a nd analysis are re c o m mende d to adequatelyevalua t e the c onfigurationpotential.

This document is presented in fo u r volumes - Volume I - St u dy Overview,i Volume II - Evaluation of Predi c tionMetho d ologies,Volume III- Effe c ts of Config u rationVariations from Baseline E205 Config u rationon Aerodynamic Characteristi c s,and Volu m e IV - R A LS RIO4 Aerodynamic C hara c t eristics and Co m parisonswith E205 C onfig u ra t ionAerodynami cC haracteristi c s.

17. Key Words (S u gg e ited by Authm(s)) 18. D i m'il m n ic m S t ate m eflT CANARD , STR A KE, AE ROD YNAMI C PREDI C TION MET H ODS U N C LAS SIFIED UN C LAS SIFIED

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Fo r saleb ) lhe N a t ,onal T , 'chn _. alln f o r ma t ,on Sen s e . S p n n E field.V,qp m a 22161 i -i_ 5.

VOLUME II - EVALUATION OF PREDI C TI O N M E THODOLO G IES TABLE OF CONTENTS Predicted vs Test E205 Baseline Configuration Aerodynamic Characteristics

1.0 U ntrimmed Longitudinal Aerodynamics 1

I.i Aerodynamic Center vs Mach No. 1 1.2 Zero Lift Pitching Moment (Cmo) 2 vs Mach No.

i 1.3 Minimum Drag vs Mach No. 2 1.4 Power Off - Untri m med Lift, Drag, 6 and Pitching Moment 1.5 Wing Trailing Edge Flap Effectiveness 9 1.6 Buffe t Onset C haracteris t ics II r 2.0 Trimmed Longitudinal Aerodynamics 13

2.1 Trimmed Power-O ff Aerodynamics 13

2.2 Trimmed Power-On Aerodynamics 15 3.0 Lateral Directional Aerodynamics 18 3.1 Rigid Sideslip Derivatives, Cn_ , 18 Cl_, Cy_ 3.2 Vertical Tail Effectiveness 19

3.3 Aileron Effectiveness 19

4.0 References 21 iii LIST OF FIGURES

PAG E

FIGURE

i-i a Effect of Mach Number on Lift and Moment 22 b Effect of Mach Number on Drag, (Expanded 23 Drag Scale) c Effect of Mach Number on Drag 24 1-2 _ Effect of Mach Number on Lift and Moment 25 b Effect of Mach Number on Drag 26 1-3 Aerodynamic Center Test / Theory Correlation 27 1-4 Comparison of Predicted and Test Aerodynamic Center Variation with Mach Number 28 i-5 Comparison of Predicted and Test E205 Model Zero Lift Pitching Moment Coefficient (CM0 ), Variation with Mach Number 29 i-6 Comparison of Minimum Drag vs. Mach Number for E205 Model Test Data and Predicted Data 30 1-7 E205 Full-Scale Airplane Cross-Sectional Area Distribution 31 1-8 E205 Wind Tunnel Model Cross-Sectional A r ea Distribution 32 1-9 Comparison of Harris Prediction of A CoMIN Due to the Aft Sting Lines Modification on the E205 Model with Experimental Results for Other Configurations 33 i-i0 Effect of Adding Excess Interference Drag (Determined from VEO-Wing Fighter Model Data) to the Predicted E205 Wind Tunnel Model Minimum Drag 34 i-ii a Lift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteris- tics of Baseline E 205 Configuration , Power-Off, Mach = .2 35 b Drag Comparison of Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E 205 Configuration, (Expanded Drag Scale), Power-Off , Mach = .2 36 :v

LIST OF FIGURES(CONT'D.)

FIG URE PAGE

1-12 a Lift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteris- tics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +i0 ° , Power-Off, Mach= .2 37 b Drag Comparison of Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +i0 ° , Power- Off, Mach = .2 38 a Lift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteris- tics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +25 e, Power-Off, Mach = .2 39 b Drag Comparison of Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trail- ing-Edge Flap Deflected +25 ° , Power-Off , Mach = .2 A0 1-14 Comparison of Incremental Effects of Canard Deflection for Predicted and Test Data, Mach = .2 41 1-15 Comparison of Incremental Canard Effects for Predicted and Test Data , Mach = .2 42 1-16 Comparison of Incremental Effects of Canard Deflected -i0 ° for Predicted and Test Data , Mach = .2 43 1-17 Comparison of Incremental Effects of Canard Deflected -20 ° for Predicted and Test Data, Mach = •2 44 1-18 a Lift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteris- tics of Baseline E 205 Configuration , Power- Off, Mach= 1.2 45 b Drag Comparison of Predicted and Test Longi- tudinal Aerodynamic Characteristics of Base- line E 205 Configuration , Power-Off, Mach = 1.2 46 v

LIST OF FIGURES(C0NT'D.)

F IGURE PAG___EE

1-19 a Lift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteris- tics of Baseline E 205 Con f iguration , 47 Power-Off , Mach= 1.6 1-19 b Dr a g Comparison o f Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E 205 Con f iguration , Power-O f f, Mach = i.6 48 1-20 a Lift and Moment Data for Baseline E205 Con f iguration , Mach= 1.8 49 + b Drag Data for Baseline E205 Con f iguration, (Expanded Drag Scale), Mach = 1.8 5 0 c Drag Data for Baseline E205 Configuration , ++ Ma ch = 1 . 8 5 1 1-21 a Lift and Moment Data for Baseline E205 Configuration, Mach = 2.0 52 b Drag Data for Baseline E205 Configuration , (Expanded Drag Scale), Mach = 2.0 53 c Drag Data for Baseline E205 Configuration , Ma c h = 2. 0 54 1-22 a Effect of Wing Trailing-Edge Flap Deflection on Lift and Moment for Test and Predicted Data, Mach = .2 55 b Effect of Wing Trailing-Edge Flap Deflection on Drag for Test and Predicted Data, Mach = .2 56 1-23 a Lift and Moment Predicted Data with Canard Deflections and Wing Trailing-Edge Flap Deflected +i0 ° , Mach = .2 57 b Drag Predicted Data with Canard Deflections and Wing Trailing-Edge Flap Deflected +i0 ° , Mach = .2 58 1-24 a Lift and Moment Predicted Data with Canard Deflections and Wing Trailing-Edge Flap Undeflected , Mach = .2 59 v i

LISTOF FIG U RES(C0NT'D.)

FIGURE PAGE

1-24 b Drag Predicted Data with Canard Deflections and Wing Trailing-Edge Flap U ndeflected , Mach = •2 60 1-25 a Lift and Moment Predicted Data with Canard Deflections and Wing Trailing-Edge Flap Deflected +25° , Mach= .2 61 b Drag Predicted Data with Canard Deflections and Wing Trailing-Edge Flap Deflected +25 ° , + Mach = .2 62 1-26 T hrust Split Between VEO-Nozzles and Ejec- tors Required to Achieve Pitch-Trim as a Function of Trimmed-_ and VEO-Nozzle / Flap Deflections for Mach = .2, CTT07AL = 1.81 63 1--27 Incremental Effects of Canard Deflection with Wing Trailing-Edge Flap Deflected +i0 ° for Test and Predicted Data, Mach = .2 64 ol-28 Incremental Effects of Canard Deflection jJ with Wing Trailing-Edge Flap Deflected +25 ° for Test and Predicted Data, Mach = .2 65 1-29 a Effect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +i0°, Mach = .6 66 b Effect of Canard Deflection on Drag With Wing Trailing-Edge Flap Deflected +i0°, Mach = .6 67 1-30 a Effect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +i0 ° , Mach= .9 68 b Effect of Canard Deflection on Drag With Wing Trailing-Edge Flap Deflected +i0 °, Mach = .9 69 1-31 a Effect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +I0 °, Mach= 1.2 70 vii

LIST OF FIGUR E S(CONT_ * D.) '

F IGURE PAGE 1-31 b Effect of Canard Deflection on Drag With Wing Trailing-Edge Flap Deflected +i0 ° , Mach = 1 . 2 7 1 1 -32 a Effect of Canard Deflectio n on Lift a n d Moment With Wi n g Trailing-Edge Flap De flecte d +250 , Ma ch = .6 7 2 b Effect of Canard Deflection on Drag With Wing Traili n g-Ed g e Flap Defl e cted +25 °, Mach = .6 73 1-33 a E ff ect of Canard De f lection on Lift and Moment With Wing Trailing-Edge Flap Deflected +25" Mach= 9 74 I • b Ef f ect o f Canard De f lection on Drag With Wing Trailing- E dge Flap Deflected +25 " , Mach = .9 75 1-34 a Effect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +25 " , Mach = 1.2 76 b Effect of Canard Deflect i on on Dr a g With Wing Trailing- E dge Flap De f lected +25 ° ,Mach = 1 . 2 7 7 1- 3 5 Incremental E ff ects Due to Defl_cting the Wing Trailing-Edge Flaps +I0 ° wh l le in Presence of Various Canard Deflections for Baseline E205 Configuration , Mach = .6 78 1-36 Incremental Effects Due to Deflecting the Wing Trailing-Edge Flaps +i0 = while in Presence of Various Canard Deflections for Baseline E205 Configuration, Mach = .9 7 9 1-3 7 Incremental Effects Due to Deflecting the Wing Trailing-Edge Flaps +i0 ° while in Presence of Various Canard Deflections for Baseline E205 Configuration , Mach= 1.2 80 1- 38 I n c re m en t al E ff e ct s Due to D e fl e c ting th e Wing Trailing-Edge Flaps +25 ° while in Presence of Various Canard Deflections for -_ Baseline E205 Configuration, Mach = .6 81 viii

LIST OE FIG U RES(C0NT'D._

PA G E

FI G URE

1-39 Incremental Ef f ects Due to De f lecting the Wing Trailing-Edge Flaps +25 ° while in Presence of Various Canard Deflections for B a seline E20 5 Configuration , Ma c h = .9 82 1-40 Incremental E f fects Due to De f lecting the Wing Trailing-Edge Flaps +25 e while in Presence of Various Canard Deflections f o r Baseline E205 Configuration , Mach= 1.2 83 1-41 I ncremental E ff e c ts o f Trailing-Edge Flap Deflection With C anard Off, Mach= .6 84 1-42 In c remental Effe c ts o f Trailing-Edge Flap Deflection With Cana r d Off, Mach = .9 8 5 1-43 Incremental Effe c ts of Tr a iling-Edge Fla p De f lection With C anard Of f , Mach= 1.2 86 .i-44 Incremental Effects of Canard Defle c tion with Wing Trailing-Edge Flap Deflected +i0 ° , Mach = i.6 87 1-45 Incremental Effects of Canard Deflection with Wing Trailing-Edge Flap Deflected +i0 " , Mach = 2.0 8 8 1-46 Ef f ect o f Mach Number on Wing Buffet, Wing Trailing-Edge Flap and Canard Undeflected 89 1-47 Effect of Mach Number on Wing Buffet, Wing Trailing-Edge Flap and Canard Undeflected 90 1-48 Effect o f Canard Longitudinal Location on Wing Buffet, Mach= .6 91 1-49 Effect of Canard Longitudinal Location on Wing Buffet, Mach= .9 92 1-50 Effect of Canard Longitudinal Location on Wing Buffet, Mach= 1.2 93 1-51 Effect of Canard Deflection on Wing Buf f et, Mach= .6 94 1-52 Effect of Canard Deflection on Wing Buffet , Mach = .9 95 ix LIST OF FIGURES (CONT'D.)

FI GURE PAG E 1-53 E ff ect o f Canard De f le c tion on Wing Buf f et , Mach= 1.2 96 1-54 Co m parison o f Predicted- _R0 with _BR - Indicators CR M s , CN, and CA f rom Test Data 97 .i-55 Effect of Canard Deflection and Canard Location on GBR 98 X LIST OF FIGURES (CONT· D. )' FIGURE

PAGE

-

Power-on and Power-off Predicted Trimmed e's 2-1 as a Function of Equivalent Lift Coeffi- cient, ClE, Mach Number, and CIJ (from Reference 1) 99 2-2 Trimmed Lift and Drag with Wing Trailing- Edge Flap Deflections and Canard Undeflected, Mach = .6 100 2-3 Trimmed Lift and Drag with wing Trailing- Edge Flap Deflections and Canard Undeflected, Mach = .9 101 2-4 Trimmed Lift and Drag with Wing Trailing- Edge Flap Deflection and Canard Undeflected, Mach = 1.2 102 2-5 Comparison of Power-off Trimmed Lift Curves and Drag Polars for Trimming with Flap- Fixed, Canard Varies and with an Envelope of

Optimum Canard and Flap Deflections, Mach =

1.6 103 '2-6 Comparison of Power-off Trimmed Lift Curves and Drag Polars for Trimming with Flap- Fixed, Canard Varies and with an Envelope of Optimum Canard and Flap Deflections, Mach = 2.0 104 Trimmed Lift and Drag for Baseline E205 2-7 Configuration Using Canard and Trailing-Edge

Flap Deflections, Mach = .6 105

Trimmed Lift and Drag for Baseline E205 2-8 Configuration Using Canard and Trailing-Edge

Flap Deflections, Mach = .9 106

2-9 Development of Envelope Trimmed Lift Curve

and Drag Polar, Mach = 1.6 107

2-10 Development of Envelope Trimmed Lift Curve

and Drag Polar, Mach = 2.0 108

2-11 Comparison of Trimmed, Power-off LID vs a for Trimming with Canard Fixed, Varying Wing Trailing-Edge Flap with Optimum Combination of Canard and Wing Trailing-Edge Flap, Mach = .6 109 xi L IST O F FI G URES _ CO NT'D . ) FIGURE ?A__KG E 2-12 Comparison of Trimmed, Power-off L / D vs for Trimming with Canard Fixed , Varying Wing Trailing-Edge Flap and with Optimum Combination of Canard and Wing Trailing-Edge Flap , Mach = .9 Ii0 2-13 Full-Scale E205 Airplane Predicted, Power- on , Trimmed Lift Curve- and Drag Polar- Envelopes for M = .2, C TT 07A L = 1.81 III 2-14 Full-Scale E205 Airplane Power-on , Trimmed Lift Curve- and Drag Polar-Envelopes from Wind Tunnel Data for M = .2 , C TT OT A L= 1.81 112 2-15 Comparison of Full-Scale E205 Airplane Trimmed, Power-on Drag Polars from Predic- tion and Test Data , M = .6 , C_ = .302 (Optimum Canard / Flap Envelope Trim for Prediction) 113 2-16 Comparison of Full-Scale E205 Airplane -- Trimmed, Power-on Drag Polars from Pre- diction and Test Data , M = .9, C_ = .159 (Optimum Canard / Flap Envelope Tr l m for Prediction) 114 2-17 Full-Scale E205 Airplane Tr i mmed Dr a g Polars from Predictions and Test Data, M = 1.2 115 xii

LIST OF FIGURES6CONT_D.)

PAGE FI GURE ------ 3-1 Variations of Predicted and Test E205 Side Force Derivative. C¥_ . with Mach Number and_ 116 3-2 Variation of Predicted and Test E205 Yawing Moment Derivative. C , _ . with Mach Number and_ 117 3-3 Variation of Predicted and Test E205 Rolling Moment Derivative , C_ , with Mach Number and_ 118 3-4 E205 Predicted and Test Vertical Tail Effectiveness at Various Mach Numbers 119 3-5 E205 Predicted and Test Aileron Effectiveness 120 xiii VOLUME II LIST OF TABLES Table No. Title Page I.i E205 Wind Tunnel Model Minimum 4 Drag Buildup from E205 Airplane Minimum Drag Buildup 1-2 Matrix of Power-On and Power-Off 7 Trimmed and Untrimmed Comparisons of Predicted and Wind Tunnel Data for the Baseline E205 Wind Tunnel Model xiv LIST OF SYMBOLS a. English Symbols A axial force, Ib (N) • a.c. aerodynamic center, %_ AR aspect ratio b span, in. (m) _, MAC m ean aer o dynami c c h o rd, in . (m) CA axial force coefficient CA axial force coefficient due to ejector eject o r CD drag coefficient aero-only drag coefficient (no thrust increments CD A ERO included) CDmi n minimum drag coefficient CDE equivalent drag coefficient CDRAM ra m -drag coefficient (engine inlet) CDt total drag coefficient CL lift co eff ic ie n t • buffet-onset lift coefficient CLbuffe t CL_ " equivalent lift coefficient CLmax maximu m lift coefficient aero-only lift coefficient (no thrust increments CLaero included) total lift coefficient CLt Cl rolling moment coefficient xv LIST OF SYMBOLS (Continued) CI_ rolling moment derivative due to sideslip, i / deg CmE equivalent pitching moment coefficient Cmx - pitching moment coefficient about x percent c Cm zero lift pitching moment coefficient o Cm t o tal p i t c hi n g moment coefficient t CN normal force coefficient Cn yawing moment coefficient Cn_ yawing moment derivative due to sideslip, i / deg T CT thr us t co eff ici e n t, qSREF side force coefficient _ side force derivative due to sideslip, i / deg CMU, C ideal thrust coefficient, w Vj / gqSRE F D drag, ib(N) e span efficiency factor T ESF engine scale factor, TESF = 1.0 IGE in gr o und effect L lift, ib(N) Lp lift due to supercirculation, ib(N) I rolling moment, ft ib (Nm) M Mach number m pitching moment, ft ib(Nm) Total Pressure NPR nozzle pressure ratio, p xvi LIST OF SYMBOLS ( Co ntinued) N n o rmal f or ce, ibtkg) n yawing moment, ft ib (Nm) OGE out of ground effect "P freestream static pressure, Ib / ft 2 (--N 2) m Po freestream total pressure, Ib / ft 2, (--N 2)m q freestream dynamic pressure, Ib / ft 2 (--N 2) in SC canard exposed area, ft 2 (mz) Sre f reference area, ft2(m 2) (usually equal to SW) STOL short takeoff or landing SW area of trapezoidal wing extended to centerline, ft 2 (m2) SVT exposed area of vertical tail, ft2(m 2) T thrust, ib (N) Vo_ freestream velocity, ft / sec, knots (m / sec) V. jet velocity based on isentropic expansion from 3 nozzle camber total pressure to freestream static pressure, ft / sec (m / sec) VSTOL vertical or short takeoff or landing VTOL vertical takeoff or landing VEO-Wing vectored engine over wing w weight flow, Ib / sec (kg / sec) X action point of circulation lift relative to - cp leading edge of MAC xvii LIST OF SYMBOLS (Continued) b. Greek Symbols a alpha angle of attack, deg ._, beta angle of sideslip, deg F super c ir cu lation 7 flight path angle, deg _C' 6i canard deflection (positive, leading-edge up), deg _TE' _F VEO-Wing nozzle and outboard flaperon deflection, deg; except for aileron action the fZaperons and VEO-Wing nozzle flaps always deflect together.

pitch attitude angle, deg 83 jet thrust deflection out of VEO-Wing nozzles when deflected, TE' deg ALE leading-edge sweep angle, deg A taper ratio tip chord ' root chord eje c to r me a s u red t h rust / isentropic s u pp l y t h rust (where isentropic supply thrust is the thrust which would be obtained from supplied air at the nozzle exit of pressures and flow rates expanded at isentropically to ambient pressure) xviii LIST O F SYMB O LS ( Con ti nu e d ) c. M o del Symbols B I VSTOL ejector configuration E-20 5 basic fuselage with fuselage strake that blends the fuselage to the inboard section to the wing.

B 2 VSTOL RALS configuration R-104 basic fuselage CI All moveable nacelle-mounted horizontal canard of VSTOL ejector configuration E-205 in the mid- location C2 Horizontal canard H_ _ in VSTOL E-205 or PALS RI04 fwd-location C3 Horizontal canard in VSTOL E-205 or RALS RI04 aft-location N VST O L ejector configuration E - 2 0 5 o r RALS RI04 r VEO-wing nacelle SI Baseline strake on E205 configuration S2 High sweep strake on E205 configuration S3 Low sweep strake on E205 configuration V All moveable vertical tail of VSTOL ejector configuration E-205 or RALS RI04 W I VSTOL ejector configuration E-205 wing with linear elements between SS 96.496 and SS 223.695 W 2 VSTOL RALS configuration R-104 wing with linear elements between SS 87.231 and SS 214.430 xix X X VOLUME I I - EVA L UATION OF PREDICTION M E THODOL O G IE S One of the primary objectives of this investigation is an evaluation of current prediction methodologies to esti- mate the aerodynamic uncertainties identified in Reference 1 for the E205 configuration described in Volume I, Section 3.1. This evaluation was accomplished by comparing pre- dicted and wind tunnel test data in three major categories: untrimmed longitudinal aerodynamics , trimmed longitudinal aerodynamics , and lateral-directional aerodynamic character- istics.

1.0 U ntrimmed Longitudinal Aerodynamics Figures i-i through i-2 demonstrate the variation o f the baseline E205 wind tunnel model lift, drag, and pitching moment with angle of attacks and Mach number f or .2 < M < 1.2. Although these data will be analyzed in some detail in subsequent sections, they are presented here to provide an overview of the basic trends; that is , increasing lift slope ( i n the linear G range ) , increasing pos i tive stabilit y and increasing minimum drag with increasing Mach No. as ex- pected. The high angle-of-attack , M = .4 characteristics obtained in the low speed wind tunnel test (Figure i-i) agree well with the M = .4 data obtained in the t ransonic test (Figure 1-2).

1.1 Aerodynamic Center vs Mach No.

The aerodynamic-center (a.c. ) travel with Mach No. is a real driver in the E205 design. As explained in Reference i, the E205 configuration is longitudinally, statically un- stable to achieve the VEO-wing nozzle benefits. The pre- dicted instability levels are greater than can be presently tolerated. The maximum allowable instability dictated by control system limitations is approximately 15-18% MAC.

Therefore , the Flight Control System (FCS) will be used to augment the stability to the required level of frequency and damping. As part of this augmentation the flight control computer will be used to schedule the canard as a function of Mach number and angle of attack to achieve the desired level of static longitudinal stability. Obviously then , the aerodynamic-center t r avel is an important parameter that must be accurately predicted. Estimates of the E205 config- uration aerodynamic-center travel with Mach No. have been made by using the Carmichael Procedure (Reference 2) and the Datcom method (Reference 3). Figure 1-3 presents a General Dynamics a.c.-prediction-accuracy correlation for the Car- michael procedure for various configurations , including the VEO-wing fighter model of Reference 4. The correction vs Ma c h No. indi c a t ed for th e V E O - w i ng fi gh t er mo d el was ap- plied to the Carmichael predictions for the E205 baseline configuration (a similar configuration) to produce the corrected Carmichael estimates, which are compared with wind tunnel results in Figure 1-4 for a zero-degree canard de- flection and with canard o ff . The Datcom estimate f or canard at a zero-degree deflection for M = .4 is also shown for reference and shows a signi f icant disparity between the prediction methods.

It is very difficult to predict the E205 a.c. with either o f these existing methods because of the unusual aspects of the configuration: the wide , flat body with sep- arated nacelles, the relatively blunt forward strake , etc.

However, the Carmichael procedure plus the correlation- correction developed by General Dynamics did yield surpris- ingly good predictions at most subsonic and transonic Mach numbers. There are , however, some signi f i c ant discrepancies f o r Ma c h No. > 1.6 ( the pred ic t i ons a r e conserv a t i ve ) . The predicted trends and levels agree rather well with the test data below Mach NO. = 1.6 and the accuracy of the predic- tions in this speed regime is certainly satis f actory f or preliminary design purposes. For Mach No.'s _ 1.6, the pre- dictions are unsatisfactory.

1.2 Zero-Lift Pitching Moment Coefficient (Cmo) v s Ma c h NO. - - Figure 1-5 compares the test and predicted variations of zero-lift pitching moment with Mach No. for the baseline E205 configur a tion ( 6c = 0° ). ( The effects of removing the canards and wings as well as shifting the canards to alter- na t e longitudinal loca t ions are also indica t ed f rom the tes t data described in Volume III).

The estimated Cmo variation presented in Figure 1-5 for the baseline configuration is the direct result of the ex- perimental data bases used for predicting the aerody a mics of each speed regime for the full-scale aircraft (as des- cribed in Volume I , Section 3.2) because no prediction method per se is av a ilable to handle predicting the Cmo for a configur a tion like the E205 with the unusual combinations of body, strake, nacelle, and wing c a mber and the subsequent interference between the components. However , the Carmich- ael procedure would have probably provided some guidance to doing a better job of predicting Cm o . Figure 1-5 emphasizes the inability to predict the Cmo variation; further analysis of the canard and wing trailing-edge flap effectiveness and the resulting trim (Section 2.0) indicate the real impor- tance of being able to accurately predict and tailor the Cmo characteristics.

1.3 Minimum Drag vs Mach No. _ The estimated minimum drag variation with Mach No. for the E205 wind t unnel model baseline configuration (6c = 0 ° ) is compared with the wind tunnel results in Figure 1-6. The estimated model minimum drag was derived from t he estimated full-scale aircraft minimum drag according to the equation shown at the bottom of Table i-i. Table l-1 from Reference 1 has been modified to demonstrate how the estimated wind tunnel model minimum drag was developed from the full scale airplane minimum drag at various Mach Nos. by removing the increments for roughness and protuberance drag (because it's a "smooth" wind tunnel model), flap scrub drag (because the model is unpowered) , and the missiles and launcher drag (because the model, unlike the airplane , has no missiles and launchers); minimum drag corrections for Reynolds Number differences between the full-scale and model were also applied as well as corrections to the supersonic wave drag for differences between the full-scale airplane lines and the wind tunnel model lines. (Note the respective cross- sectional area distributions in Figures 1-7 and 1-8.) The inlet spill drag increments were determined experimentally for the E205 wind tunnel model as described in Volume I and added to the estimated model minimum drag variation with Mach No. , making possible the direct comparison of model es- timated and test data shown in Figure i-6.

This comparison indicates that the prediction methods described in Volume I, are very effective in the subsonic and transonic speed regimes. In fact , good agreement is achieved for M < 1.2. For 1.2 <Mach No. < 2.0 the prediction methods tend to underestimate t_e wind tunnel data. There are probably two causes for this low estimate: (i) uncertainty in the estimated increment due to the lines modifications for the aft sting and (2) uncertainty in the interference drag for this type of configuration.

Figure 1-9 shows a plot of the variation with Mach No.

of the increment in minimum drag due to the aft sting for several General Dynamics wind tunnel models; these incre- ments were experimentally determined by subtracting drag levels using alternate mounting methods. Note that the added cross-sectional area required for the sting installa- tion (relative to the actual airplane lines) results in a reduction in drag for all of the configurations tested except the VEO-fighter model (Figure 1-9). The complete VEO-fighter configuration results in positive drag incre- ments subsonically and varies from positive to negative to positive as Mach No. varies from 1.2 to 2.0.

The prediction method (the Harris procedure) employed to estimate the supersonic wave drag of the E205 wind tunnel model indicated a drag reduction for the aft sting instal- lation rather than the drag increase indicated (at some Mach numbers by the VEO-fighter configuration model) as shown in Table I-I E20 5 WIND TUNNEL M O DEL MINIMUM DR AG BU I LDUP FROM E205 AIRPLANE MINIMUM DRAG BUILDUP (SEE EQUATION (I) BELOW S ref = 38 4 ft 2 M A C H NUMB E R DRA G COMP O NENT . 2 .4 .6 . 8 .9 1.2 1. 6 1. 8 (1)rag i n C o u nt_ Fri c tion 166. 5 14 9 . 3 13 9 .0 1 3 0.4 1 2 6. 5 116 . 0 103.0 90. 8 Form 17. 2 15. 5 14. 2 1 3 .4 13.1 - - - In t erference 8.2 6.9 10.9 21.0 22.2 - - - Wing Camber 2.1 2.1 2.1 2 .1 6. 3 9.4 10.4 14. 6 Roughnes s + 25.9 25.9 25.9 25.9 25.9 32 .5 28.8 2 5.4 Prot u berance Flap Scrub 3 2.7 10.9 5.5 4.4 4.4 2 .2 I.I l.l W a v e ..... 292 . 3 289.4 28 1. 4 Mi s siles + Launcher s (2) W ing-Ti p LCLM 7. 7 7.7 7 . 7 7. 9 8 . 7 1 6 . 1 14.3 12 . 2 (2) NAC-MT'D AM R AAM 7.7 7.7 7.7 7.9 8.9 13.5 I I .0 7.5 T ot al C _. in F u ll S c al e 268 226 2 1 3 2 1 3 2 1 6 4 82 4 58 4 33 _ CDmjn Scal e / R E 9. 0 2 7.2 36 . 5 4 2 .1 44. 2 44.7 44.7 44.7 _ CDw . . . - .... 3 3 . 0 33 . 6 4 5 . 6 ( F u _ w sca l e to m od e } _ CDspill (Model ) 0 O 2 5. 0 8 .0 1 6.O 15 .O O T o tal CDmin M o de l 203 201 2 04. 7 21 4 22 0. 3 51 1 4 96 47 7 E q (1) PREDI C TED CDH IN HO DEL = CDH IN - A C D H IN - A CDH IN A CD NIN + A CDM IN + 4 C D M I N + A CD w Av E F ULI _ ROUG H + F LAP M I SSIL ES I SCALE / R E MODEL BET W EE N S CAL E FROT. S CRUB + AI RPL ANE L A U N CHE R S FULLSC A L E SP I LL MODEL AN DFULL TO MO DEL SC A L E L I NES INCL .

I i : (ST I NG HO D) .

i

i )

Figure i-9. In f ac t , t he Harr is pro c edure usually predicts a drag reduction f o r aft-sting modifications where area in- creases result in reductions in the aft slopes of the cross- sectional area distribution as seen with the E205 wind tunnel model. Since the aft-sting increment was not evalu- ated experimentally for the E205 configuration , it's impos- sible to know how much the prediction is o f f due to using the Harris procedure.

The second reason f or the discrepancy between the pre- dicted and test CDm_, mavlie with the inab i l i ty to predict the interference d_. V EO-wing fighter model Wind _unnel data indicate that the inter f erence drag produced by the combination of configuration components is higher than predicted using the same prediction methods as those em- ployed in the E205 prediction. T his "excess" interference drag increment from the f ighter model has been corrected for reference area and added to the E205 minimum drag prediction shown in Figure i-i0 to demonstrate the upper bound of the predicted drag level that might be expected i f the E205 "excess" inter f erence drag were the same as that o f the VEO- wing f ighter model. Actually , the excess interference drag of the E205 configuration is expected to be different from t he VEO-wing fighter because the geometric sources of the interference, i.e., the channel shape between the nacelles and fuselage spine where shocks form at certain speeds , will differ substantially between the two configurations. Since the E205's channel is much more "open , " the excess inter- ference is expected to be smaller than that shown in Figure 1-9 based on the VEO-wing fighter model.

No t e t hat the results of two methods for predicting the Mach= .2 minimum drag are compared in Figure 1-6. The CDm_. = .0203 was estimated using the m_thod desqribed abO V e, that is using the norma± pre a ic_1on me_no a s emp_oye a by General Dynamics for an arbitrary con f iguration for which there is n o prev i ous wind tunnel data. The CDm i n. = .0236 was predicted using the equations described in vol u me Section 3.2 to correct the wind tunnel data from the VEO- wing fighter model and powered research model to a predic- tion of the E205 configuration. This is the method employed to develop the low speed power-on (and power-off) aerodynam- ics used for transition and STOL in Reference 1 and des- cribed in Section 1.4. It is somewhat surprising that bet- ter agreement is achieved with a "generalized" method than with a prediction built up from previous wind tunnel data.

One of the reasons for this is a change in component inter- ference drag between the fighter model and the E205 configu- ration (narrow channel vs wide open strake , etc.).

1.4 Power-Off U ntrimmed Lift , Drag, -_ and Pitching Moment Table 1-2 summarizes the matrix of power-on and power-off, untrimmed and trimmed comparisons of predicted and wind tunnel data for the baseline E205 wind tunnel model that are included in this report. In this section the power-on and power-off untrimmed comparisons are presented while the trimmed power-on and power-off comparisons are presented in Section 2.1.

The predicted power-off , untrimmed lift , drag , and pitching moment curves for the E205 wind tunnel model in the low speed and supersonic speed regimmes are compared with the corresponding wind tunnel data in Figures l-ll through 1-13 _nd 1-18 through 1-19. The predicted, transonicpower- off untrimmed curves were no t developedfor the reasons ex- plained in Volume I, Section 3.2.

At Mach No. = .2 , the predicted wing-body and wing- body-canard (_© = 0 ° and _ TE = 0 ° ) power-off lift , drag, and pitching moment coefficients are compared with the wind tunnel data in Figure i-ii. The predicted data were devel- oped as described in Volume I , Section 3.2 based on the VEO-fighter-model wing-body characteristics and research- model canard , flap , and supercirculation increments. The wind tunnel lift, drag, and pitching moment characteristics are in general more favorable than predicted; that is, for a given angle of attack the wind tunnel data exhibits more lift , less drag, and a more negative pitching moment. There is reasonably good agreement between predicted and test wing-body C L _ and dCM / dCL in the attached flow region (_ > 8-10 ° ); the agreement with predicted minimum drag is also-- acceptable and could be even better if the prediction method employed at other Mach No's were employed at M = .2 (See Section 1.3). Both the predicted and test data indicate early wing separation beginning at _ = 8-10 ° with the wind tunnel model actually producing slightly higher lift and less drag than predicted. The major difference between the predicted and test wing-body data lies in the inability to accurately predict the Cmo of the configuration; the config- uration exhibits more effective positive camber (and hence more nose down moment) than expected. This is a surprise since each component was considered geometrically uncambered (except the fuselage which has a large upswept negatively c_nbered boatail which should produce a positive moment increment).

With the canard on at zero deflection , the agreement between prediction and test data is still reasonably good.

However, there is still a substantial difference between the predicted and test Cmo of about the same magnitude observed Table 1-2 MATRIX OF POWER-ON AND POWER-OFF, TRIMMED AND UNTRIMMED COMPARISONS.OF PREDICTED AND WIND TUNNEL DATA FOR THE BASELINE E205 WIND TUNNEL MODEL POWER SUBSONIC TRANSONIC SUPERSONIC UNTRIMMED TRIMMED UNTRIMMED TRIMMED UNTRIMMED TRIMMED POWER-OFF PREDICTED x - - x x - WIND TUNNEL x - - x x x -J POWER-ON PREDICTED x x - x - - WIND TUNNEL x x - x - - in the wing-body case. As described in Volume III, Section 2.0, the canard effectiveness (Figure I-II) is larger than predicted because the upwash induced by the E205 wing-body is substantially different from that of the research model which does not have a lifting strake area between the nacelle and fuselage spine like the E205. The highly swept, sharp edged strake of the baseline E205 configuration probably creates a substantial vortex flowfield about the strake which extends outboard even past the nacelle to influence the canard flowfield resulting in a higher upwash than expected using the research model canard effectiveness. The changes in the incremental lift, drag, and pitching moment due to the canard in and out of the presence of the wing (Figures 1-14 through 1-17)indicate the effects on the canard flowfiel_ caused by the wing. Please note that these "canard increments" are (canard on - canard off) at a given canard location and de- flection.

The wind tunnel and predicted effects of wing trailing edge flap deflection are shown in Figures 1-12 and 1-13 for - + _T_ = i0 ° and 25 ° . Again , the wind tunnel data is more favorable than predicted with more lift, less drag, and a more negative pitching moment. The major differences are due to the c u mulative errors in predicting the wing-body and wing-body-canard configuration (as descussed above) coupled with the errors in predicting the wing trailing edge flap effectiveness. The low speed wing trailing edge flap ef- fectiveness was also derrived from the Research model data which yields more pessimistic flap increments than obtained with the E205 wind tunnel model as seen in Secti o n 1.5.

The predicted and test untrimmed wing-body and wing- body-canard lift, drag, and pitching curves for Mach numbers of 1.2 and 1.6 are compared in Figures 1-18 and 1-19.

Canard deflections of 0 ° and +i0 ° are presented. These comparisons indicate very good agreement between test and predictions at M = 1.2 primarily because the predictions are based on VEO-wing fighter model data (rather than Research model data) which, although not totally like the E205 configuration , is more similar than the Research model used for low speed and transonic predictions. At M = 1.6 the errors in predicting Cm0 and CDmi. are not acceptable. It should be noted that there are no power effects to be added to the supersonic aerodynamics. Figures 1-20 and 1-21 present the Mach = 1.8 and 2.0 wind tunnel results for _c = 0 ° , 6T E = 0 ° for completeness.

Comparisons of t he M = .2 predic t ed and wind tunnel untri m med " t otal" aerodynamic coefficien t s (which include power effects) were developed according t o the equa t ionsof Volume I, Section 3.2. These "total" aerodynamiccoefficientswere developedby adding the several components to t he unpowered "aero-only" co- efficients described above; the componen t s are as follows: the incremental lift, drag, and pitching moment due supercirculation (from the Powered Research model testing described in Reference I), the engine ram drag, the ejector ram drag, and the vectored thrust components from the forward ejectors and the Veo-nozzles.

There fo re the same p o wer e ff e cts f o r a g i ven f ligh t c o ndi- tion are added to the unpowered predicted and wind tunnel aerodynamics to arrive at the untrimmed powered comparisons shown in Figures 1-22 through 1-25 at Mach n u mber .2 and CT TOTA L = 1.81. T hese data were of course developed to determine comparisons between the trimmed power-on aerody- namics presented in Section 2.0. As explained in Section 2.0 the thrust split between the VEO-nozzles and the forward ejector (Figure 1-26) is a function of angle of attack and flap deflection for a given Mach number and power setting (which fixes CT TOTAL) and are prescribed to arrive at a reasonable trimmed angle of attack range. Therefore , the thrust split is varying along the lift, drag, and pitching moment curves of Figures 1-22 through 1-25 for a given flap deflection according to Figure 1-26.

Since the same power effects have been added to the predicted and wind tunnel unpowered data to arrive at the comparisons in Figure 1-22 , the differences are still primarily attributable to those described above with the unpowered comparisons. In general, the test data shows more favorable aerodynamics than predicted with higher lift , less drag, and a more nose down pitching moment for a given angle of attack. A major part of the difference between the wind tunnel and predicted characteristics lies with the error in predicting Cmo but there is also an error in predicting the trailing-edge-flap effectiveness as discussed in Section 1.5.

1.5 Wing Trailing-Edge Flap Effectiveness As noted in Section 1.4 , the M = .2 predicted wing trailing-edge flap effectiveness for the E205 configuration was derived from the Research model described in Volume I; the Research model power-off and power-on flap increments were presented for the E205 configuration in Reference 1 as a function of flap deflection , angle of attack, and C .

These Research model incremental data were developed in the presence of the undeflected canard and with a fuselage / nacelle / strake arrangement (Volume I) which differs substan- tially from that of the E205 configuration.

Figures 1-11 through 1-13 together with Figures 1-22 through 1-25 present the M = .2 wind tunnel lift , drag, and pitching moment curves for variations in trailing-edge flap deflections of 0 ° i0 ° and 25 " and variations in canard deflections. These data form the basis for developing the lift , drag, and pitching moment increments due to flap de- flection which are shown in Figures 1-3 5 through 1-45 and compared wi t h t he predicted flap i n cremen t s (rela t ive to flap unde- flec t ed) in Figures 1-27 and 1-28. No t e t ha t there is a variation of wind-tunnel-flap effec t iveness wi t h canard deflec t ion while the predicted increments were developed with the canard undeflected.

The wind tunnel data indicates that at M = .2 , the 10°-flap produces more lift than predicted at low _'s (< 8 ° ) for all canard deflections. As the canard deflection is increased or decreased from the undeflected position , the downwash on the wing from the canard produces a decrease in flap effec- tiveness; the flap effectiveness also decreases with in- creasing angle of attack. The center of pressure and hence the flap pitching moment increment is much less affected by canard deflection. The drag increment is affected by canard deflection but increases with increasing angle of attack.

With the flaps deflected twenty five degrees the same trends are indicated with canard deflection and angle of attack at M = .2. However , the decrease in wind-tunnel-flap lift increment with _ and 5C is much more dramatic than at _TE = i0 ° • In general, the low speed flap effective- ness is about as predicted, especially at low _'s and _C'S near zero degrees.

• Figures 1-29 through 1-34 provide the transonic (Mach= .6 to 1.2) lift , drag, and pitching moment curves that indicate the effect of wing trailing-edge flap deflection with varying canard deflection and angle of attack. Because of the way the transonic predictions were developed (see Volume I, Section 3.2) predicted flap incremental effects are not available to compare directly with the wind tunnel flap increments presented in Figure 1-35 through 1-43 for o-- 6c = i0 ° and 25 ° at Mach numbers of .6 , .9 and 1.2. The effect of increasing Mach number is , in general, to decrease the flap lift increment at a given canard deflection and angle of attack with a small change in pitching moment increment but little change in drag increment.

Figures 1-41 through 1-43 indicate the flap lift, drag, and moment increments at Mach = .6 to 1.2 with the canard removed for _ TE = i0° and 25 ° . When these increments are compared with Figures 1-35 through 1-40 , they indicate that the i0 ° and 25 ° flap increments with the canard removed are almost the same as with the canard on at zero degrees deflection (with the exception of the increments for M = .6 and _ 's > 4 ° where the canard removed increment behaves like the 6C = -20 ° case).

Subsonically, the addition of the canard produces a downwash on the wing, lowering the effective local angle of attack of the wing which reduces the adverse pressure gradient on the upper surface of the wing at the wing trail- ing edge-flap allowing the flap to work to higher _'s be- fore flap separation begins. As the Mach number is in- creased to .9 the boundary layer is thinned reducing the adverse pressure gradient at the flap so that the addition i0 of the canard has a less dramatic effect on the flap incre- ment than at lower Mach numbers. Supersonically, (M = 1.2) the flap increments are slightly higher in the presence of the canard (and increase with increasing _ ).

Figures 1-44 and 1-45 illustrate the lift , drag, and pitching moment increments at M = 1.6 and 2.0 due to de- flecting the trailing-edge flap ten degrees in the presence of the canard of varying deflection and with the canard removed. These curves indicate that the canard deflection or presence of the canard has little influence on the flap effectiveness.

1.6 Buffet Onset Characteristics The buffet characteristics of the ejector vehicle are portrayed in Figures 1-46 through 1-55. The wing bending moment coefficient (Crms), an indicator of buffet, is plotted as a function of angle of attack and Mach numbers in Figures 1-46 and 1-47. At Mach= 0.6 , the onset of buffet occur near an angle of attack of 8 degrees. Buffet progres- ses with increasing angle of attack until the wing stall.

As the wing stalls , outboard to inboard , the wing bending moment coefficient decreases and then levels off.

The effect of canard location on the buffet character- istics is displayed in Figures 1-48 through 1-50 for Mach

numbers of 0.6 through 1.2. A t M = .6, compared to the mid loca t ed

canard (CI), the forward located canard (C2) has a milder buildup to approximately the same level of intensity. The aft located canard (C3) has a more direct effect in that the intensity is higher and occurs at a lower angle of attack.

This same effect is much more pronounced at Mach= 0.9. The intensity with the forward located canard follows the trend of the wing body configurations. At supersonic speeds (M = 1.2) this level of intensity for all the canard location is much milder.

Figures 1-51 through 1-53 contain the variation of Crms for various canard deflections. For modest deflections (+i0 , -i0) the levels of intensity differ little. For the case of large negative canard deflections (-20 °) the unporting of the canard causes an earlier buffet onset angle of attack and has a higher intensity as the wing outboard portion stalls first. This stalling is evident from the pitching moment data as well.

II Figures 1-54 and 1-55 compare the predicted and wind- tunnel buffet onset angle o f attack (_ B . 0 .) variation with Mach number for the E205 configuration. In Figure 1-54 three types of _ B . O . indicators are compared, CRMS , CN , and CA. As noted in Reference i, the predicted _ B . 0 . for the baseline E205 configuration were determined by analyzing the axial force data of the VEO-wing fighter configuration force model (from test TF512 conducted at the AEDC PWT 4T Tran- sonic wind tunnel) by using the methods described in Refer- ence ii of Reference i. As noted in Reference I , this VEO-wing fighter configuration model was not specifically instr u mented to obtain buffet data but it appears from Figures 1-54 and 1-55 that the data was adequate to do a relatively good job of predicting the wind tunnel results for the E205 model , especially the trends if not the abs o lute values.

With the 6 C = 6 TE = O ° (Figure 1-54) the wind tunnel data actually indicates a higher (more favorable) _ B . 0 .

than predicted with the Crms and Ca indicators, providing the best agreement with the test data. With the canard rem o ved the limited test data agrees well with the predictions out to Mach .9.

Zero and +i0 ° canard deflections produce higher _ B . 0 .

than predicted (Figure 1-55;) while the negative canard def l e c ti o ns indu c e ear l i e r _ B.0 . tha n predi ct ed.

Figure 1-55 also indicates that the baseline longi- tudinal canard location produces a higher G B . 0 . than either the forw a rd o f aft canard locations.

1 2 2.0 TRIMMED LONGITUDINAL AERODYNAMICS 2.1 Trimmed Pow e r - O ff Aer o dynamic s As noted in Table 1-2 , the E205 trimmed power-o ff transonic and supersonic aerodynamics developed from the untrimmed power-o f f wind tunnel and predicted characteris- tics of Section 1.4 are compared in this section as well as the trimmed _ x)wer-on subsonic and tr a nsonic wind tunnel and predicted aerodynamlcs. Power-off trimmed comparisons between predictions and test data are made at model scale while power-on trimmed comparisons are made for the full scale airplane to confirm the airplane aerodynamics used in the design effort o f Reference i. The subsonic , trimmed power-o ff aerodynamics were not developed from the test data because the M = .6 power-of f da t a indicates the same trends that would be observed if the M = .2 power-off data were trimmed. The supersonic, t rimmed, power-on data are the same as the supersonic power-of f data since there are no supersoni c power effects anticipated.

One o f the original intents in the analysis and com- parison of the predicted and wind tunnel data was to deter- mine and compare e's that are comparable to those used in Reference 1 and displayed in Figure 2-i. These e's were developed from the Research model data by removing camber e f fects to arrive at e's based on an undisplaced drag polar; the source of thecamber effects was experimentally deter- mined to be the wing camber (the fuselage being a body of revolution). Because of the fuselage shape of the E205 configuration , there is an apparent fuselage camber effect that was not determined experimentally. Therefore it is impossible to develop comparable polar e's to those pre- sented in Reference 1 so the drag polar comparisons between predictions and test results are discussed in terms of polar "shape" and C0mi, for power-o f f and power-on cases which is probably more meaningful to understanding airplane performance than e's any way.

Figures 2-2, 2-3, and 2-6 compare the wind tunnel and predicted power-off trimmed drag polars for Mach numbers of .6 , .9, and 2.0 , respectively, while both t rimmed lift and drag curves are shown in Figures 2-4 and 2-5 for Mach numbers of 1.2 and 1.6.

Trimmed lift curves are not presented for Mach = .6 and .9 because of the manner in which these polar predictions were developed directly from envelope trimmed e's (from Re- search model data) as explained in Volume I , Section 3.2.

T he pr e d ict ed p o wer- o f f t rimmed drag polar s of F i gures 2-2 and 2-3 at M = .6 and .9 were developed by using the CDmin estimated for the wind tunnel model (Table i-i) and the power-o f f t rim m ed e's from Figure 2-i which , as noted above , were developed f rom the Research model data using the w i ng trailing-edge f lap deflections 6nly f or trim ( with the canard undeflected) . It was recognized in Reference 1 that these e's were not necessarily the optimum achievable using both canard and flap deflections for trim but they were assumed "representative" o f wha t could be achieved with c anard / f l a p comb i nations given eno u gh exper i mental data.

Figures 2-2 and 2-6 indicate that this was certainly a reasonable assumption.

Two types of trimmed _ower-off wind tunnel drag polars are compared with the predicted polars in Figures 2-2 and 2-6: ( i ) t r imming b y v ar y ing only the wing t rai ling-edge f lap leaving the canard unde f lected, and ( 2 ) trimming with an envelope o f optimum canard and wing trailing-edge f lap combinations (within the experimental data limitations).

These envelope lift curves and drag polars obtained by trim- ming with the optimum canard / f lap combinations were devel- oped as indicated in Figures 2-7 through 2-10. At M = .6 , the wind tunnel polars trimmed with the trailing-edge flap only (canard undeflected) are not as good as predicted; the polar shape is worse and the minimum drag appears to be a little higher than predicted (additional flap-deflection data would be required to accurately determine the trimmed minimum drag). There is closer agreement between the pre- dicted and test trimmed data when using the optimum canard / flap combination d e rived from Figure 2-7 which yields the envelope polar shown in Figure 2-2.

In f act, if the wind-tunnel-developed envelope polar for M = .6 (trimming with the optimum canard / flap combina- tion) is compared with the predicted polar ( trimming with the flap only) at a common C D m in , the test data actually shows a better polar shape than the prediction for C t < .76.

Although trimming with the flap only yields a worse polar shape than trimming with the canard and flap , the trimmed lift for a given angle of attack is lower when trimming with the flap alone but the trimmed L / D is better with the canard and flap combination as shown in Figure 2-11.

At M = .9 (Figure 2-3), the error in predicting C Dmi n I is larger than for M = .6. The actual trimmed wind tunne_ data (trimming with trailing-edge flap only or with canard / flap combinations) has a slightly better polar shape than the estimated polar. If the minimum drag had been estimated correctly , the agreement between prediction and test would have been quite good. Its interesting to note t hat trimming with the trailing-edge flap only or with the canard / flap combination makes very little difference in the lift curve and drag polar shape (as evidenced by Figures 2-2 and 2-3); however , the maximum trimmable _-range with each trim- method is probably different and more test data would be required to determine the _-limits. Figure 2-12 provides a similar comparison of the power-off, trimmed, L / D vs _ with each trim method indicating that trim with the optimum canard / flap combination is better for _ < 3 ° .

Figure 2-4 compares M = 1.2 predicted and wind tunnel trimmed lift and drag curves. Trimming with the optimum combination of canard and wing trailing-edge flap yields a better polar shape and a higher lift slope than predicted.

Trimming with the wing trailing-edge flap only (_C = 0 ° ) produces a higher C Dmi , than pre 4 ic t e 4 but a better polar shape.

The supersonic predictions were based on the VEO-wing fighter model data rather than the Research model data which was used subsonically. The supersonic predictions were trimmed using the canard only with zero trailing-edge flap deflections which compares with the low-C L region of the canard / flap trimmed envelope wind tunnel polar of Figure 2-4 (where the zero trailing-edge flap deflection is used at the low-CL region for trimming). The canard / flap combina- tion does yield a better polar at all CL,s. The differ- ence in the predicted and test values lies primarily then in the inability to predict a.c. and Cmo satisfactorily at the supersonic Mach numbers.

At M =1.6 and 2.0 the errors in predicting CDmin , Cmo and a.c. result in somewhat optimistic predictions as shown in Figures 2-5 and 2-6; the trimmed CDmin is higher than predicted , the polar shape is somewhat worse than predicted, while the CLo is also somewhat lower than predicted. The C L however is as predicted.

2.2 Trimmed Power-On Aerodynamics Power-on , trimmed aerodynamics were developed for the full scale E205 airplane configuration by applying minimum drag corrections to the model-scal e , power-on , untrimmed data of Section 1.4. Comparisons between predicted and wind- tunnel data corrected to full scale were developed at subsonic and transonic Mach numbers. (There are no power effects at supersonic Mach numbers.)

T he low speed, power-on, trimmed co m p a ri s o ns be t we e n the predictions and wind tunnel results (corrected to full scale) are indicated in Figures 2-13 and 2-14. The aerody- namic coefficients presented are "total" coefficients which include the thrust related forces and moments and are de- fined by the equations of Section 3.2 of Volume I.

The maximum total thrust coefficient (C TTOTAD and the ram drag of the engine and ejector are also functions of _ Mach number. Given the hot-day dynamic pressure , q , the engine and ejector ram drag coefficients are calculated according to the equation defined in Section 4.1.1 of Ref- erence i. These ram drag coefficients are assumed to act in the axial direction according as shown in Section 3.2 of Vol u me I. The ejector ram drag coefficient is a function of the air diverted to the ejector, i.e. , the thrust split be- tween the main engine and the ejector. These ram drag and thrust components are also used in the referenced equations of Volume I, Section 3.2 t o determine the thrust induced forces and moments which are then added to the unpowered aerodynamic data to develop the power-on aerodynamic data.

The power-on aerodynamic data for zero canard deflection (from wind tunnel and predictions) are trimmed at Mach = 0.2 using the wing trailing-edge flap and VEO-nozzle plus the forward ejector as trimming devices (see Figures 2-13 and 2-14). The E205 airplane can be trimmed at any reasonable an@le of attack depending on the thrust split (between VEO- nozzle and e_ector) and the flap /nozzle deflection.

The variation of nozzle deflection with the ratio (VEO-nozzle thrust / total thrust) was shown in Figure 1-26.

The relationships between the thrust splits and nozzle deflection were developed and used with the predicted E205 aerodynamic power-off untrimmed data to obtain the pre- dicted, trimmed, power-on lift and drag curves presented in Figure 2-13; these thrust splits and nozzle deflection combinations were also combined with the E205 power-off wind tunnel data to develop the power-on trimmed curves shown in Figure 2-14.

Figures 2-13 and 2-14 indicate that an envelope exists for various attitudes and flap / nozzle deflection where the airplane can be trimmed in pitch. The flap / nozzle deflec- tion can be scheduled as a function of angle-of-attack for operation at a Mach number anywhere in the envelope des- cribed in Figures 2-13 and 2-14. It should be recognized that a set of these envelope trim lift-drag polars exists at each Mach number and that at each Mach number , a thrust split is determined which is a function of the flap / nozzle deflection and angle of attack.

Also in Figures 2-13 and 2-14 a constraint line is indicated which represents the case of all thrust being diverted to the nozzles with none to the ejectors; this line represents a limiting case for aircraft operation.

]6 The shape o f th e drag polar s in Figures 2-1 3 and 2-14 at large flap deflections is typical of a normal airplane con f iguration; this is not the case at lower flap deflec- tions , as the flap de f lection is reduced to zero , increasing angle o f a t tack results in decreased drag. This is the re- sult o f higher thrust required from the ejector for trim a t low angles of attack (causing an increase in ejector ram drag) than at higher angles of attack. The net result is a decrease in the total drag at the higher angles of attack and lower flap settings which in turn results in the polar shapes exhibited.

Comparing Figures 2-13 and 2-14 f urther , the wind tunnel corrected data actually shows better performance than the estimated data due to better power-off flap performance.

Although not shown in Figures 2-13 and 2-14, the variation in canard deflection would result in an in c rease in the size o f the trim envelope but is considered a second order e f fect relative to the trim with nozzle deflection and ejector thrust.

Figures 2-15 and 2-16 compare the Mach= .6 and .9 trimmmed power-on polars for the full scale E205 configura- tion. The transonic power effects were determined for the predicted and Wind tunnel data by calculating the change in induced drag due to power from the p ower-on and power-off e'_ (determined from the Research model as described in Figure 2-1. At Mach = .6 , CM = .302 (combat at i0 , 000 ft altitude) the predicted minimum drag is 14% lower than the corrected wind tunnel data while the corrected test data has a better polar shape than the prediction for C L <.76 and worse for C L > .76. The predictions at M = .6 are certainly within acceptable accuracy levels , especially for prelimi- nary design purposes. At M = .9 , CM = .159 the wind tunnel minimum trimmed drag is higher than predicted while the polar shape is substantially better than predicted. The differences in the predicteed and test minimum trimmed drag are probably primarily due to an inability to predict the minimum trim drag penalty since the unpowered, untrimmed CDmi. differed only by 15 counts (.0237 vs .0222).

t Although there are no supersonic power effects , a comparison of the trimmed predicted and wind tunnel drag polars corrected to full-scale is shown in Figure 2-17 for Mach= 1.2. T he trimmed mi n imum d r ag from the test data and predictions agree well while the polar shape derived from the test data is actually better than that of the predicted polar (just as at m o del scale).

3 . 0 LA T ER A L - D IRECTIONA L AERO D YNAMI CS Comparisons between the predicted and tested lateral- directional characteristics of the baseline E205 configura- tion are examined in this section. In general, the DATCOM procedures were used to develop the predicted character- istics.

3.1 R i gid Sideslip Derivatives Sideforce coefficient de r iv a tive, C y s, i s displayed in Figure 3 -1 as a function of angle of attack for discrete Mach numbers and as a function of Mach number at zero angle of attack. The trend with angle o f attack is rather well predicted. However , the level o f prediction varia t ion with Mach number is somewhat lower. From the analysis o f the Rl04-model transonic data , it was found that the lower- t han- expected variation wi t h Mach number c an be principally attributed to t he sidewash gr a dient. The sidewash for the E205 c on f igurationderivative was predicted as 1.0 while the test data indicates a gradient on the order o f 0.3. This sidewash gr a dient ef f ects the vertical tail contribution to sideforce slope which results in the lower side f orce gradient in the test data.

The consequences of the sidewash gradient is even more apparent in the directional stability parameter, Cn_ , as _ " shown in Figure 3-2. The test level of Cn 8 is approx- imately half that predicted. This is directly attributable to the inaccurate prediction of the sidewash gradient.

The variation of Cn R with angle o f attack is also shown in this f igure and_indic at es that while the DATCOM prediction gives only the slightest variation with angle of attack, the test data shows deteriorating stability as angle of attack is increased. This is mostnotable at the super- sonic Mach numbers.

The level of static directional stability, Cn R , for the complete E205 baseline configuration is stable_at small angles of attack but degrades to static instability at 18 degrees at M = 0.6 as shown in Figure 3-2. This angle of attack for static directional instability decreases with Mach number. At Mach = 0.9 it is 15 degrees and Mach = 1.2 it is 7.6 degrees. The vertical tail contributes to sta- bility to at least 12 degrees angle of attack at M = 0.9 and possibly higher. ( T he largest angle of attack tested with the vertical tail off was 12 degrees.) This indicates that the large forward fuselage is a destabilizing element. The effect of the canard either on or off is very small on the directional stability while undeflected. At small angles of ]8 _ _ attack, the addition of the canard is slightly destabi- lizing. Only slight differences in rolling moment due to sideslip can be noted for canard on or off.

The lateral-directional coefficients are fairly linear in th e r a nge of sides l ip ang le s t es t ed, -6 ° to + I 0 ° . The variation of rolling moment coef f icient with the vertical tail off is only slightly nonlinear. The derivatives dis- cussed above are based on a least square curve fit over a range o f two degrees (0 to 2 " ) .

The dihedral ef f ect, CL_ , is shown in Figure 3 - 3 . The trend with angle o f attack _s predicted fairly well in the linear region when compared to the test data. After the wing becomes ineffective (angles of attack of 8 to i0 de- grees because of no leading-edge protection) the prediction- test correlation is less f avorable. The characteris t ic level o f CI_ with Mach n um ber at a constant angle o f attack is predicted well subsonically but the comparison diverges as supersonic speeds are attained.

3 .2 Vertical Tail Ef f ectiveness Several wind tunnel data runs are available from which the vertical tail effectiveness could be determined. These were the tail-off and vertical-tail-deflection runs for zero angle of attack (_ _ 0.2 " ). Runs with vertical tail deflec- tion were also conducted at 18 degrees angle of attack but unfortunately there was no vertical t a il off data at that angle of attack. The results of the analysis of this data indicate a sidewash-sideslip gradient of approximately 1 / 4.

This gradient degrades the effective lift curve slope of the vertical tail. As noted above, the predicted sideforce derivative is higher than the test value primarily because a sidewash gradient of 1.0 was used in predicting the E205 lateral-directional derivatives. This is also reflected in the predicted-test comparis o n o f Cn_ . The va ri at i on of dihedral effect, however, which is primarily a wing func- tion , was predicted fairly well.

Directional control effectiveness is displayed in Figure 3 -4. The level of effectiveness is predicted well at low speed but deteriorates as speed is increased. Apparently the force generated by the vertical tail surface is less than predicted.

3.3 AILERON EFFECTIVENESS Aileron e ff ectivenes s i s presented in Figure 3- 5 . The predicted level is lower than the test results at zero angle of attack. This may be attributable to the DATCOM predic- tion technique. The location of the ailerons on E205 are immediately behind and outboard of the nacelle. The DATCOM m e t h od d oes n ot pr o v i de fo r any inter fe ren c e o f th is t ype , either favorable or adverse. As the wing becomes less effective at angles of attack from 6-10 degrees , the aileron effectiveness drops to become less than predicted.

4.0 REFERENCES i. Lummus, J. R., Study of Aerodynamic Technology for a VSTOL Fighter / Attack Aircraft, NASA CR-152128 , May, 1978.

2. Carmichael, R. L., Costellano , C. R., and Chen , F. C., "The Use of Finite Element Methods for Predicting the Aerodynamics of Wing-Body Combinations , " Analytical Methods in Aircraft Aerodynamics, NASA SP-228, October, 1969.

3. Hoak , D. E. , USAF Stability and Control DATCOM, October, 1960.

4. Heim , E. R., Basic Aerodynamic Data for a Vectored- En@ine-Over-Win@ Confi@uration, AEDC-TR-78-1 , February 1978.

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f:::T- Figure 1-1 7 Comparison of Incremental Effects' of Canard"'''' 3:L._ Deflected -20 for Test Data, ,Mach =-.2 f;;I,:;::~~:;;~:;;: E20S,

_i! _ _!i_

_+ F i g u r e i 'lS a Lift and M om en t Com pa rison of P r e dicted an d Test L on_ gitudinal A e rody n amic Cha r act e r istics of B as e l i n e E 205 Co n figuration , Powe r - Off, Mach = 1.2 TEST RUN MACH LEF TEF HORIZ.

X 32 4 22 1. 20 0 . 0 0 . 0 0 .0 D ' Predic t ed 1.20 0.0 0.0 0.0 /k 324 78 1.20 0.0 0.0 I0.0 <> P r edi ct ed 1. 20 0 . 0 0 . 0 I 0 . 0 CRE F = E205 F i gurel-18bDrag C o mparison o f Predicted and Test Longltudinal Aerodynam i c Characteristics of Baseline E 205 Configuration , Power-Off , M _ icli = 1.2 :

• • •

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Aerodynamic Characteristics o f Baseline E 205 Configuration , Power-Of f , Ma c h = i.6 {: : | ARC - 9X7 - 324 _(1 _:._1 E205 _ii i !i !_il Pr e dl c t e d I 1. 60 0 . 0 0 .O i 0 . 0 ii ii! l 324 646- 1 .60 0 . 0 0 .0 1 0 . 0 . , .1 P red i c t e c ] 1. 60 0 . 0 0.0 I I 0. 0 _Li :: : ; | ; i_it + j ;:: : ] ; :: : I

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... ~: T: Figurel-21cDrag Data for Baseline E205 Configuration, Mach = 2.0 1 ) 0. 20 0.0 0 . 0 0 .0 Tun nel [3 0. 2 0 o.o 1o.o o.o 0 .2 0 o.o 2 s. o o .o l() W_ ; )i ()i in :!ii E205 :i, * .... _ _!_ ', ;1 : ARC-12-327 , , i_ = 1. 81 !ii _ - i CTToTAL !_I_ i S REF = 38 4. 0 0 f t.2 l.l IIIIII1 ; II ; II;IUI = 142.68 i i_i!i, u : NOZZLE See Figure 1-26) Ui

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_U F i gurel-2 2 aEffect of Wing T ra L ling- E dge Flap Deflect L 0 n on Lift and Moment f o r T est: and Predicted Data , Ma c h= .2 I !!!! !_i: ARC- 12-327 i fl it: i11: :I:: I E 205 0.20 0.0 0.0 0.0 Tunnel 0 . 20 0 . 0 1 0 . 0 0 . 0 /k 0. 20 0 . 0 25 . 0 0.0

S CHL FT F.OR ZW DPR D CTED

_:! .... SRE F = 384.00 ft _!i itt !!i E205 iii ilii :h iil :I :ll Figurel-22bEffect of Wing Trailing-Edge Flap Deflection on Drag for Test and Predicted Data , Mach= .2 SYM MACH L E F T E F HORIZ• X 0 .2 0 0 . 0 1 0 . 0 - 10 . 0 [] 0 .2 0 0 . 0 1 0 . 0 0 . 0 A 0.20 0 .0 I0.0 I 0.0 E205 Figurel-23aLift and M o ment Predicted Data with Canard D efle c ti o ns and Wing Trailing-Edge Flap Deflected +i0 ° , Mach = .2 .

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X 0 . 20 0 . 0 25 . 0 - I 0 . 0 _i_ [] 0 . 20 0 . 0 25 . 0 0 . 0 A 0 . 20 0 . 0 25 . 0 1 0 . 0 i!i : iSREF = 384.00 ft 2 E2O5 iili : _ : _ : !!!

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Figurel-25bDrag Predicted Data with Canard Deflections and Wing Trailing-Edge Flap Deflected +25 ° , Mach = .2

t ) '

ill i[ ii i![ _ Thrust Split Vs. Flap Deflectior

H

:[ +H _Thrust Split _ t

TMAI

6F Deg urel-2( Thrust Split Between VEO-Nozzles and Ejec-: tors Required to Achieve Pitch-Trim as a Function of Trimmed-G and VEO-Nozzle / Flap Deflections for Mach 2 , C 1 81 Sp. TEST MAC}{ C ONFIG. HORIZ. i TEF.

ARC-12-327 327 0.20 BSNVW C I I0.0 I0.0

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i : i .

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i ,. w1th W1ng Tra1l1ng-Edge Flap Deflected +25 0 ~I.:,;V·

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!!! [ ] :[i i!{i I ;; i i : _I : , i _iii ! iiil g_ il; ll [ l l !1_ :i , ,i,i, Figuretl-29aEffect 0-f Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +10 ° , Mach = .6 '::::i, i"i::: ::j:"'.':' :':1"':,::1':: ,: .

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E205 ............... I ............ !........

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Figurel-30aEffect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +I0 ° , Mach = .9 ) ) J !iiJ.................................

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• • •

Figurei-31bEffect of Canard Deflection on Drag With-Wing Traili n g-Edge Flap Deflected +i0 ° , Mach = 1.2 - 4 !.: :_i_, _ ii_ ,i : Figurel-32aEffect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +25 ° , Mach = .6

• • •

E205 i Figurel-3 2 bEff ect o f C anard De f le ct ion on Drag Wit h W i ng T r ai ling- E dge Flap Deflected +25 ° , Mach = .6 o E205 '.: . I , .I i .. , _ _._ .... i _i_: ¸,. . 2_ _,, , . J ..... b..... _ .

Figurel-33aEffect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +25 ° , Mach = .9 E205 ._ j J.++i ..l. + i lJi! ..... __ . t . __ _ t __ _ .__ Figurel-33bEffect of Canard Deflection on Drag With Wing Trailing-Edge Flap Deflected +25 ° , Mach = .9 E205 "1 : ." : I ] . ' .

! it! ; - . 8 i!_ :i_2 t i : ' !it;_ . 4 j : ; ' : H :. 2 .. , - I _.. _k ....

Figure i -34aEffect of Canard Deflection on Lift and Moment With Wing Trailing-Edge Flap Deflected +25 ° , Mach= 1.2

• • •

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__ i '_._-J_-_,_ F i gure 1- 3 6 In c remen t a l Ef f ec ts D u e to D e f le ct ing th e , Wing Trailing-Edge Flaps +i0 ° while in ..... Presence of Various Canard Deflections Baseline E205 Configuration , Mach = .9 .... = ............... '"":: : : ......... ' ....................... _-; : -' L,T: i:_ii-_ , _ ......_ '_L :i - TiT-: - .T_ 7 ] - _ : -. -- _L : _ 7 / =" 7- L • _: : l : z ] ! ! _ : !.i c - --- c j 4 _ _ . h l_ ' €_ : . :t .: . _ . .: : : :. :. v : , t __ . l: : l i : : : " . : !: : !I_' _ I ' _''I , : : i_ . _ :: . _ ' _ .... i _: i......... ' ..::.

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:--.- -8.1_ :i-,!---:_-.-,_L-I- J . I. ,-1,- d--t- L . __ ... ~_._L.' .. Lr--:;""f"'l'--+-+C,j -.- Figure 1-47 Effect of Mach Number on Wing Buffet, Wing -e-- Trailing-Edge Flap and Canard Undeflected _t~ I_ i_ 1_. : Figure 1-48 Effect o f Canard L o ngitudinal L oc ati o n o n Wing Bu f fe t , Ma ch = .6 ======================= = = = === = =-_: ; -:-I:::r. : r: _::1::::_: : --_-:t ::]':......:=:::: : ::m= : :: = : _ - _ = =: : :::t: : : : : : b: : :t. : :_ : :: : ._::=: , =:=T=:_ :_ .- :F:_:t:_ . .l_=_._== : _:: ::-_* .= : -.

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• • , _ _ I ;77_!i_ii!I ....... I....i::i,:::112_.,_ [ 20 _ r - / v _1

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: O CRMS "

tm A-

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-- t .......... '........ _ .........................................................................................

:: :-:_::::t: -', , : : - _: - : '. .... ;. Mach Number - :-: : : --: . 1::: .... :::: Figure 1-54Comparis0n- 6-f -- Pred - icted-_K 0 " with _B. O . - Indicators C_ S , CN, and CA from Test Data ......... - t - .... '...,.7_-Fr .....

',-:-::::::', , ................. - .;- ............. -i-_ - ::, .......... :--_-::,_ .... .... -" - -- :- :.... ::_ - ;:---_ - ::ii _ :i_ii-E:__ : i:-. ' ,i . _ - l-:'i:! - F - qti-::Ei -- ........ q . i ....... : "_:! ::tT: '_ :-" ; .... ': - _- ' _ : ___ _n___{ ................... i ......... I .......... Buffet Onset _ ............ ;-- : 1 ....... " ; " .......

..... . . . ._ : .: i. -1:. - i : : i .... :: :" " • t ., . .......

• _ .... : " ' : : : " Basic " _.... ' ........................ _ :": : "[ :ii / _ ..... ' ....... Configuration ......... : :]":;:: ' ".... I i"" Effect of Canard Deflection ,.. - :: 6c Based on CRM S Evaluation ........................... _ - ......... " .... • _

, 4__ O 0

[3 +i0 ° •v^--lO - " ................... ....... 0 . .

_BO ............... _C - = 0 ° ? Predictions

e

_k _ / " " " E] O Mach Number _2 Effe ct of C anard Posit i on : J :.

........ 1 .. . ....

o

/ Cl _BO

- -" ou ............................ . - ! ............... /

........... _-- e_ . .. . t

- L .:_'-_ _._ 1 . .. ; . . . ; M : ------' .

,¢.- ¢ : . :. _[l ] . _ ; . . i. .1 . . ' " • - - i .......... F ............................ ' Et. Ei ; " " i-::: . _ :. _ : Mach Number .... i : .{EE : ' ':"_ " " ' " Figure 1-55 Effect of Canard-befie : ct__nd Ca ' nard Location on _B.0 .

I B MAX POWER 159 MAX A / B POWER . Figure2-1_, Power-on and Power-off iPredictedTrimmed,e's as a Function of Equivalent : Lift Coefficient, C LE, Mach Number, and C_ (from Reference il) Figure 2 ' 2 Trimmed L i ft and Drag wi t h Wing Trailing - Edge _Flap De f lec t i o ns and Canard U ndeflected , Mach = .6 ,i -- Y r-- 0. 9 0 Tr im w i t h Can ard Fix ed ; Var y TEF --.-- 0 . 90 T rim with Opt Co mb o f C an a r d an d T EF ii!i

!!!!

0. 9 0 P red i c ted Pow e r - off mod e l t r im; Fi x e d _ : ': _.

C ana rd V a r y T EF iii ill 0 : ' ! 1 ' iii :it m ! '.; !ii iil Figure 2- 3 Trimmed Li f t and Drag wi t h W ing Trailing-Edge Flap Defle c ti o ns and Canard Undeflected , Mach = .9 ---- -- --.Tri m wi t h C a na rd Fixe d , V a ry T EF O-- . --Tr i m w ith Opt C o mb o f C an ar d / TEF M=I'2 .... Pre d ic t e d-Canard V a ries Fi x e dTEF [ 28 5_ - ......

Figure 2 - 4 Trimmed Lift and Drag with Wing Trailing-Edge Flap Deflec t ion and Canard Undeflected, Mach = 1.2

) )

ARC-9X7-324 o Figure 2'5 Comparison of Power-off 'Primmed - Lift Curves and Drag Polars for Trimming with Flap-Fixed , Canard Varies and with an Envelope of Opti m um Canard and Flap Deflections , Mach = 1.6 : :: i 20 5 SYM M ac h ,!_ .

2 . 00 Wi n d T_ n el E n v elope-Flap + _ SRE F = 384. 0 0 ft. 2 : - -- 2.0 0 Pre d ic t e d - Fl a F ix e d - C a n a rd Va r ie s _I: CGRE F = 3. 00% r:I !:1 E205 _ ar d i_it ; . r - i)il i7 !i!{_ rla_ :il l lli!i !.._ • i : :1 _ • _1 i $ i! ;t _iiligi: t i:i -.-, : ::1 _ ! i_ii i!! l if:; Figure 2-6 Comparison of Power-off Trimmed Lift Curves and Drag _i_0-].a-ra for Trimming with Flap-Fixed, Canard Varies and with an Envelope of optimum Canard and Flap Deflections , Mach = 2.0 i_ -- !_" _ E205 _ I I i : :x |:x: t-- _ .

_ .

L

E

Figure 2-7 Trimmed Lift and Drag for Baseline E205 Configuration Using Canard and Trailing-Edge Flap Deflections , Mach = .6 :::| :::| O O% Figure 2-B Trimmed Lift and Drag for Baseline E205 Configuration Using Canard and Trailing-Edge Flap Deflections , Mach = .9 E205 o Figure 2-9 Development of Envelope Trimmed Lift Curve and Drag Polar, Mach = 1.6 E2O5 o Figure 2 - 10 Development of Envelope Trimmed Lift Curve and Drag Poiar , Mach= 2.0 i L / DTRIMMED _anard and TEF

g

kO Figure 2-11 Comparison of Trimmed , Power-off vs _ for Trimming with Canard Fixed , Varying Wing Trailing-Edge Flap with lOptimum Combination of Canard and Wing Trailing-Edge Flap , Mach = .6 M L / DTRIMMED of D-d Angle of Attack _Degs i Figure 2-12 Comparison of Trimmed , Power-off L / D vs _ for Trimming with Canard Fixed , Varying Wing Trailing-Edge Flap and with Optimum Combination of Canard and Wing Trailing-Edge Flap , Mach = .9 and Curve- Lift Power-on, Trimmed Figure 2-13 Full-Scale E20S Airplane Predicted, = 1.81

Drag Polar- Envelopes for M = .2, CT

TOTM Figure 2-14 Full-Scale E205 Airplane Power-on , Trimmed Lift Curve- and Drag Polar-Envelopes from Wind Tunnel Data for M = .2 , CTTOTA L = 1.81 i j f t. 2 E s tim a t e d!

Wi nd Data Co rr e cted C G = t o F. S . + P o wer II [!

II ii !! - , !!

!!

ii N-" ii H- H- i, of Full-Scale E205 Airplane Trimmed , Power-on Drag Polars from Prediction and Test Data , M = .6 , C_ = .302 (Optimum Canard / Flap Envelope Trim for Prediction) SREF = 384.00 ft.2 .159 Estimated Ao / Ai = 1.0 0 9 .159 Wind Tunnel Data Corrected Trimmed to F.S + Power r l i ] N x_ • _ : I tl H! _,1 tl: ! t t_

t , i ll

t l Figure 2-16 Comparison of Full-Scale E205 Airplane Trimmed , Power-on Drag Polars from Prediction and Test Data , M = .9 ,, C_ = .159 (Optimum Canard / Flap Envelope Trim for Predicti o n)

1 )

SRE F = 384.00 ft. 2 Ao / Ai

.0 3

Figure 2-17 Full-Scale E205 Airplane Tr mmed D rag Polars from Pred _lons and Test Data , M = 1.2 MACH NUMBER.

Figure 3-1 Variations of Predicted and Test E205 Side , , ,,,:,_: ;_:Force_ Derivative , , Cy_, wi , th ' _!aC,h ,Number and +_ E205 RIGID S IDESLIP DERIVATIVE

308.86 FS

14 5 WL j C _ 0.6 - 79X [i. 2 - 81X I: i: '! M -O.

i ; € 4 + ;i i Predicted -- _=MACH NUMBER Figure 3-2 Variation of Predicted and Test E205 Yawing = _ Moment Derivative , C , _ , with Mach Number and_ Figure 3-3 Variation of Predicted and Test E205 Rolling ____ Moment Derivative , C[_ , with Mach Number and _ !18 E205 RIGID VERTICAL TAIL EFFECTIVENESS _r

!! REF _ 3O8.86 FS

145 WL MACH NUMBER MACH Figure 3-4 E205 ,redicted and Test Vertical Ta, Effectiveness at Various_ Mach Numbers - + + l Figure 3 - 5 E20 5 Predicted and Test Aileron Effec t iveness

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