Analysis of wind tunnel test results for a 9.39-per cent scale model of a VSTOL fighter/attack aircraft. Volume 1: Study overview

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ANALYSIS OF WIND TUNNEL TEST RESULTS FOR A 9.39-PER CENT SCALE MODEL OF A VSTOL FIGHTER / ATTACK AIRCRAFT VOLUME I - STUDY OVERVIEW DR. J. R. LUMMUS G. T. JOYCE O C. D. O'MALLEY Prepared under Contract NAS2-103A4 by General Dynamics r : ;, -,,,,,_.:, , ,_: ? ._,_ Fort Worth Division i :-- _ .... ' :,, for

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4. Title and _ i _ 6. Re1 _ m't _ " An al ysis o f Wi nd T unn e l Test R es u lts for a O cto b e r,198 0 9.39-per c ent S c aleModel of a VSTOL Figh t er / Atta c k e. pe_ m, _o_,,i u ,i_ c _ Air c raft 7. Autl _ r($) 8. I _ w f _ m _ ; _ n , z = z,_ R e po r t No.

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P. O. Box 748 NAS2-10344 Fort Worth, Te x as 7 6101 1 3 . T v _ o t , ,=o _mdP . . ._ _,._ _ . s_ , _ , cv _ ,, _ _ x _ rm C o ntractor Final Report Sept.i0, 1979-Feb.lOt 1981 NASA, AMES R e s e a r c h C e n ter, 1 4 . sl,omm , +_ A _- c v C o_ Moffe t tF i eld, C a 9403 5 15. _ D _ emen _ r¥ Notl _ AM E S R e s e ar c h C ent er T ec hni c a l MonitorW . P. N e lms (41 5 ) 96 5- 5880 16. , ll , l_i x ra _ The resultsof a seriesof NASA AMES wind tunnel tes t sof a General Dyna m i c sve c tored-englne-over Wing,Navy VSTOL figh t er / a t tack c onfiguration have teen analyze d to (I) assess p redi c t i on m e t ho d c apabilities, (2) evalua t egeometr y var i ationssu c h as m d ltiple c ana r dlongitudinal lo c a ti ons and strakeshapes,and (3) evaluatethe effe c tsof configuration c hangesasso c iate d with varyingthe propulsivel i ft system from a j et- diffuse r e j e c torto a Rem o teA u g m enta t ion Lift System (RALS). C onflgura- _ t l onmodif ic ation an d add i tional testingand analysisare re c ommended to . a d eq u ately eval u a t ethe co nf i gura t ion po t en t ia l .

This doc u mentis presen t ed in f o u r volumes- V o lume I - Study Overview Volu m e II - Evaluationof Pred ic t i on Methodologies, Volu m e III- Effectsof Config u rat i o n V ariationsfr o m BaselineE205 C onfiguration on Aerodynami c Chara c teristics, and Volu m e IV - R A LS RI04 Aerodyna m i c Chara c teristi c s and Co m parisons with E 20 5 C onfigurat i on Aerodynamic Charac t eristics.

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VOLUME I - STUDY OVERVIEW TABLE OF CONTENTS i. Su mm ary 1 - 2. Introduction 3 3. Review of Airplane Configuration Development Study 6 3.1 Configuration Descriptions 8 3.1.1 E205 Ejector Configuration 8 3.1.2 RI04 RALS Configuration 13 3.2 Predicted E205 Full-Scale-Airplane 19 Aerodynamic Characteristics 3.3 Resulting Aerodynamic Uncertainties 27 3.4 Description of Wind Tunnel Models 30

O 3.5 Wind Tunnel Test Programs 33

4. Conclusions and Recommendations 35 5 . Referen c es 42

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iii LIST OF FIGURES Figure

2-1 VEO-Wing Powered Lift Concept 43

2-2 E205 Three-View Drawing 44 2-3 RI04 Three-View Drawing 45 2-4 Artist Concept of E205 Configuration 46 2-5 Artist Concept of RI04 Configuration 47 3-1 DLI Mission Profile 48 3-2 E205 Airplane Cross-Sectional Area Distribution 49 3-3 RI04 Airplane Cross-Sectional Area Distribution 50 3-4 VEO-Wing Experimental Data Base 51 3-5 Longitudinal Forces Acting on E205 - 52 3-6 Equations for Building U p Aero-Only Coefficients 53 for E205 Airplane Configuration 3-7 E205 Minimum Trimmed Drag vs Mach Number 54 O 3-8 Power-Off Wing Body Lift , Drag, and Pitching 55 Moment Characteristics for the VEO-Wing Fighter Model of Reference 4, M=.2 3-9a Lift and Moment Comparison of Predicted and Test 56 Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration , Power-Off , Mach= .2 3-9b Drag Comparison of Predicted and Test Longitudinal 57 Aerodynamic Characteristics of Baseline E205 Configuration, (Expanded Drag Scale), Power-Off, Mach = .2 3-10a Lift and Moment Comparison of Predicted and Test 58 Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +i0 °, Power-Off, Mach = .2 3-10b Drag Comparison of Predicted and Test Longitudinal 59 Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +i0 °, Power-Off, Mach = .2 3-11a Lift and Moment Comparison of Predicted and Test 60 Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +25 °, Power-Off, Mach = .2 iv O LIST OF FIGURES Figure Page 3-11b Drag Comparison of Predicted and Test Longitudinal 61 Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +25 °, Power-Off, Mach= .2 3-12 Full-Scale E205 Airplane Power-on, Trimmed Lift £2 Curve- and Drag Polar-Envelopes from Wind Tunnel Data for M = .2 , CT TOTAL = 1.81 3-13 E205 Trimmed Cruise / Maneuver Drag Polars 63 3-14 VEO-Wing Trim Method for Maneuver 64 3-15 Power-on and Power-off Predicted Trimmed e's as a 65 Function of Equivalent Lift Coefficient , CL_ Mach Number, and C_ (from Reference i) 3-16a E205 Lift and Pitching Moment Curves at 66 M = 1.2 with Canard Deflection 3-16b E205 Drag Polars at M = 1.2 With Canard Deflection 67 D 3-17a E205 Lift and Pitching Moment Curves at M = 1.6 68 With Canard Deflection 3-17b E205Drag Polars at M = 1.6 with Canard Deflection 69 3-18 E205 Predicted Aerodynamic Center 70 3-19 Aerodynamic Center Test / Theory Correlation 71 3-20 E205 Buffet Onset Angle of Attack 72 3-21 E205 Lateral-Directional Characteristics 73 a. C¥_ vs and Mach No. 73 b. C n vs and Mach No. 7 4 c. C_ vs and Mach No. 75 d. Vertical Tail Effectiveness 76 e. Aileron Effectiveness 77 3-22 Three View Drawing of E205 Wind Tunnel Model 78 D 3-23 Photos of E205 Wind Tunnel Model 79 V LIST OF FIGURES Figure Page O 3-24 Cross-sectional Area Distribution of E205 80 Wind Tunnel Model 3-25 Three View Drawing of RIO4 Wind Tunnel Model 81 3-26 Cross-sectional Area Distribution of RIO4 82 Wind Tunnel Model 3-27 Photo of RIO4 Wind Tunnel Model 83 3-28a Variations in Canard Locations for E205 and RIO4 84 Wind Tunnel Models 3-28b Variations in Strake Shape for E205 Wind Tunnel Model 84 4-1 Recommended E205 Canard and Wing Planform Change 85

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vi O L I ST OF TABLES TABLE PAGE 3-1 Dimensional and Design Data i0 For Ejector E205 Configuration 3-2 Dimensional and Design Data 15 For R104 Configuration 3-3 Methods Summary 20 3-4 E205 Minimum Drag Buildup 23 3-5 Summary of Constants 31 3-6 Available Component Deflections 32

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LIST OF SYMBOLS 1- a. English Symbols A axial force, ib (N) • a.c. aerodynamic center, %_ AR aspect ratio b span, in. (m) c, MAC mean aerodynamic chord, in. (m) CA axial force coefficient CA axial force coefficient due to ejector ejector CD drag coefficient CDAER 0 aero-onlYincluded) drag coefficient (no thrust increments O CDmi n minimum drag coefficient CDE equivalent drag coefficient CDRAM ram-drag coefficient (engine inlet) CDt total drag coefficient CL lift coefficient CLbuffe t buffet-onset lift coefficient CLE equivalent lift coefficient CLmax maximum lift coefficient aero-only lift coefficient (no thrust increments CLaer° included) CLt total lift coefficient CI rolling moment coefficient O viii O LIST OF SYMBOLS (Continued) CI_ rolling moment derivative due to sideslip, I / deg CmE equivalent pitching moment coefficient Cmx- pitching moment c o efficient about x percent c C zero lift pitching moment coefficient m o Cm total pitching moment coefficient t CN normal force coefficient Cn yawing moment coefficient Cn_ yawing moment derivative due to sideslip, i / deg T CT thrust coefficient, qSREF side force coefficient O _ 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 ground effect L lift, ib(N) i F lift due to supercirculation, ib(N) i rolling moment, ft ib (Nm) M Mach number m pitching moment, ft Ib(Nm) Total Pressure Q NPR nozzle pressure ratio, p ix LIST OF SYMBOLS (Continued) O N normal force, ib(kg) 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) m S C canard exposed area, ft2 (mZ) Sre f reference area, ft2(m2) (usually equal to SW) STOL short takeoff or landing area of trapezoidal wing extended to centerline, O SW ft2(m 2) 2 (m2) SVT exposed area of vertical tail, ft T thrust, Ib(N) Vo_ freestream velocity, ft / sec, knots (m / sec) V. jet velocity based on isentropic expansion from J 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 weight flow, ib / sec (kg / sec) X action point of circulation lift relative to cp leading edge of MAC

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x Fo _ Wo _ h Divis _ n _ _ 0. Box 7 48, For t Wor _ , Te xas76 101 • 81 _ 7 32 _ 811 LG:mb / 10344-FW#060-15 _ / _t___-_ _ . _._ 3 February 1981 ___ _ _ __0 Subject: Contract NAS2-I0344 Contract Compliance Submittal Final Report - NASA CR-152391 To: National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035 Attention: W. P. Nelms N-227-2 Enclosure: (A) One (I) reproducible and twenty-five (25) copies of NASA CR-152391 "Analysis of Wind Tunnel Test Results For A 9.39-Per Cent Scale Model of a VSTOL Fighter / Attack Air- c raft" Final Report September 10, 1979 - February 10, 1981, dated October 1980 Volume I - Study Overview Volume I - Evaluation of Prediction Methodologies Volume III- Effects of Configuratio n VariationsFrom Baseline E205 Configuration on Aerodynamic Characteristics Volume IV - RALS RI04 Aerodynamic Characteristics and Compari- sons with E205 Configuration Aerodynamic Characteristics (B) Distribution List for NASA CR-152391 I. The Report, submitted as Enclosure (A), covers work conducted by General Dynamics under NASA Contract NAS2-I0334 for the joint AMES / NSRDC VSTOL Aerodynamic Technology Program.

2. Enclosure (A) is submitted in compliance with Article II.C.1.c and Article V.B.3 of Contract NAS2-I0344.

3. Distribution is being made in accordance with Enclosure (B) list suDplied by Mr. W. P. Nelms, Contract Technical Monitor, NASA-Ames Research Center.

GENERAL Dynamics Fort Worth Division

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Director of Engineering Administration Page 2 I0344-FW#060-15 To: NASA / W. P. Nelms cc: w / Encl. (A) Ames Research Center Patent Counsel N200-11A Ames Research Center Technology U tilization Office N240-2 c c : Ames Research Center Technical Information Division N241- 1 2 Ames Research Center Contracting Office, Carolyn S. LaFollette N241-I Direct transmittal is approved by the Air Force. Your answer , if any, will be forwarded through regular Air Force channels to General Dynamics, Fort Worth Division, Fort Worth, Texas.

Page 3 I0344-FW#060-15 To- NASA / W. P. Nelms CC : J. R. Lummus, 2882 W. R. Kayl, 2870 T. M. Pettigrew, 1649 L. G. Graham, 2236 (2) Engr. Files, 2228 Cortes. Control, 1137 Contract Records, 6699 tw Enclosure (B) to I0344-FW#060-I5 DISTRIBUTION LIST FOR NASA CR-152391 Mr. W. S. Aiken,'RJ-2 Mr. Hal Andrews National Aeronautics & Room 412 Space Administration Jefferson Plaza #I Washington, D.C. 20546 1411 Jefferson-Davis Highway Arlington, VA 20306 Mr. John Ward, RJL-2 National Aeronautics & Mr. M. W. Brown Space Administration Room 904 Washington, D.C. 20546 Jefferson Plaza #2 1421 Jefferson-Davis Highway Mr. Jack Levine, RJH-2 Arlington, VA 20360 National Aeronau t ics & Space Administration Capt. D. C. Troutman Washington, D.C. 20546 Project Manager, V / STOL Room 674 Mr. P. J. Bobbitt Jefferson Plaza #I M.S. 285 1411 Jefferson-Davis Highway NASA La n gley Research Center Arlington, VA 20360 Langley Station Hampton, VA 23665 Mr. R. F. Siewert O U SDR&D Mr. R. V. Harris Room 3D1089 M.S. 407 Pentagon NASA Langley Research Center Washington, D.C. 20301 Langley Station Hampton, VA 23665 Mr. T. J. Brennan Naval Air Development Center Mr. L. W. Gertsma Code IVA M.S. 500-208 Warminster, PA 18974 Lewis Research Center NASA Co. Peter J. Butkewicz 21000 Brookpark Road Chief, Aeromechanics Div.

Cleveland, Ohio 44135 AFWAL / FIM Flight Dynamics Laboratory Mr. H. R. Chaplin Wright-Patterson AFB, Ohio NRDC, Room 219, Blag. 7 45433 Corderock, Maryland 20034 Mr. John Chuprun, Jr.

Mr. James H. Nichols, Jr. 3 Cys. ASD / XRHD NSRDC, 208A, Bldg. 13 Wright-Patterson AFB, Ohio Corderock, Maryland 20034 45433 • Ir " I ,, Q LIST OF SYmbOLS (Continued) b. Greek Symbols a alpha angle of attack, deg 8, beta angle of sideslip, deg F supercirculation y flight path angle, deg 6c,_i canard deflection (positive, leading-edge up), deg _TE' 6 F VEO-Wing nozzle and outboard flaperon deflection, deg; except for aileron action the flaperons and VEO-Wing nozzle flaps always deflect together.

pitch attitude angle, deg 8 3 jet thrust deflection out of VEO-Wing nozzles when deflected, deg O TE' ALE leading-edge sweep angle, deg taper ratio, tip chord root chord ejector measured thrust / isentropic supply thrust (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)

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xi LIST OF SYMBOLS (Continued) O c. Model Symbols BI VSTOL ejector configuration E-205 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 in VSTOL E-205 or RALS RI04 fwd-location C3 Horizontal canard in VSTOL E-205 or PALS RI04 aft-location N VSTOL ejector configuration E-205 or RALS RI04 VEO-wing nacelle Q 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 PALS configuration R-104 wing with linear elements between SS 87.231 and SS 214.430

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i . o SUMMARY This document presents the analysis of a series of NASA AMES wind tunnel t ests of two General Dyn a mics Vectored- Engine-Over (VEO-) Wing, Navy VSTOL fighter / attack con- figurations; the airplane configurations and wind tunnel models were developed during a previous NASA AMES contracted effort described in Reference i. The two configurations differ primarily in the propulsive lift systems employed for VTOL operations, the jet-diffuser ejector (E205 configuration) and the Remote-Augmentation-_ife System (RALS RI04 configura- tion); both configurations employ the VEO-wing concept for improved maneuver and STOL performance.

The wind tunnel data has been analyzed to (I) assess the prediction method capabilities, (2) evaluate geometry variations such as multiple canard longitudinal locations and strake shapes and (3) evaluate the differences in the aerodynamic characteristics of the two configurations, i.e., the effect of the propulsive lift system on the airplane arrangement and subsequent performance.

I The existing prediction methods were found to be sur- prisingly effective for such unusual configurations but areas of concern were uncovered where improvements in prediction capabilities are certainly worthwhile. The experimental data base gathered in this series of tests forms one of the most complete, systematic parametric variations of con- figuration variables existing in the literature available to the designer and as such, should represent a very valu- able aid in years to come. The analysis of these variations presented in this report will also hopefully become a worth- while design aid.

Comparison of the overall performance of the two con- figuration concepts showed the E205 configuration to be superior.

The major limitation of the E205 concept has been un- covered and several suggestions for future research have been recommended to resolve the limitation and check the effect of the required configuration changes on the effects of the geometry variations described above.

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This document is presented in four volumes - Volume I - Study Overview, Volume II - Evaluation of Prediction Method- ologies, Volume III - Effects of Configuration Variations from Baseline E205 Configuration on Aerodynamic Characteris- tics, and Volume IV - RALS RI04 Aerodynamic Characteristics and Comparisons with E205 Configuration Aerodynamic Charact- eristics. The figures are placed at the end of each volume for the reader's convenience while the tables are integrated into the text as they occur.

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Q 2. INTRODUCTION Many potential advantages for incorporating VSTOL capa- bility into future Navy fighter / attack aircraft have been perceived by both the government and the aerospace industry.

Among the advantages are tactical benefits resulting from dispersal of air strength through operation from ships smaller than aircraft carriers , improved combat tactics via in-flight use of vertical-lift propulsive systems, reduced costs from requirments for construction of smaller ships, and improved close support through short takeoff and land- ing. Presently, the integration of a vertical-lift propul- sive system penalizes subsonic cruise performance and super- sonic dash capability , degrades the ship-board deck enviro- nment , and imposes additional operational requirements.

However , innovative aircraft design , including advances in propulsive system , flight control, structural, and aero- dynamic technologies projected to the 1990 time period, has led to the emergence of VSTOL concepts with significant transonic maneuver and supersonic performance potential.

Nevertheless, detailed configuration design of these VSTOL aircraft concepts is generally lacking, and only limited experimental data to define the aerodynamic / propulsive characteristics of such vehicles are available. Therefore, studies were commissioned jointly by the Navy (David Taylor Naval Ship Research and Development Center and Naval Air the to O Systems Command) and NASA Ames investigate aerodyna- mic technology associated with various VSTOL fighter / attack aircraft concepts.

In Phase I of the contracted program , (Reference I) four contractors provided conceptual designs, estimated the aerodynamics of the designs, identified aerodynamic uncertainties of the concepts and proposed a wind-tunnel pr o gram t o expl o re these uncertainties. In Phase II of the contracted program , two contractors designed and built wind-tunnel models for tests in the Ames Unitary and 12-Foot Wind Tunnels covering a Mach number range of 0.2 to 2.0.

This rep o rt presents the analysis of the testing accom- plished with two models designed and built in Phase II of the contracted program by the General Dynamics Corporation (Reference 1 and 2). This analysis was conducted by General Dynamics under a separate contract to NASA Ames under Con- tract No. NAS2-I0344. The two wind-tunnel models investi- gated in this report represent horizontal attitude takeoff and landing VSTOL fighter attack aircraft derivatives of General Dynamics' Vectored-Engine-Over-Wing (VEO-Wing) concept. This concept (see Figure 2-1) achieves improved

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transonic maneuvering and short takeoff and landing (STOL) performance by utilizing the full engine momentum from over-wing-mounted engines to augment the external aero- dynamics through a jet-flap effect and vortex augmentation.

The major difference between the two configurations is the propulsive system utilized for vertical lift. These propul- sive systems are the jet-diffuser ejector and the General Electric developed Remote Augmentation Lift System (RALS).

These systems represent the range of cold-vs-hot deck en- vironments currently being considered for vertical propul- sion concepts. Both systems afford thrust / lift augmenta- tion, which allows reduced vehicle size for a given payload capability. The aerodynamic lift augmentation achieved from the VEO-Wing nozzles through upper circulation also leads to reduced vehicle size. Three-view drawings and artist con- cept drawings of the E205 and RI04 airplane configurations are shown in Figures 2-2 through 2-5. The models repre- senting these configurations are described in Section 3.4.

The primary objectives of this analysis effort are: i. To evaluate the ability of current methodologies to accurately predict the aerodynamic characteristics identified as uncertainties for the two aircraft con- figurations developed in the Phase I study program described above.

2. To analyze the results of the three test entries to O determine the effects of configuration variations with the baseline ejector and RALS configurations.

3. To analyze the results of the three test entries to determine the differences in the aerodynamic characteristics of the ejector and RALS baseline configurations.

These objectives are accomplished in this report by tracing the development of the airplane configurations through the design process and noting the predicted aerodynamic characteristics used in the sizing studies for the E205 configuration and the resulting aerodynamic uncertainties that evolved. Following a description of the wind tunnel models and tests that were accomplished, an analysis of the test data yields an appraisal of current prediction methodo- logies to resolve the aerodynamic uncertainties. The analy- ses also aptly demonstrate the variation in aerodynamic characteristics resulting from the wide variations in configuration variables as well as the differences that result from variations in the vertical propulsion concept, i.e., RALS vs ejectors.

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O Comparisons of predicted and wind tunnel aerodynamic characteristics are limited to the E205 configuration only because (i) either one of the configurations could serve as a sufficient test case to determine the capabilities the prediction methods for this generic class of VEO configura- tions and (2) the E205 wind tunnel model more closely repre- sents the full scale E205 airplane configuration than does the wind tunnel model of the RALS RIO4 airplane configura- tion (see Section 3.4).

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3 . REVIEW OF A I RP L ANE CONF I GURAT I ON A

DEVELOPMENT STUDY w

The E205 and RI04 airplane configurations (Figures 2-2 and 2-3)were developed during the Reference I , Phase I, contracted study effort. This contracted study limited the scope of the analysis to only one concept, the jet-diffuser ejector concept, E205. This configuration concept was selected because it offered more potential shipboard opera- tional benefits due to its benign footprint that other members of the VEO-generic class of VSTOL fighter concepts.

Further , the ejector configuration exhibits more aero- dynamic uncertainties and differs more from the existing data base than does the RALS. The RALS configuration more closely resembles the VEO-Wing fighter for which an un- powered experimental data base already exists. General Dynamics continued to pursue the RALS configuration through in-house funding in a somewhat parallel study program; in fact, the RALS and ejector configuration were first compared using NAVAIR-supplied ground rules in Reference 3.

However, the E205 and RIO4 configurations were further developed using the sizing ground rules of Reference i. The Reference 1 study was structured to assess the import a nce of the various aerodynamic uncertainties involved in the con- cepts by actually designing and sizing the airplanes to a set of requirements suggested by the NASA contract guide- lines and by General Dynamics' experience in previous Navy VSTOL fighter studies. The requirements shown below, re- flect the desire for the aircraft to have good supersonic fighter combat performance (with reasonable mission "legs") when operating in VTO and good attack-support capability when operating in STOL: Mission: VTO Deck Launch Intercept (DLI) with (Standard Day) radius of action = 150 n.mi and design dash M = 1.6 (See Figure 3-1 for detailed mission profile definition).

C o mbat Performance: Sustained load factor of 6.2 at Mach (Standard Day) 0.6, i0,000 ft of altitude at 88% VTOL gross weight Specific excess power of 900 fps at ig, Mach 0.9, i0,000 ft of altitude at 88% VTOL gross weight.

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O VTOL: Vertical acceleration = 1.05 g (IGE) (Tropical Day) while achieving maximum design control rates simultaneously in all axes , where maximum design control acceleration rates are: Roll = .96 rad / sec Pitch = .28 rad / sec Yaw = .40 rad / sec STOL Operational from land and from ships smaller than CV's without catapults and arresting gear;sea-based gross weight = VTO maximum gross weight + i0,000 ib; sea-based WOD = 20 kt for overload.

VTO Takeoff Maximum = 35 , 000 lb.

Gross Weight Fuel Flow U se minimum engine without 5% fuel flow Conservatism conservatism (approximately same as using average engine with 5% fuel flow conservatism).

O It should be noted that the RALS 104 configuration, as drawn in Figure 2-3, does not represent an aircraft exactly sized to meet all of the ground rules described above but, instead , served as a sizing baseline configuration for synthesis studies. Comparisons between the ejector and RALS configurations sized to these ground rules are shown in Reference 1 based on synthesis studies.

The VEO-Wing Ejector and RALS fighter / attack configuration layouts have been influenced heavily by the necessity to meet the following three criteria simultaneously: i. Static margin variation (center of gravity /aerody- namic center) with Mach number from approximately -18% (unstable) subsonically to + 10% (stable) sup- ersonically. The maximum allowble instability of -18% static margin subsonically is a value set by the aerodynamic control and control-system-response capability expected in the 1990 time period.

2. Center-of-gravit y , wing, nozzle , and canard-loca- tion relationships (as well as static margin) to achieve the supercirculation benefits of the VEO- Wing concept for cruise maneuvering and STOL. To O achieve the benefits of the VEO-Wing concept re- quires the c.g. to be as far aft as possible (with i static margin as in Criterion I). This has a large impact on the configuration design; it is difficult to achieve the desired c.g. location without get- ting so much "real estate" ahead of the c.g. that the resulting forward-located aerodynamic center produces more instability than can be tolerated.

3. Center-of-Gravity, aircraft-inertias , and thruster locations that meet the hover requirements.

3.1 Airplane Configuration Descriptions This section provides a brief description of the physi- cal characteristics and design features of the E205 and RI04 airplane configurations. For a more detailed description of these configurations , the reader is referred to References 1 and 3 which deal with the following design features in some detail and which are omitted from this report: geometry, propulsion systems, mass properties , structural design, flight controls and the major subsystems such as avionics, crew station equipment , secondary power generation and weapons, electrical, hydraulic, ECS , oxygen and fuel systems, etc.

Both the E205 and RI04 supersonic fighter designs are configured to provide propulsive enhancement of external A

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aerodynamics. This unique integrated airframe / propulsion system, known as VEO-Wing (Vectored Engine Over Wing, see Figure 2-1) utilizes the full engine momentum from the over-wing-mounted engines to augment the external aerodynamics through a jet-flap effect. The VEO-Wing feature is combined with spanwise blowing in which a portion of the engine exhaust is used at high angles of attack to produce leading-edge vortex augmentation. This unique system is thus capable of providing lift / drag polar improvements in the full angle of attack range, resulting in improved maneuverability and STOL performance.

3.1.1 Ejector E205 Airplane Configuration The three view drawing of the jet-diffuser ejector VSTOL fighter / attack conceptual design (E205) sized to meet the mission, hover and combat performance requirements described in Section 3.0 is presented in Figure 2-2.

The concept utilizes a high-canard, low-wing arrange- ment with podded engines located for over-the-wing blowing.

Four chord-wise bays between the center body and nacelles (two forward and two aft) are provided for location of the jet diffuser ejectors.

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Unique fea t ures of th is conf i guration approach are the incorporation of movable doors to form the ejector nozzle and the stowable pri m ary nozzles , which result in a rela- tively compact arrangement when the ejectors are not in use.

A strake is extended forward and a beaver tail aft to fair off the depth of the ejectors when folded into their cruise position. The ejector design is based on application of the research discussed in detail in Reference I. The ejectors are sized to be operated with intermediate power airflow from P&WA .35-25-2800 parametric engines. Air is diverted to the ejector primary and throat nozzles through the duct- ing arrangement shown in Figure 2-2. The augmentation ratio is 1.98 in free air and 1.70 at lift off. Thrust modulation at the forward and aft ejectors , by varying airflow at the ejector primary nozzles , is used for pitch control during hover and transition. Yaw control is achieved by vectoring the ejector flow. Engine exhaust air is ducted to upward and downward firing thrusters for roll control. No lift is produced by this reaction control system in hover. The VEO-Wing engine nozzles can be operated in afterburner power setting with the ejectors running.

VTOL transitions are accomplished by diverting the excess thrust required to hover out of ground effect from the ejectors to the vectorable VEO-Wing nozzles. For STOL are aero- O operations the hover controls blended with the dynamic controls (the canard, elevons, and all-moving verti- cal tail) to provide control about the pitch, roll , and yaw axes. T he reaction controls are fired fore and aft or up and down as required to provide yaw and roll control at very low speeds or high angles of attack to augment the aerodynamic controls.

During conventional flight, control about the three axes is provided with canards , elevons , and the VEO-Wing nozzle for pitch, flaperons for roll, and the vertical tail and flaperon for yaw. The reaction controls are also available at low speeds for augmenting the aerodynamic c o ntrols to extend the lateral-directional c o ntr o l capabilities at high angles of attack. Due to the high longitudinal instability levels that result with this type of configuration, the flight control system is used to schedule the canard , VEO-Wing nozzle, and wing flaps as a function of Mach number, angle of attack, and power setting to achieve desired levels of static longitudinal stability and to augment the longitudinal stability to the required frequency and damping levels.

Dimensional and pertinent design data for the E205 con- figuration are presented in Table 3-1. The wetted area component buildup is included. The cross sectional area

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TABLE 3-1 DIMENSIONAL AND DESIGN DATA W FOR EJECTOR E205 CONFIGURATION WING Area (Ref) 384 ft2 (35.67 m 2) Aspect Ratio 3.62 Taper Ratio .19 b 37.28 ft (11.36 m) • b / 2 223.70 in. (5.682 m) CR 207.72 in. (5.276 m) CT 39.47 in. (1.003 m) c 142.680 in. (3.624 m) y 86.473 in. (2.196 m) Airfoil Root & Tip NACA 64A204 Sweep-Leading-Edge 40 ° Sweep - c / 4 32 ° Incidence -0 ° Dihedral 0 ° CANARD Area (Exp) 76.9 ft2 (7.14 m 2) Aspect Ratio 2.16 Taper Ratio .37 b (Tip to Tip) 28.6 ft (8.72 m)

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b / 2 (Exp) 77.33 in. (1.96 m) CR (Exp) 104.58 in. (2.655 m) CT 38.67 in. (.982 m) c 76.65 in. (1.947 m) Y 32.74 in. (.832 m) Airfoil Root NACA 64A005 Airfoil Tip NACA 64A003 Sweep-Leading-Edge 45 ° Sweep - c / 4 37 ° Incidence & Dihedral 0 ° Lc (LE _ Wing to _ / 4 Canard) 6.7 ft (2.04 m) Vc (Volume) 514.3 ft3 (14.565 m 3)

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i0 Q TABLE 3- 1 DIMENSIONAL AND DESIGN DATA FOR EJECTOR E20 5 CONFIGURATION (Continued) VERTICAL TAIL Area (Exp) 47.5 ft2 (4.41 m 2) Aspect Ratio 1.27 Taper Rati o .43 b 7.8 ft (2.38 m) CR 102.6 in. (2.606 m) CT 44.1 in. (1.120 m) c 77.3 in. (1.963 m) Y 40 in. (1.016 m) Airfoil Root 5.3% Biconvex Airfoil Tip 4% Biconvex Sweep - Leading-Edge 47.5 ° LVT (LE c Wing to c / 4 VT 17.8 ft (5.43 m) VVT (Volume) 845.1 ft 3 (23.933 m 3 ) WETTED AREAS Fuselage 451 ft2 (41.90 m 2) Canopy 33 ft2 (3.07 m 2) Nacelle 461 ft2 (42.83 m2) Wing 919.5 ft_ (85.42 m 2) f t2 O Canard 153.8 (14.29 m 2) Vertical Tail 95 ft2 (8.83 m 2) Dorsal 8 ft 2 (.74 m2) Wing Aft of Nac 27.7 ft 2 (2.57 m2) TOTAL 2159 ft 2 (199.64 m 2) Fineness Ratio (l / de) 7.66 Fuel Fraction 27.2 % Structural Fraction (w / o Ejectors) 32.4% Composites (% of Struct Wt) (w / o Landing Gear) 23.1% Advanced Metallics (Incl Ejectors) (% of Stru c Wt w / o Landing Gear) 72.8% C G Location ( % c) 3.0% VTO TOGW / Max TOGW 34987 / 44987 (15853 / 20384 kg) Combat Wt (88% VTOGW) 30789 ib (13950 kg) Flight Design Wt (88% VTOGW ) 30789 ib (13950 kg) Empty Wt 23402 ib (10603 kg) Payload (VTO / Max Overlead) 1146 / 11 , 146 ib (520 / 5055 kg) I n stalled Gun Sys. Weight / Am m o 521 / 500 ib (236 / 227 kg) Avionic Wt (Installed / Uninstalled) 1057 / 846 Ib (479 / 384 kg)

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TABLE 3-1 DIMENSIONAL AND DE S IGN DATA A

FOR EJECTOR E205 CONFIGURATION w

(Continued) Internal Fuel Volume Fuselage (Bladder) 877 ib (3974 kg) Wing (Integral, Halon Inerted) 750 ib (340 kg) Total 9521 ib (4314 kg) Design Mission Fuel 9521 ib (4314 kg) Number of Engines & Types (2) P&W Parametric Eng FB ABTF BPR=.362 OPR=25 TIT=2800 ° F (1537.8°C) Thrust (Max A / B SLS- U nistalled Each) 22,718 ib (10 , 294 kg) Inlet Type Axisymmetric Normal Shock Shock A1 Per Engine 4.86 ft2 (.451 m 2) W / S At VTOGW 91 ib / ft 2 (444 kg / m 2) T / W At VTOGW (Max A / B SLS U ninstalled Thrust) 1.3 Max Cross Section Area Minus A1 33.2 ft2 (3.084 m 2) Airplane Overall Dimensions Overall Length 53.3 ft (16.25 m) Overall Span (Including Missiles) 39.4 ft (12.0 m) Overall Height 15.4 ft (4.69 m) Flight Design Limit Load Factor 7.5 g Design Rate of Sink 15 fps (4.57 mps)

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c.g.

O distribution is shown in Figure 3-2. The is maintained by fuel burn sequencing at +.03c (F.S. 308.86) as long as possible to achieve t he VEO benefits for combat (until about 3000 of the 9521 ib of fuel for the DLI mission remains).

The control devices and deflection limits are as follows: Max deflection i. VEO-Wing nozzle -I0 ° to +30 ° 2. Flaperon (outboard of VEO-Wing -20 ° to +30 ° nozzle) 3. Canard -25 ° to +25 ° 4. Reaction controls -90 ° and +90 ° 5. All-moving vertical tail -25 ° to +25 ° The flaperon acts with the VEO-Wing nozzle for high lift but also acts as an aileron from the deflected flap position to provide roll control.

3.1.2 RALS RI04 Airplane Configuration O The RALS RI04 configuration (Figure 2-3) utilizes a high-canard, low-wing arrangement with podded engines l o cated for over-the-wing blowing like the E205 configura- tion. The RALS concept provides a vectorable force forward of the aircraft c.g. by augmenting fan-discharge air (with a burner) from the GE 16 / VFI9-DI (.6 bypass ratio) variable- cycle engine in a remote duct buring system. A "three poster" configuration is thus achieved for vertical flight by vectoring the reheated VEO Wing exhaust nozzles 90 degrees downward (in conjunction with the RALS nozzles).

The nominal exhaust tempertaure of both the remote and primary nozzles is 2800 ° F. Thrust modulation of the RALS and VEO-Wing nozzle burner is used for pitch control during hover and transition which results in temperatures up to 3200 ° F. Yaw control is provided by vectoring of the VEO-Wing nozzle. Roll control is achieved with downward- firing reaction control thrusters , which are located in the wing tips and always contribute to lift. Transition from hover to wingborne flight is accomplished by gradually diverting the thrust from the forward RALS to the aft VEO-Wing nozzles as wingborne flight is approached. To achieve the VEO-Wing benefits for up-and-away flight , the c.g. is held as far aft as possible (c.g. = +i.1%c).

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T h i s co n f iguration (like E 20 5) also f eatures spanwise blowing louvers (upstream of the VEO-Wing nozzle burner rings), which exhaust engine thrust out over the wing (Figure 2-3) to augment the leading vortex, thus delaying stall to higher angles of attack and producing excellent STOL performance (although the effects of spanwise blowing are not incorporated in this report).

For operations at low speeds, the configuration features an all-moving verticl tail.

The RALS nozzles are gimballed to provide 15 degrees of deflection from 90 degrees downward and 360 degrees in the planform view , except straight aft, where a maximum deflec- tion of 30 degrees is possible. These RALS nozzles not only are used to achieve vertical transitions or hover by pro- viding pitch and yaw control, but also are used as forward thrusters to achieve nose-wheel rotation at very low flight speeds t o provide excellent STOL performance.

The control devices for up-and-away flight and their deflecti o n limits are as follows: Max Deflection i. VEO-Wing nozzles -i0 ° to +90 ° 2. Flaperon (outboard of VEO- -20 ° to +30 ° O Wing nozzle) 3. Canard -25 ° to +25 ° 4. Reaction controls -90° and +90 ° 5. All-moving vertical tail -25 ° to +25 ° Just as on the E205 , the flaperon not only acts with the VEO-Wing nozzle for high lift, but also acts as an aileron from the deflected flap position to provide roll control in the transition , STOL or conventional flight modes.

Like E205, IR-guided missiles are carried on the wing tip; all other payload on RI04 is carried on the nacelles and the wide flat fuselage between the nacelles.

Dimensional and pertinent design data for the RI04 configuration are presented in Table 3-2. A cross-sectional area distribution is shown in Figure 3-3; the large volumes (and consequently large cross-sectional areas) required to install the RALS ducts in the fuselage plus the large engines required for V / STOL operation result in the high peak in the are distribution. The wetted area component build-up is also shown in Table 3-2.

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O TABLE 3-2 RI04 DIMENSIONAL AND DESIGN DATA WING Area (Ref) 300 Ft 2 (27.867m 2) Aspect Ratio 3.6 Taper Ratio .20 b 32.9 Ft (i0.028m) b / 2 197.2 In (5.009m) CR 182.5 In (4636m) Ct 36.5 In (.927m) c 126 In (3.200m) Y 75.7 In (1.923m) t / c Root & Tip NACA 64A204 Sweep-Leading-Edge 40 ° Sweep - c / 4 33 ° Incidence & Dihedral 0 ° CANARD Area (Exp) 66 Ft2 (6.131m 2) Aspect Ratio 2.16 Taper Ratio .37 b (Tip to Tip) 25.4 Ft (7.742m) O b / 2 (Exp) 71.7 In (1.821m) CR (Exp) 96.5 In (2.451m) CT 36 In (.914m) c 71 In (1.803m) Y 32 In (.813m) t / C Root NACA 64A005 t / C Tip NACA 64A003 Sweep-Leading-Edge 45 ° Sweep - c / 4 37° Incidence & Dihedral 0 ° Lc (LE c / 4 Wing to c / 4 Canard) 6.5 Ft (1.981m) Vc 429 Ft3 (12.148m3) VERTICA L TAIL Area (Exp) 43 Ft2 (3.995m2) Aspect Ratio 1.25 Taper Ratio .43 b 7.3 ft (2.225m) CR 98.5 Ft (30.023m) CT 42.34 In (12.91m) c 74.7 In (22.769m) Y 38 In (.965m) t / c Root 5.3% Biconvex t / c Tip 4% Biconvex Sweep - Leading Edge 47.5 °

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TABLE 3 -2 RI04 DIMENS I ONAL AND DESIGN DATA (Continued) LVT (LE _ Wing to _ / 4 VT) 16.08 Ft (4.901m) VVT 691.6 Ft3 (19.584m3) WETTED AREAS Fuselage 479 Ft2 (44.5m 2) Canopy 42 Ft2 (3.90m2) Na c elle 4 73 Ft 2 (43.94m 2 ) Wing 338 Ft2 Canard 132 Ft2 (12.26m 2) Vertical Tail 86 Ft 2 (7.99m2) Dorsal 6 _t2 (.557m 2) Wing Aft of Nac 28 Ft2 (2.60m 2) TOTAL 1584 Ft2 (147.158m 2) FinenessRatio (l / de) 7.12 Fuel Fraction 34% Structural Fraction 28.7% Composites (% o f Stru c t Wt) (W / O Landing Gear) 26.1% Advanced Metallics (% of Struct Wt g W / O Landing Gear) 64.3% C.G. Location (% c) 0.6% VTO TOGW / Max TOGW 31,940 / 41,94 0 (14,485 Kg / 19 0 20c Combat Wt (TOGW-40% WF) 27 , 590 Lb (12512 Kg) Flight Design Wt (TOGW- 40% WF) 27 , 59 0 Lb (12512 Kg) Empty Wt 19,000 Lb Payload (VTO / Max Overload) 1146 / ii , 146 Lb (520 Kg / 5055 Kg) Installed Gun Sys. Weight / Ammo 521 / 5 0 0 Lb (236 Kg / 226 Kg) Avionic Wt (Installed / Uninstalled) 1057 / 846 Lb (479 Kg) Internal Fuel Volume Fuselage (Bladder) 10,315 Lb (4678 Kg) Wing (Integtral, Halon Inerted) 560 Lb (254 Kg) Total 10 , 875 Lb (4932 Kg) Number of Engines & Types (2) GEI6 / VVCE 5 Study D 2 Engines With Remote Aug Lift Sys.

Thrust (SLS- U ninstalled Each) 21,747 Lbs (9863 Kg) 0, I TABLE 3-2 RI04 DIMENSIONAL AND DESIGN DATA ( Con t inued) Inlet Type Axi-Symmetgric Normal Shock A1 Per Engine 4.55 Ft 2 (1.387m2) W / S At TOGW 106.5 Lb / Ft 2 (520 Kg / m 2) T / W At TOGW (SLS Uninst a lled) 1.36 Max Cross Section Area 38 Ft 2 (3.530m2) Airplane Overall Dimensions Overall Length 48.8 Ft (14.874m) Overall Span 32.9 Ft (i0.028m) Overall Height 11.54 Ft (3.517m) Flight Design Limi t Load F a ctor 7.5 g Design Rate of Sink 15 fps (4.572m / sec)

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L i ke t he E 20 5 c o nfigurat i on , t he s t ructure of the RI04 O configuration is designed for a limit load factor of 7.5 g's and a design sink speed of 15 ft / sec. Both aircraft employ advanced metals and composites as well as advances in avionics to achieve weight savings considered feasible by the 1995 time period.

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O 3.2 Predicted Aerod ynamic Character istics One of the primary objectives of this analysis effort is to determine the capabilities of current prediction methodologies to predict the aerodynamic characteristics of the E205 type of configuration. In Reference 1 predictions of the longitudinal and lateral-directional aerod y namic characteristics of the full-scale E205 airplane configur- ation were presented for the subsonic (STOL / VTOL , M <.3) and transonic (cruise / maneuver, .3 <M <i.0) and supersonic (dash, M >i.0) flight regimes of the design DLI mission (de- fined in Section 3.0). These full scale aerodynamic predi- ctions are briefly reviewed in this section because they form the basis for developing the predicted.'m 6 del-scale wind-tunnel aerodynamic characteristics that are compared in Volume II with the actual wind-tunnel data to determine the prediction-methodology capabilities for the E205 type of configuration. Predictions of the full-scale E205 longi- tudinal aerodynamic characteristics including the lift, drag, and pitching moment curves, aerodynamic center travel and buffet onset angle of attack are presented in this sec- tion as well as estimates of some of the lateral-directional characteristics including the rigid sideslip derivatives, vertical tail effectiveness and aileron effectiveness.

The predicted longitudinal aerodynamics of the full-scale E205 aircraft configuration are based on minimum while the Q estimated values of drag lift , pitching moment , and drag due-to-lift rely heavily on the experimental da_a base developed from wind-tunnel tests of the powered GeneralDynamics Research Model (Ref.4) and the unpowered VEO-Wing fighter model (Figure 3-4) (Ref. 5). Figure 3-5 illustrates the forces and moments considered in predicting the full scale E205 aircraft aerodynamic characteristics. Table 3-3 provides a summary of analysis schemes with other pertinent data included for each flight regime. Three types of data coefficients were developed in Reference 1 for analyzing the configuration in these flight regimes and are defined by the following equations: i. Total Coefficients (Subscript t): All aerodynamic plus thrust forces included = CTv.N. CTEj Ejector CL t CLAero + sin (_F +_ ) + cos _ - CDram sin = CT cos_(6 F +(_) + CTE J sins+ CDEng +CDram cos (x CDt CDAero -Inlet Ejector Ram Drag

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Table 3-3 Methods Summary FLIGHT PROPULSION VEO NOZZLE THRUST DATA TYPE REGIME SYSTEM EMPLOYED ANGLE _GNITUDE BOOKKEEPING AND TRIM METHODS PRESENTED STOL, VTOL THRUST VECT FROM VEO 6TE = 30 ° CT _ 7.4 GROSS THRUST APPLIED / REMOVED AERO M _.3 NOZZLE + EJECTOR IN (+_ TE ) DIRECTIONeNOZZLE EQUIVALEN_ THRUST + SPANWISE & INLET FORCES APPLIED TOTAL BLOWING (NOT USED EXTERNALLY TRIM w FLAPERON, IN REPORTED CANARD, EJECTORS, VEO WING ANALYSIS BUT NOZZLE POSSIBLE) o CRUISE / THRUST VECT FROM 6TE = 15° CT !.2 NET THRUST APPLIED / REMOVED EQUIVALENT MANEUVER VEO NOZZLE IN DIRECTION INDUCED & .3_MXI.0 VECTORING EFFECTS INCLUDED IN "EQUIVALENT" POLAR NOZZLE & INLET FORCES IN PROP. TRIM WITH FLAPERON, CANARD, VEO WING NOZZLE DASH NO THRUST 8TE= _ CT < .15 CONVENTIONAL A / C METHODS EQUIVALENT M>I.O VECTORING (USE UNPOWERED DATA) NO SPANWISE TRIM WITH CANARD BLOWING

• ' 0 •

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_t = _Aero + CT cos _F ( C.G.W.L. - _ V.N.W.L. ) C sin 6 F (C.G.F.S. - V.N.F.S.)

+ CTv'N" c

+ ( C.G.F.S.- EJ.F.S.)+ CDEng• sins(C.G.F.S.-Inlet F.S.)

CTEj _ -Inlet - c c Where C.G.W.L. = waterline for center'of-gravity location V.N.W.L. = waterline of VEO-Wing nozzle thrust vector C.G.F.S. = fuselage station for center-of-gravity location V.N.F.S. = fuselage station of VEO-Wing nozzle thrust vector O EJ.F.S. = fuselage station of ejector thrust vector.

Assuming ejector thrust always 90 ° to W.L., i.e. no thrust recovery• 2. Equivalent Coefficients (Subscript E): Aerodynamic plus thrust forces with thrust angle-of-attack effects removed are: sins cosS + sin_ C_ = C_ - CTv.N" - CTE J CDram Ejector + C cos_ sins- cos_ CDE-" CDT TV.N. - C;rEj CDrnm Ejector CME= CMT - CTv _ - CTE J • . .- _ " N. (C.G.W.L.-V.N.W.L. i (C G.F S EJ.F S.)

• C C

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3 . Aerod ynam ic-Only Coefficients (Subscript Aer o) : Longitudinal force and moment coefficients with all

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thrust effects removed: _ero DAero ro The coefficients were de- veloped by app l y i ng corrections (for super-circulation , VEO-Wing nozzle deflection, and canard deflections , dreived from the General Dynamics Research model plus differences between the AFFDL VEO-Wing fighter model and E205 geometry) to the unpowered VEO-Wing fighter model wing-body data of Reference 4.

A detailed set of equations for developing the aero-only coefficients for the E205 configuration in the STO / VTO flight regime is presented in Figure 3-6.

The predicted trimmed lift,drag and pitching moment curves for the subsonic, transonic and supersonic flight regimes for the full scale E205 aircraft are presented from Reference 1 in Figures 3-12 through 3-17 ; power effects are included where appropriate. In the subsonic regime, the trimmed lift and drag curves differ fr o m those published in Re f erence 1 due to an error discovered in previous calculations.

Table 3-4 presents the estimated minimum trimmed drag component-buildup for the full shale E205 configuration from .2 <M <2.0 with canard, VEO-Wing nozzle , and flaperons at a i zero degrees deflection. Estimated trimmed minimum drag is plotted versus Mach number in Figure 3-7. The subsonic and supersonic friction_ form , wing camber, and interference drag were estimated by an empirical aircraft aerodynamic prediction method developed by General Dynamics for AFFDL (Reference 6). The supersonic wave drag was estimated by a modified version of the Harris area rule procedure. The roughness + protuberance drag was estimated at 18% of the friction plus form drag subsonically and 28% of friction drag supersonically based on F-16 flight test experience.

This trimmed drag also includes increments for flap scrub drag, and installed missiles and launchers for the DLI mission (2 Low-Cost Lightweight missiles and 2 Advanced Medium Range Air-to-Air Missiles + Launchers).

The STOL / VTOL longitudinal aerodynamics were estimated by adding increments in lift , drag , and pitching moment to the wing-body fighter model data of Reference 4 (Figure 3-8) according to the equations described above and in Figure 3-6. These increments , derived from the powered research model of Reference 2 (for canard deflection , wing trailing edge flap and VEO-Wing nozzle deflection as well as thrust level) , were corrected, where applicable, for geometry differences between the powered research and fighter

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Table 3-4 E 205 MINI_flJM DRAG BUILDUP Sre f = 384 ft2 - MACH NUMBER DRAG COMPONENT .2 .4 .6 .8 .9 I.2 i.6 2.0 (Drag in Counts) Friction 166.5 149.3 139.0 130.4 126.5 116.0 103.0 90.8 Form 17.2 15.5 14.2 13.4 13. i - - - Interference 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 Roughness + Protuberance 25.9 25.9 25.9 25.9 25.9 32.5 28.8 25.4 I iFlap Scrub 32.7 10.9 5.5 4.4 4.4 2.2 1.1 1.1 'Wave ..... 292.3 289.4 281.4 Missiles + Launchers (1) Wing-Tip LCLM 7.7 7.7 7.7 7.9 8.7 16.1 14.3 12.2 (2) NAC-MT'D AMRRAM 7.7 7.7 7.7 7.9 8.9 13.5 ii.0 7.5 , Trim 0 0 0 0 0 0 8 8 h _ CDmi n ITotal 268 226 213 213 216 482 466 441 models and the E205 configuration. Similar corrections were also made to the wing-body fighter model data to account for geometry differences with the E205 configuration. Figures 3-9 through 3-12 present the resulting predicted full scale E205 STOL / VTOL power-off and power-on lift , drag, and pitching moment curves. Figures 3-9, 3-10, and 3-11 have been lifted from Volume II to illustrate the predicted wind tunnel-model low speed (M=.2) lift , drag and pitching moment characteristics for variations in canard and wing trailing- edge flap deflection. The full scale airplane predictions (power-off) would differ from these curves only by a scale effects correction to minimum drag. Figure 3-12 illustrates the envelope of trimmed lift curves and drag polars that are obtained when the power effects are included with the un- powered data of Figures 3-9 through 3-11 (corrected to full scale airplane). Figure 3-12 demonstrates that by using the forward ejectors in conjunction with the vectored VEO-Wing nozzles, virtually any reasonable angle-of-attack range can be achieved for STO / VTOL operations. This is discussed in more detail in Volume II.

The full-scale airplane cruise / maneuver (transonic) aero data estimates presented in Figure 3-13 required an alternate approach from the STOL / VTOL data estimates. The limited existing data base prevented the development of aerodynamic estimates for variations in all of the desired parameter combinations (canard deflection , VEO-Wing nozzle

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deflection, flaperon deflection , C_, and Mach number). Out of necessity an alternate approach was sought, which led directly to representative estimates of the trimmed cruise maneuver drag polars without developing the untrimmed data as follows. Figure 3-14 schematically illustrates how a set of equivalent trimmed, optimum-span-efficiency (e) envelopes (vs C_ and Mach number) were developed from the powered VEO-Wing nozzle.only for trim with a zero-degree canard deflection, maxlmum negative static margin = -18% at M = .2 , and c.g. = +.03c (like Configuration E205).

For a given Mach No., VEO-Wing nozzle CB, and with canard undeflected, the equivalent wing span efficiency is derived and plotted as a function VEO-Wing nozzle deflec- tion, 6TE, and equivalent lift coefficient in Figure 3-15.

The _TE required to trim (with undeflected canard) at various angles of attack and equivalent lift coefficient is determined from the equivalent lift and pitching moment curves and allows the determination of the optimum trimmed span efficiency envelope as a function of equivalent lift coefficient.

However, since the estimated static margin for _c = 0 ° is more unstable than allowable for the E205 configuration (as explained in subsequent paragraphs) , the flight control computer schedules the canard with Mach No. and angle of i attack to achieve the desired stability level. Therefore, a of O reoptimization the canard / VEO-wing-nozzle deflections would be required at each Mach No. and blowing-momentum-co- efficient combination to achieve the maximum obtainable e-envelopes. Since the existing data base was inadequate for developing these max-obtainable e-envelopes , the e-envelopes using the VEO-Wing nozzle only for trim were used in making these predictions. Although the e's are not necessarily the optimum achievable with the canard / VEO-Wing nozzle trim , they are considered representative of what can be achieved with canard / nozzle deflection combinations given enough experimental data.

The C L at Cilni n is a fallout of the way the e's are de- rived; these e's are used directly with the estimated mini- mum drags to produce the resulting cruise / maneuver equiva- lent trimmed drag polars shown in Figure 3-13. (The minimum-drag trim penalty was negligible based on the powered model data.) This approach does not afford the development and visibility of the untrimmed lift, drag , and pitching moment curves directly because this would require enough experimental data to determine the canard / VEO-Wing nozzle deflection schedule with angle of attack, Mach No., and C T (or C_).

The estimated supersonic (M = 1.2 and 1.6) lift,drag, and pitching moment curves were developed for the E205 O onfiguration by correcting the unpowered VEO-Wing fighter model data (zero degrees VEO-Wing nozzle deflection) of Reference 4 for changes in CM o , canard arm , and reference areas. These data are presented in Figures 3-15 and 3-17 along with the trimmed lift curves and drag polars which are developed from these data and shown in Figure 3-13.

Aerod ynamic Center As discussed in Reference I, the aerodynamic center travel with Mach number plays a major role in the design of the VEO-Wing configurations.

Estimates of the E205 configuration aerodynamic-center travel with Mach no. have been made by use of the Carmichael Procedure (Reference 7) and the Datcom method (Reference 8).

Figure 3-19 presents a General Dynamics a.c. pre- diction accuracy correlation for the Carmichael procedure for various configurations, including the VEO-Wing fighter model of Reference 5. The correction vs Mach no. (Figure 3-19) indicated for the VEO-Wing fighter model was applied to the Carmichael predictions for the E205 configuration (a similar configuration) to produce the corrected Carmichael estimates , shown in Figure 3-18 , for a zero-degree canard deflection and with canard off. The Datcom estimate for canard at a zero-degree deflection for M = .4 is also shown

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for reference and shows a significant disparity between the W prediction methods.

It was thought to be very difficult t o predict the E205 a.c. with either of these existing methods because of the unusual aspects of the configuration: the wide, flat body with separated nacelles, the relatively blunt forward strake, etc. The methods do not lend themselves to this type of configuration. The configuration is being driven hard by the predicted instability levels. This is a major aerodynamic uncertainty that must be resolved with an experimental test program. The methods above predict the a.c. in the linear attached flow (low _) regions only; as non-linear effects are experienced at high _'s, the a.c.-variation prediction methods are less reliable and experimental data must be used as a guide. So many aspects of the design are dependent on these high-_ stability characteristics; a wind tunnel program must be conducted to develop and tune the E205 configuration with any confidence.

The E205 configuration is longitudinally statically un- stable to achieve the VEO-Wing nozzle benefits. As noted above , the predicted 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.

Buffet Onset The estimated buffet onset angle o f attack variation with Mach no., canard deflection and wing trailing-edge flap deflection are presented in Figure 3-20; these estimates do n o t include the effects of thrust deflection which are as yet unknown but are expected to be favorable. These estimates were determined from analysis of the axial force data of the VEO-Wing fighter configuration force model of Reference 5 using the methods of Reference 9.

Lateral-Directional Characteristics Figures 3-21a through 3-21e present the estimated static lateral-directional characteristics for the E205 configuration. The variation in the rigid sideslip deriva- tions , Cy R , C ,_ , and C_ , with angle of attack and Mach no. have _een estimated using Datcom procedures. The derivatives are determined for _'s from -2 ° to +2 ° . It should be noted that the wide forward fuselage fairing and

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predicting O nac e lles c o n t ribute to t he uncertainties in the static lateral-directional instability and the effects on sidewash. The variation of the directional characteristics is largely dependent on this unorthodox forebody loading , which is not easily predicted by standard methods. The dihedral effect, which is dependent on C_, will be greatly affected by the induced super-circulation lift. The lateral characteristics were also expected to be affected by the canard and canard deflections. Data in these figures has been predicted for the zero-canard-deflection case.

Directional control for configuration E205 is obtained with an all-movable vertical tail. Control effectiveness of this surface is presented in Figure 3-21d. Standard DATCOM methods for these predictions were used.

Lateral control for configuration E205 is obtained with ailerons located from immediately outboard of the VEO-Wing nozzle to approximately 85% semi-span. The predicted values of roll-control effectiveness are presented in Figure 3-21e.

The augmentation in rolling moment due to the VEO-Wing has not been included because of lack of available data.

Side force and yawing moments due to aileron deflection were not predicted. The yawing moment is caused by the pressure gradient against the side of the fuselage. There was not O enough experimental data available for correlation to any reliable prediction method.

3.3 Resulting Aerodynamic Uncertainties In Reference i, the critical aerodynamic uncertainties were identified as those aerodynamic parameters which had the greatest effect on the E205 design and performance and which could not be accurately predicted with confidence.

The power-off aerodynamic uncertainties identified were the prime targets for experimental investigation in the wind tunnel tests described in Section 3.4.

Both the ejector and RALS configurations have large , wide , flat fuselage / strake areas end-plated by nacelles with the primary lifting surfaces located outboard. The aero- dynamics for this type of configuration are difficult to predict with existing tools. The unpowered-aerodynamic-test data base approximates the RALS much better than the ejector configuration , but no powered data exists for this separated nacelle type configuration.

A reiteration follows of how these aerodynamic uncer- tainties are related to the VEO-Wing / ejector design. While the following discussion focuses primarily on the ejector configuration , E205 , many of the c omments apply to the in- O house RALS configuration due to the similarity in external arrangements.

o e - The optimum e envelopes that can be obtained _ith canard / VEO-Wing nozzle combinations for trim must be confirmed since the experimental data base does not exist to allow optimizing the canard / VEO- Wing nozzle dfelctions (especially at transonic speeds). The transonic maneuver performance (and subsequent aircraft sizing) presented in this study are predicated on the assumption that the envelope e's developed for the more stable VEO-Wing fighter config u ration (Reference 4) with trim provided by the VEO-Wing nozzle only (i.e., canard fixed) can be duplicated with the canard / VEO-Wing nozzle deflec- tions, which also provide effective augmented in- stability levels of -18%. The effect of canard location will be examined in this effort as well as the influence of the strake / inner-body region on aerodynamic center and the resulting trimmed e.

° Minimum Dra@ - Large volumes are required for inst- allation of the vertical-lift ejector system (ejector bays and ducting). As a result, increases in wave drag and friction drag are incurred compared to a c o nventional takeoff-and-landing (CTOL) configuration. Minimizing the impact of increased cross-sectional area on supersonic wave drag requires considerable experimental configuration tailoring. Integration of the wing / nacelle / strake must be examined experimentally to minimize interference drag at transonic speeds because it is very difficult to predict the flow field and subsequent interference drag between the fuselage body and the nacelles.

o Trim Dra@ - Trim-drag penalty is critical for transonic maneuvering and the supersonic dash. The effect of canard location and schedule optimization will be investigated, as well as the effect of a.c.

on trim drag.

o _L m_ x - Maximum lift coefficient is critical for transonic maneuvering, STOL, and VTOL transitions.

The usable CLma x is determined by both longitudinal pitching moment and lateral-directional control characteristics at high _ , which are virtually impossible to predict with the wide-bodied configuration being studied with power effects.

" Body /Wing Design for Cmo - During maneuver at transonic speeds, it isdesirable to camber the body / wing for a positive Cmo contribution to alleviate the nose-down pitching moments induced by the vectored over-wing nozzles. Without this Cmo

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O contribution, the (larger) positive canard deflec- tions required to trim degrade the configuration transonic maneuver capability. U nfortunately, the E205 cDnfiguration , with jet diffuser ejectors in- stalled in the fuselage, precludes the use of a de- sign body camber. (This restriction does not exist for the RALS configuration, which has body camber incorporated.) One possible approach to obtain the desired positive Cmo shift for the ejector con- figuration is to employ a fuselage beaver tail de- flected upward during maneuvers.

o Aerod ynamic Center Location - Power-off aerodynamic- center predictions versus Mach number were shown in for the ejector configuration (canard off and zero canard deflection). The estimates are based on the Carmichael method (Reference 7) , with a comparison point against the DATCOM method (Reference 8 ) at Mach 0.4 and zero canard deflection. There is a significant disparity between the predicted values for the two methods, which must be resolved during the testing.

It is very difficult to predict the aerodynamic center with either of these methods because of the O unusual aspects of the configuration, the wide flat body with separated nacelles , and the relatively blunt forward strake. Also , the variation of aero- dynamic center with power setting, angle of attack, and trailing-edge flap deflection is difficult to estimate with conventional methods.

° Buffet Characteristics - The close-coupled canard of the E205 configuration is expected to delay the buffet onset to higher angles of attack in much the sam_ way as does the forebody strake on the F-16 aircraft. However, data to substantiate this favorable effect is lacking for canard / wing con- figurations. An additional increase in angle of attack for buffet onset should result from power-on supercirculation effects. It is desirable to obtain root-bending-mo m ent (CORMS) strain-gage data during the wind tunnel tests to verify the estimated configuration buffet characteristics.

o Lateral-Directional Characteristics - Lateral-direc- tional characteristics determine the a max and re- sulting usable CL max , which contributes to airplane sizing for transonic maneuvering STO , and VTOL operations. The effectiveness of the all-moving vertical tail in preserving lateral-directional O control at high angles of attack when influenced by A the large flat body and strake of the E205 (or RALS) configurations is difficult to estimate.

3.4 Description of Wind Tunnel Models The 0.0939 scale wind tunnel model representing the E205 aircraft configuration is defined in Figures 3-22 and 3-23,and was constructed by General Dynamics and tested at AMES Research Center. The overall dimensions of the model are illustrated in the three view drawing in Figure 3-22 and Table 3-5 while the cross-sectional area distribution of the model is shown in Figure 3-24; note that the model cross-sectional area distribution differs somewhat from that of the airplane (Figure 3-2). The model scale was dictated by the Ames requirement that the model be sized to match the airplane engine maximum inlet airflow with that of the XM2R / CMAPS engine simulator being developed for future Ames investigations into propulsion / airframe interactions. The model is defined in detail in Reference 2.

A second wind tunnel model was constructed to investi- gate s o me aspects of the RALS RI04 configuration. The three view drawing of the RALS model is shown in Figure 3-25 while the cross sectional area distribution and photographs of the model are shown in Figures 3-26 and 3-27. This model was constructed by building a new center fuselage section (simulating the RALS fuselage lines aft of the canopy / nose A

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area); (the transition in lines from the E205 nose and canopy to the RALS fuselage lines resulted in an aerodynamically unfavorable concave depression which may be faired-out in future investigations). The E205 model wings, canards, vertical tail, nacelles , nose and canopy were used on the RALS model; the major changes are a narrower fuselage resulting in more closely spaced nacelles and a thicker strake, and a deeper, narrower vertical channel between the nacelles and fuselage sides. Fairings on the E205 nacelles were removed to more cl o sely simulate the nacelle shape of th RALS configuration. The use of the wing panels outboard of the nacelles on the RI04 configuration resulted in a theoretical wing area of 357 f_ instead of the 300 ft 2 of the RI04 aircraft.

Both the E205 and RI04 model configurations have several very useful geometric variables which allow tailor- ing the aerodynamic performance of the configurations and provide real insight into the mechanisms of the uncer- tainties described above. Table 3-6 lists these geometric variables and the ranges of associated deflections or locations for each. Both models allowed three longitudinal canard locations (baseline (C I ), forward (C 2 ), and aft (C 3 ) Figure 3-28a); the canards have leading and trailing-edge

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Table 3-5 SUMMARY OF CONSTANTS w PARAmeTER E-205 R-104 i 2 2 Fuselage Base Area, Ab 9.5991 in 9.2941 in 2 2 Nacelle Exit Area, Ae(per side) 4.8982 in 4.8982 in 2 2 Nacelle Inlet Area, Ai(per side) 6.1685 in 6.1685 in 2 2 Nacelle Plug Base Area, _(per side) 3.2484 in 3.2484 in Wing Reference Area, S 3.3858 ft2 3.1541 ft 2 Wing Reference Span, b 42.010 in 40.270 in Wing Reference Mean Aerodynamic Chord, E 13.398 in 12.973 in Longitudinal Transfer Distance, X 1.002 in 1.479 in O Lateral Transfer Distance, Y 0.000 in 0.000 in Vertical Transfer Distance, Z 0 700 in 0.419 in Moment Reference Center @ FS 29.002 29.463

13.050 13.050

BL 0.000 0.000 Balance Incidence Angle, i 0°00 ' 2°00 ' . m (positive incidence alrplane nose-up)

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TABLE 3-6 AVAILABLE COMPONENT DEFLECTIONS Component Deflection Wing Outboard Trailing-Edge Flap 0°00 ' - I0°13 , 20°24 ' 25°29 ' Wing Leading-Edge Flap 0° 15° 30 ° Wing Inboard Trailing-Edge Flap 0° I0° 20 ° '0 2.5° Canard Leading-Edge Flap 0° 15° Canard Trailing-Edge Flap 0° 20° All Moveable Horizontal Canard 0°

+I0 °

+20 °

All M o veable Vertical Tail 0°

5°

15°

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flaps (inboard and outboard segments); the vertical tail is all moving to several deflections. The E205 model has three strake variations (baseline S I ) mid sweep (S 2 ) anda low sweep (S 3 ), as shown in Figure 3-28b.

Both models are sting mounted through the base of the fuselage with forces and moments measured with a 2.5in. Task MK-XXA six component strain gage balance. The moment refer- ence center for both models was located longitudinally at 3-percent of the mean aerodynamic chord which represents fuselage station 29.002 on the E205 model and fuselage sta- tion 29.463 on the RI04 model along waterline 13.050. The model angle of attack is referenced to the wing reference chord plane which is along a waterline. Fuselage-sting cawity pressure from three static orifices (two located immediately aft of the balance and one located 0.5 inches forward of the fuselage base) and nacelle nozzle exit plug base pressure from static orifices located in each nozzle plug base were measured concurrently with the force data.

Nacelle nozzle exit static and total pressures were recorded during internal drag runs. Two pressure rakes , each con- taining 20 total pressure orifices were calibrated prior to the wind tunnel tests against known mass flows measured by ASME nozzles; these rakes were used to gather the total pressure data for the internal drag calculations while the e static pressures were obtained from the manifolded orifices located in the nozzle exit plug walls. Fluctuating wing bending moments are measured as a buffet indicator (RMS value) using Kulite diffused semiconductor four-arm strain gage sensor, type 5B-3-350-300-4, locatged at 40% chord and span station 9.600 on the E205 and 8.73 on the RI04 models.

3.5 Wind Tunnel Test Programs The wind tunnel models described in the previous sec- tion were tested once in each of the following NASA Ames Research Center wind tunnels: the 12-foot pressure tunnel , the llxll foot transonic tunnel, and the 9x7 foot supersonic tunnel. These are single return tunnels with controls that allow independent variation of Mach number, density, temp- erature and humidity. Tests were conducted at Mach numbers ranging from .2 to 2.0 at a constant Reynolds number of 9.84 x 10 6/ M (3.0 x 10G / ft). To evaluate Reynolds number effects, selected configurations were tested at additional Reynolds numbers for several Mach numbers. The angle of attack range was -5 ° to 90 ° in the 12-foot , -5 ° to 27 ° (maximum allowable loading) in the llxll foot , and -5 to 15 ° in the 9x7 foot tunnel. The angle of sideslip ranged from -4 to +8 ° .

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A The measured angles of attack and sideslip were corrected for wind-tunnel-flow misalignment and for balance and sting deflections caused by aerodynamic loads.

The measured axial forces have been adjusted to a con- dition corresponding to that of having free-stream static pressure acting on the fuselage cavity and on the base areas of the two nacelle choke plugs. The data in this report has also been adjusted for internal forces acting in the flow-through nacelles using the internal and normal forces derived from the series of runs employing the duct exits rakes in each nacelle.

To assure a turbulent boundary layer, transition grit strips were placed near the leading-edges of the wing, canard and vertical tail , around the fuselage nose and around the nacelle leading-edges.

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4.0 CONCLUSIONS AND RECOMMENDATIONS The initial objectives of this research effort have been accomplished. The capabilities of current prediction methodologies to evaluate configurations of the VEO-Wing- VSTOL generic family have been evaluated, the effects of geometry variations have been examined, and two versions of this generic family, the E205 and RI04 configurations, have been aerodynamically evaluated and compared. As a result of these evaluations and analysis additional research is recommended.

CAPABILITIES OF PREDICTION METHODS Because of the unusual geometry involved with the E205 and RI04 configurations there was substantial concern that current prediction methods would not do an adequate job on predicting many of the areas of "aerodynamic uncertainty" for the E205 configuration.

In summary, the minimum drag is predicted reasonably well in the subsonic and transonic speed regimeswith existing methods; there are some substantial problems with the predictions supersonically where the actual drag is substantially higher than predicted. Part of this discrep- ancy many be due to a failure to accurately predict the model sting interference effects and the excess interference suspected due to the model components.

The aerodynamic center variation with Mach number was reasonably well predicted for the E205 configuration with the Carmichael Woodward procedure. However, a failure to consult the Carmichael Woodward results regarding the ef- fective camber of the configuration resulted in the error in predicting C m which in turn resulted in larger trim penal- ties than anticipated as some speeds.

In fact, the agreement between the predicted and test untrimmed lift, drag, and pitching moment curves is rather good (except for the error in Cm ) with the canard and flap undeflected at most speeds. The°canard is a little more effective than predicted because of the higher upwash than expected caused by the E205 wing-body. The wing trailing- edge flaps were also found to be slightly more effective than predicted ; the flap effectiveness was also found to vary with canard deflection which was not accounted for in the original predictions. However, the flap effectiveness was predicted by ratioing data from a similar configuration using appropriate parameters. This continues to be a basically reliable method of predicting the flap effects.

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Trimming at low speeds was accomplished using com- i binations of wing trailing-edge flaps and deflected thrust from the VEO-Wing nozzles balanced by the modulated thrust fro m the forward ejectors. The canard was not used for trim but could be deflected for maneuvering from trim (gust response, etc.). Transonically and supersonically, leaving the canard undeflected and varying the wing trailing-edge flap deflection yielded trimmed drag polars that were almost as good as with an optimum combination of canard and flap indicating that it may be possible to justify fixing the canard and using only a canard trailing-edge flap for aiding maneuvering from trim. The trimmed wind tunnel data cor- rected to full scale and including the power effects (super- circulation and deflected thrust) of Reference 1 were compared with the predicted power-on trimmed full scale aircraft characteristics of Reference 1 on which the E205 performance was evaluated. In general the test results indicate better characteristics than predicted except for Cm at some transonic and supersonic speeds, which can largely be traced back to a failure to accurately predict Cmo, resulting in substantial trim penalties. The polar shapes are, for the most part, better than predicted indi- cated that the maneuvering cruise , and dash performance which played a substantial role in sizing the aircraft to the ground rules described in Section 2.0 can be achieved, especially if some redesign to provide a more acceptable Cmo variation is employed, i The buffet-onset angle of attack variation with Mach number developed from the test data (using the RMS value of the wing-root bending-moment strain gage output) indicates that the E205 buffet characteristics are better than predicted but do agree in trend with the predictions.

The Datcom prediction techniques for the lateral-direc- tional characteristics appear to adequately account f o r the configuration effects except for sidewash generated at angles of attack. Most prediction techniques are only concerned with the linear range of the parameters involved.

The sidewash determined from these tests indicates a need to examine the configuration to at least determine what com- ponents effect the sidewash using a procedure such as Carmichael Woodward since Datcom does not adequately handle the sidewash prediction at angle of attack.

EFFECTS OF GEOMETRY VARIATIONS A matrix of combinations of canard longitudinal loca- tion and strake shape were investigated for the E205 con- figuration. This investigation indicated that there are major "first order" effects for varying either canard location or strake shape , but the influence of the strake shape on the canard effectiveness or the canard location "second O on the changes produced by varying strake shape are order" for this type of configuration; i.e., in the preliminary design stage the effects of canard location and strake shape must be considered, but the mutual interference of the canard and strake is not that important to the type of nacelle-strake-canard arrangement exhibited on E205.

Canard location produced the expected changes with the m o st dra m atic change being the substantial variati o n in d.C.

caused largely by changes in the canard / wing interaction.

Also of major importance is the change in effective configu- ration camber produced by changing canard location resulting in C and subsequent trim variations. The mid-canard location of the baseline E205 configuration offered the best overall trimmed characteristics across the Mach range as well as the best lateral-directional characteristics.

Variations in strake shape were found to be a very effective means of tailoring the pitching moment curve at high angles of attack to avoid pitch-up. Strake effects become more pronounced with increasing angles of attack and Mach number. The S1 C1 combination offers the best overall pitch-trimmed characteristics over the Mach range tested.

At low speeds, the S1 C1 combination also exhibits the best overall lateral-directional characteristics; at transonic and supersonic Mach numbers , the limited amount of data O precluded determining the best canard / strake combination for lateral-directional characteristics.

COMPARISON OF E205 _D RI04 CHARACTERISTICS The configuration body-nacelle-strake arrangement does influence the performance of the canard and wing. Compari- son of the E205 and RI04 wind tunnel model results indicate that the wide, flat strake arrangement of the E205 configura- tion acts as an effective lifting surface inducing a substantial upwash on the canard and wing. This results in the E20 5 wing and canard each performing better (alone and in the presence of each other) on the E205 configuration than with the narrow strake arrangement on the RI04 configuration.

The untrimmed minimum drag of the E205 and RI04 configura- tions is about the same subsonically; the RI04 transonic drag rise is much more severe , followed by substantially higher supersonic drag. This increased drag of the RI04 is due to its larger maximum cross-sectionalarea, the higher interference drag due to the

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A narrow channel between the nacelle and fuselage,and the W adverse effects of the concavity formed by using the E205 nose and canopy with the RI04 fuselage.

The aerodynamic center variation with Mach number for the two configurations are similar. The addition of the canard makes approximately the s a me a.c. shift for both vehicles.

The flap effectiveness for the two configurations is almost identical.

Comparisons between the E205 and RI04 wind tunnel model (unpowered) trimmed polars indicate that at transonic Mach numbers the E205 has a better trimmed drag polar thanthe RI04 for CL's >.35, so for combat maneuvering the E205 looks superior; The RI04 appears slightly superior for transonic cruise at low CL'S. At M = 1.2 the E205 is superior at all CL'S primarily because of the lower minimum drag. Although the untrimmed minimum drag of the RI04 is higher than that of the E205 at all supersonic Mach numbers, the trimmed minimum drag however is less for the RI04 at Mach numbers < 2.0 primarily because of differences in Cmo of the two con- figurations. O The lateral-directional characteristics of the two configurations are very similar except at M = 1.2 where there is apparently a more favorable flow at the vertical tail for the RI04 than the E205.

The major deficiency of both the E205 and RI04 configurations appears to be the inability to trim to angles of attack higher than 6 to 8 degrees at low speeds with either the canard or wing trailing-edge flap with the power off. The large configuration instabilities at low speed require large negative canard deflections to trim with the power off; _c = -20o was the largest negative deflection tested and this was not adequate to trim over a reasonable - range. The canard appears stalled at a canard-local angle of attack of near -20 (in the presence of the wing), so that the resulting trim limit appears to be about 8°model angle of attack. Part of the problemis, of course, the fact that the wing i t self is stalling at about _ = 8° because it has been provided no leading-edge protection, the leading- edge flaps not being employed due to a lack of test time.

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The airpl a nes could be said to be balanced at the wrong point but movement of the c.g. forward would result in de- teriorating canard controlpower even more. Several areas of additional research are suggested below to investigate the "real" aerodynamic characteristics of the configuration if the wing is working properly and to solve the problem of limited trimmed o (- range at low speeds.

I. Test the existing E205 wind tunnel model at low speeds to develop an acceptable trimmed _- range by varying the wing leading flaps across the wing span in combination with the canard at various deflections (deflecting the canard to negative deflections _<-20°)plus varyin_ the strake shape.

The objective of this testing would be to allow the wing to work effectively to higher a's;of course, as the pitch-trim limit is raised, the next operational constraint is the _ - limit due to deteriorating lateral-directional control ( a _ 16°).

Hopefully, aiding the wing performance and working on the strake shape will have positive effects on increasing the lateral-directional _- limits. With the configuration working better at low speeds, higher Mach number improvements can be investigated. Also, the effects of the canard loca- l tion movements and trimmed characteristics should be rechecked from the current tests to see if the magnitudes have substantially changed (trends are expected to be the s a me but magnitudes will probably substantially change, e.g., "wing-alone" performance or "canard-alone").

2. Remove the canard and develop E205 wing-body characteristics that are acceptable to higher _'s at low speed in the same manner described in the preceding para- graph recalling that the strake shape as well as the wing leading-edge flaps are effective means of modifying pitching moment characteristics and avoiding pitch up. When this is accomplished, mount the canards on the top of the vertical tail to be used as horizontal tails and test to develop the trimmed characteristics across the Mach number range using combinations of horizontal tail and wing leading and trailing- edge flaps. Estimates show that the existing canards are an acceptable size and planform for use as horizontal tails on this configuration.

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3. Based on the results of the data analysis presented in this report, a change in canard and wing planform is suggested for the E205 configuration which should alleviate the problem of low speed, power-off limited trimmed a - range.

General Dynamics has had considerable experience with wind tunnel testing delta planform configurations - with and without horizontal canards. A recent ex a mple was the Convair Division's lift-plus-lift-cruise V / STOL program.

In general, optimum low-speed and transonic maneuvering performance was obtained with this configuration balanced sufficiently unstable, canard-on, so that trailing-edge- down elevon deflection was required for trim. Pitch re- covery control power at high angles of attack was obtained by unloading the canard (high authority actuators are required). Pure tailless configurations do not possess sufficient nose-down control authority through the trailing- edge surfaces to permit maneuvering flight at negative static margins much greater than 1-2%. For modest canard- wing area and span ratios, the size of the canard only determines the center-of-gravity position at which the con- figuration should be balanced. O Based on this Convair experience, the modifications to the wind tunnel model indicated in Figure 4-1 are recom- mended as a potential direction toward effecting improve- ments in Configuration E-205 balance. Specifically, the wing has been enlarged and moved aft, in order to position the wing-body a.c. slightly behind the 3% reference point.

The planform selected is based upon previous wind tunnel experience, and is expected to give linear pitching moment characteristics at operational angles of attack. (Some wing leading-edge treatment may be required at extremely high maneuvering angles of attack). At the same time, the canard has been decreased in size and moved forward to the forward pivot location (MS 21.036). Summary geometric data are noted on the figure.

Aerodynamic predictions have been made, utilizing readily available wing-body and canard surface methods, in conjunction with wind-tunnel extracted canard flow fields(E205data). These data are inset on Figure 4-1. The wind-tunnel Cmo has been included, although a portion of further configuration development efforts would be directed

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toward the elimination of this phenomenon. Note that a nominal static margin of -8.5% is indicated (M -- 0.2), and that the configuration is significantly more controllable in the normal range of lift coefficients.

Construction and testing of the new canard and wing planform arrangement is strongly suggested because (I) it can be accomplished with minimal cost utilizing the existing E205 model hardware (only a hew canard, canard deflection bracket, and wing panel outboard of the nacelle are required) and (2) it would allow the confirmation of the expected solution to the low speed trim problem.

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5.0 REFERENCES i. Lummus, J. R., Study of Aerodynamic TechnoloKv for a VSTOL Fighter /Attack Aircraft_ NASA CR-152128, May, 1978.

2. Walker, J. J., Model and Test Information Report, .0939-scale VSTOL Fighter /Attack Aircraft Wind Tunnel Force Model , General Dynamics Rept FZT-344, Dec., 1978.

3. Advanced VSTOL Fighter Attack Aircraft Preconcept Formulation Study, General Dynamics Fort Worth Division Report NAV-GD-006, March, 1978.

4. Woodrey, R. W., et al., An Experimental Investigation of a Vectored-Engine-Over-Wing Powered Lift Concept, AFFDL- TR-76-92, Volumes I and II, September, 1976, 5. Heim, E. R., Basic Aerodynamic Data for a Vectored- Engine-Over-Wing Configuration, AEDC-TR-78-1, February 1978.

6. Schemensky, R. T., Development of an Empirically Based Computer program to Predict the Aerodynamic Characteristics of Aircraft, AFFDL-TR-73-144, Vols. I and 2, November, 1973.

7. 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.

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

9. Ray, Edward J_, Techniques for Determining Buffet Onset, NASA TMS 2103, November, 1970.

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• HIGH LIFT COEFFICIENT S (Low Approach S peeds) AT TAKE-OFF / LANDING CONDITIONS 6. 0 5.0 ( _ ) 4.0 CL 3 , 0 2. 0 1 , VE0 1. 2. SPA N WlSE LO IN

1; 30

= DEG Q , • FULL ANGLE OF ATTACK POLAR IMP ROVEMENT AT TRAN S ONIC S PEED S -- MACH = . 9 1 2 .

• COMBINES SU P ERCIRCULATION VE O -Wl N G F R O M OV ER-WING - M O UNTE D \ =_ .8 ENGINES WITH LEADING - EDGE VORTEX AUG M E NT AT I O N "" I, - , ,,=1 • UT I LI ZE S F U LL E NG INE _ J . 4 MOMENTUM (No Bleed) 0 .04 . 08 .12 .16 2 .0 CDTRIM Figure 2-1 VEO-Wing Powered Lift Concept ____ ... _ -L ,,..

~~ ""') ......

" .Q""

-

Figure 2-2 E205 Three-View Drawing ~~a.-..~ ~TUDY VpTOL l:> r ON~IG 12104 -' .. - -:::.a;:::--I fWlO~OOl ..

Figure 2-3 RI04 Three-View Drawing Figure 2-4 Ejector Configuration in Cruise Mode Figure 2-5 Artist Concept of RALS R104 Configuration

MISSION DEFINITION

, M = OPT, h = OPT PRIMARY (SIZING) MISSION DECK·LAUNCHED INTERCEPT (DLI) '-=" h = 40K FT, M = 1.6

COMBAT: I

10-MIN LOITER 2-MIN @ MAX AlB + 5% INITIAL FUEL M=1.6 + 45 SEC VTO THRUST • LOADING: • (2) LCLM MISSILES • RADIUS • • (2) AMRAAM MISSILES WARMUP & T.O.: 2.0 MIN'@ INTERMEDIATE PWR + 30 SEC VTO THRUST 612201 Figure 3-1 DLI Mission Profile V/STOL B PROGRAM: j: h' • '"1 ....

. 'l

' .

I' . , , I· N 2: ex: w ~ a: l.O ex: - •...•. + 2: C I-

I

c.J I w .. " 2000 en en en c a: c.J .4·1· .... ·.

I'

. .. . .~

600 500 200 100

o - 60

FUSElAGE STATION·IN.

Figure 3-2 E205 Cross-Sectional Area Distribution 5000 r-----f------t--Ir..r:- .~~~_-:.~~,+----~-V.:..:£::..R::..T..:.T::..AI;.:l--1------1 4000 t------+-----.y-----+----~~~-----f-----+------1 VI ....

o !!

I C looO ...

a: c 2000 1__----~--1~c......---+--I----~--~---~-+---~.I------~~-------4 1000 I- +-I~----~t..------+-----+--\-----.p,,---~,....--I-------I OLL .::Ir;::;;.. __ .L.....M:::...:::::a... __ ~ ....... ~-......I-~~-...I_--~~~-----' o 100 200 JOO 400 600 fUSST,. -IN.

Figure J.J Cross·Sectional Area Distribution. R104 AFFDl VEO FIGHTER MODEL AEDC TEST GD/NASAIAFFDl RESEARCH MODEl "---.--.---- .. - - (Ret"4) - -_. _._-- (Ref. 5) • POWERED MODEL .. UNPOWERED flOW THROUGH MODEl Tested Subsonic to Supersonic Q' 's to 40 Subsonic • PROVIDES-Power Effects • PROVIDES-Power Off Baseline for VSTO l Aero Buildup -Canard Effects (Subsonic) -Supersonic Trimmed Polar Shapes -Transonic e's FOR THIS STUDY FOR THIS STUDY VEO-Wing Experimental Data Base Figure 3-4 T FWD EJECTOR \JI N

.-----

--

----.-

.--------

--

..------

- -~-_.

----------_ ...

Figure 3-5 Longitudinal Forces Acting on E205

0 0 •

A E R O - O N LY COEFF I C I ENTS BU I LDUP SREFTEST _ S CE205 SREFTEsT 6 SWE2 05 SREFTEST C LAERO" CLWBTEST " SWE205 . SREFTEST + _ C LcHANGE + C L c ANARD + CLFLAP • • SWTES T SREFE205 SREFE20 5 SCTEST' SREFE20 5 SWTEST SREFE205 L r l . + SREFTE s T . J_t SCE205 SREFTEsT _ C DF LA P SNE 205 SREFTEsT SWE205 SREFTEsT _ C D c HANG E C D c ANARD + Go CDA E R O" CD N BTEST S WTE ST S REF E20 5 S R E FE2 0 5 SCTEST S REF E 25 0 S W T E sT SRE F E2 05 C M AER O " CH w B , 1 . ES T • SWE205 SREFTEsT. CTEST + _ CHcH AN GE SREFTEST -CTEST + _ CHcA NA R D S CE205 SREFTEST • CTEST SWTEST SREFE205 " _E2 0 5 SRE F E2 0 5 " _E 20 5 SC T E S T SR EF E20 5 " _E 2 05 + _ C L C A N ARDE 20 5 XCTEST " XCE205 _REFTEST + _ C L F LAP XCp _REFTEsT + ( cgz20§ mom e nt '_

_EFTEST_EFE20S _TES'---W " _EFE20S C L_ERO " "_'_ST J

Figure 3-6 Equations for Building Up Aero-Only Coefficients for E205 Airplane Configuration " " ........... _ * "_ 1 o 'J _ ' i ' ;t ..... ; ! ...... ' !.... : i .......... ! i ' ............................................. I ' ..... '' ] ; i! ....

.... I ..... i_t.-. . ...... ! ....... i * _ ................ i ...................... i ....... __ .... _ ..... ,_ ................ I ........ _.-

i * [

L.l.._iilll .... 1 .... I..: :I .... I ....... i ....... i...i.i_ i .... i .......... i -i-l-- :-i ....... ! ___ i... _, ........... ii ..... ,.....

• " : i : : . _L ] .................... : ......

Figure 3-7 E205 Minimum Trimmed Drag Versus Mach Number

• • •

O

1.1 2 2 • UNP O W[R L D I • • STE • 0 0 W A E DC TC t,1 2 .!

t 2 / 12 t.I J CL / !

. I / .i | JI / / 4" 1ll ,0 I -Z l

// '/

- .2 1 O 4 | 12 1 S Z 0 Z 4 4 3 2 I 0 - ! - , 2 ar W l kG CM0 3 € • 6 T[* _o ( R U _ t 0!

• A[DCTC St? _ / 2 0 /

,, /

a I ! f D ,

'if

• 1 , 7 €D Fig ure3 - 8 Power-Off Wing Body Lift, Drag, and Pitching Moment C har a cteristics I f o r the VEO-Win_ Fighter Model of Reference 4 ,Mach = .2 S YM TEST - R UNIMACHI LEF TEF HO R I Z . _5 P r e di c t e d 0 .20 0 . 0 0 . 0 - I 0 . 0 : li P r e dic te d 0 .2 0 0 . 0 0.0 i O . O t: : ' _ ) P r ed i c t e d 0 . 2 0. 0. 0 0. 0 0 .0 t,. P redic t ed 0.20 Wing - Body ' I : 327 43 0. 2 0 0 . 0 0 . 0 0.0 " : i= _' ' 327 30 0 .2 0 0 . 0 0 . 0 O FF !!_ : : i : : . :1 :.: 1 : I: .... _ :-| , !

:[ !.

i@ --_.

I ! • : r . : . !i .[ ! : = _ -'1_ CRE F ...... i : t : t •1 I : : _ L -- L- - l : -i ........

_-#-- -+.-_ .... t .--t i i ! ! !

[._ ..i, i I.

I I I l f I I I ' I : 1: 1 t 2 : : : :i [ _ 1 ,t_ 1 ;;t!_ 1 2 t = H_I_I _ ti ! ........

_:H:H!H+H_i L ! t li!:Ii_ ...... :, ':_

-i ..... . . -' - .........

Figure 3-9a Lift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E 205 Configuration , Power'Off , Mach = .2

• • •

• • •

.J ARC - 12 -3 27 i E 2 05 TEST R U N _% C H LEF ' TE F HORIZ. i P red L c t ed 0.20 0.O 0. 0 -I0.0 _ [ 3 P red i ct e d 0. 2 0 0 . O 0. 0 0 .0 !

/k Predicted 0.2 0 0.0 0.0 1 0. 0 . i!

3 2 7 43 0 .2 0 0. 0 0. 0 P r e di c t ed 0 . 2 0 W i ng - Body o . o _ _ 3 27 30 0 .2 0 0 . 0 :1 _ , !

' _ ,' _ EF = 3 8 4.00 f t . 2 !! _ .

•' - :-r=: ...... ilii ...... i .

_l . , _ / ; . IL . j - 4 f,. ?i_- ,l , _: i

2 "

E205 ...

.... V ITFTI -

.... , ._ .I .. .. l.. ............ _ .......

i i"i ii I : : i ; [71!ii Ii' : _ ii F i gure 3 - 9b D rag Co mparis o n o f Pred ict e d and T est Longi t u di na l Aer o d y nami c . C hara ct er istic s o f Basel i n e E 205 Co nfigura t i o n , (Ex panded Drag Sc ale) , ~f-' '" \:- _. i SYH T~-;T- 'R~-;"T ~~~ , LEF .. _~~~ ··'-"~~~l·~''- -;'. -"- : .. _~ --'~".-. -I -. ~r·j I I : ..

i . : -- .. ~ .. 'ib ''--: ... t'!' E H:~H~:~ g:~g g:g ~gj -:g:g 'T' i ... ··r "" : ("':' ·"rt·~ ':"-'- - ... -:', _.+ .. ~

Lr - _ ---- - 0 327 33 0.20 0.0 10.0 0.0 \-- r--'" ._--! r-'-' '-:--1 ' -f--f'

~. r--' t. f' ~~~~ ~t-- ~+.... i- ..... ;.. ~- - ''''.~~' I ! .-_..... ~ ... :- '-t·: .- ~~, .. :.- \... 1'-:" ---\ .. h- . ~ ~-,- .~t-.-w-.·.

i : : ··1 . , ~'t- ! .. ! ;' .... .~. I" 'I.!:; ., I : 1--.' :-.'. .... - 1-- .. ....,.-. iI. .:.- 1--- . 1-- ... --.1--.... ;.:.; ",111'-.:-- I' - ....... - ... +- - .. it'. "-' f--o--"':' . ~ - _ .... f .. ·.. -- ': . ..- -·t-,l-,-.

[~~~ ... .!: 0: : . .'fl,.... I.: j I·; :.,.j. · .. ·ji:.

~+ .... ·f·61 ~...;- --.: . -~-~~l'7 'i ... -_ ... ~ --:-------,:i .. r~··+ .. ~- -'-,---',( : oW " : I ~T':1

S'~!~ ~-'}+' , ;' i/~"f~~'f"; I • ---r .~_. ·~c---~~J .. ~ .. ~-~T:l ~. ; i ·i· ' \ -l-j

r+.. -:- .. pt- ~~ :.'- -;-- ,.......~ f'-' .' _ .. -I---.. . .. -.. :-: 1--;-. :.: ... -. -.-. I-.-.-~ - ... '.-- "·~ ... -· .. I-· .... -. I·.. I~.-+-' t-+ !--~'-'''-r~' ... ; ....•. .." .. ':"j' . "j" 17 [z" :'>.."". ,"! '!'j ...

H- --"-. -- - .: I" .. .: . I '~,~~-+L: ..... ..±-'-

.. !;~.+.+.. 1-.:.) ~;j.z . .. ., .. ,.. . i· .. . . . . I .'\.~ I . t·.· .i..! ., -;. : 1·-:,· ...., ..

r- t-- .... t.l-- --~ _.. : : VI), )i r- -' --~... I '1'-" '-; ..... - ... _- --t---f---l-' .~.- .. '1)': ! "'~ .. :! : +- I-i I.. .. "'!: ... I: I·' ',;) . I ., .. '. . ! .: .... ,..: :.. L: 'i.' l' \~. i'" I .. , .. --,' I i,,· , i . -:-iAt'rfTf- --+- -i--i-f- ARC-12-327 .. -f-,- r--' 1-,-1--- -, \.J\'l-' . 1. i I-I . ;. I I' .....

T'\ !'''I/~JlCI ~ i' -1. i' I I 'rt"- ---~. ~- ... : .. ).- Lm:l\~'!:f.~ 'b j'1'~ .! .. T"'I 'j E ~:r"f' .:- ... j .. j .. :.'~AVr·· 1 .. .. ; .. : : .! .. j .. ; 'I ?o .. l i·:· .;::. l·i \ 1\ r~'" ~\I"·~ ... : ... + .. j.++:. ·I·'- .. !

I . : • I • M';' M : I . .:.. T 5 2 l,'t' ~" I

· .. i ; I· _ .. ··\·'··1 .. :··· ,:;" . :'" -.! ! ....... i REF = 384.00 ft •... , 1: . : .j : I": I ... : .. ,... : ....... i' 4· 1--;..·!t"D :.: ...J' :. -- ~. .. --j---i '-H--'--\--!--,- . --8: ~- .~-.p+-- ~·;Ofs ;, '. , .!:. : 1·:_· .. ·I·~7.~/I .. ·' .. ··i .. ·-i· .. ·i ... i.. I ..• ~ EF = 142.68 in.'; ... I, I: .. · ... ! 1"1.:\. m'~:" ",i.-I'."." .. r'· .. ··.. · '. ., I··· i ! . :.. i f:7,/J, :..!.. ·f ' · ~ ·'···1: .~ I :T: I i j.:; . ':'" --;'~i'-:"'I-: ... :. ·,Tl \.[ I··: ... ..: ., i'~"'i'- i--l- . -- 0,., . " 'J/-ff' :. -'.- 1 • : -- :. ..~~ .. ..::..,·ia.... ~, . . I ...... ·j_·-·,

. ; .:. (;';1 .. . I .1 .. : .j.. i··, .... i .:. . ... I .:. ;. ~;\. .: ... I. '. . . . j

; ; I~'" -.: lY~~ : I .. :. ; .~ .[=': .. '1 :. E205 __ = ~l-=~I ; ·:.-+-r1·1~. ·j'''':·~(\r5:. -f\~;":"-~~ "~'T"'l'i--:' .,.-tOo '--"]U.+' -#)--.r---. ., ;--I--i--I---. ~=-== f- - .. -I- ..... _. -_. -.,.....,....... \: : ... - '-'" . i . P/,.: :. '1:· .. ---- -,-=:'~-'::::-- ----~i :.: .... . :·1 :.; i\i~ 1"\: 't ~ I' +. 1 :: c-t ... j..... .... 1-- ... ~-l-. . . ... -- --'- --- ..1 j" '."- 1-.,..... '-"1-;·1--·1 .... -f--+-. "'t'ir'1-I1,' . ---t--'r-i- i- ii • ) ,/·1 \1:. i f': ~ ~ I··: ! i: " .! .. 1 / \i 'fi , ¥!., ·fl·· . i i: I-A· l--'~' J: . ·-t- -'- --r-I--- --- -- --- r·· _ / " .. L·~+- __ I: : i"'· ·.. m--t--" ·-,-1·- '. i .' . " . , .:: 1\: ' i·· , i :.1 I' 'I . , . ...; . I . . . I' .. , !. • I' lri··· "1 I'· i .

~-I"'" or;" ;...i....:. .. _ _ --- ----- "-' !'" ., ,'''" + .. _ .. -1--'-1·'· ,·_: ~"-l--l"": .j : L I .. : 'l---I'~'-r~ ';-! .. ·-1o .. -1·~~r-+· .. -!~ j

"{ I ! . ! I. ~ ! . . !"., '

t.l,.... + r· '/f '+: 20 i .... ~~ i 21l : .3 .' a.S q~ 0·3 D.:? ;r, i.~.o':' i -0.·J.. -0·3 E

\ .. -kz \. i IlLPdtl (b l6Il ES); .. _\ -0.2; ..: C 1'1

I I

i

I I I I 1. .. I !

Figure 3-l0a.Lift and Moment Comparison of Predicted and Test Longitudinal Aerouynamic Characteristics of Baseline E205 Configuration with Wing D

Trailing-Edge Flap Deflected +10 , Power-Off, Mach = .2

~:-----r--- - ,. "---r- -,.------ ----- ----. -_ . ..,.----,.--.-,-- - ---~- .. --~ -. - -1' - _. -- - - ~--r-..,----r-'----'-"-r--r--....---r--r-> ·1 .; 1I! S~ TEST RUN HACH LEF TEF HORIZ." " i" ;.. I . ' . i .; ri··· ... , .. -'" .... L. 0 :~:~~~~:~ ~:~~ ~:~ ~~:~ -1~:~ 1-- ..... --- ... _l~ .. ;_ •.. ,. " "'1---;'1-'-1-'--'" -- .. -:- ._- _._+ 1:1' ! t::. Predicted 0.20 0.0 10.0 10.0 I . :.:! !' ~' .•.. \.

, ~. --:--r': .'.1fff . i 3 27.. i 33, I 0.20 - 0.0 .. 10. ~ I 0.0 :- .... -- -::+-:- '-'~-J' ... :. I'" .. !. ~ Ii -~-:::~t4" ... ;.\.. .. '.. ;.- H+ C\.f"H -i-1-; .. - ... : .... - -r~- f- .. - - .. - _ ..... i- -,- ... ;-+,-+..... + ~ E"" e-,- --:-r-' '-. r:-r ~TIJ~' ' .. _; :. ':'. ' . ! :. !. .; . , : i: ' i' . : 1 I 1/ ,~. . I ,. . .,,> r) I . I" 'j.

H

H·· ... ~ ... , .. - ARC'-l2-327 .- -'T- -:-1-'''' -;'-f-.- ·-rr---;--t-·r--j .... >1:0.. '~~-r'--"- ... ,. ... +::::- /'F--' ._. ~ .. -.. , , '+--i, :- ....... ;.--+-,.-l-----I---;---+-~f---+-...,.....f

I:i t· : i' ,.;.... . .. , .. , 'j' ..... '.' ,. " £/'1' h .: : 7"1 '·1· ·f··

,.-,-." 0;7 1-"1' -j-, . I .. L " f-,-- - ,-r'-ri", .. , 1-,- --r- -.- t::::~ ...... ~?I-"-- -_.. - --1- , , !i . "IE 205,. Ii,. ; i ., " -.. '" '/ft i .; ':""" .i.

lit r I .... : , . " , ~~./'.....: ,. i:'

-,. .. -_ ... ----r-:-,-i- - .. .,-,--.; ..... ---.-... ~rl -/' ._. --.]_ .. _\ ... _-.. "; -I--- ., i I .1"., , .. ,.. . ; ;.e ".; . : ,Y "V ' : , :~ .,' I··:·' '1' I .. • .. -··!T : ....

f--'-. -10;" . S REF = 384.00 ft. 2 -' - -'" ... _ .. =-:.- •. -. ~;;:r' + ~ ....... - "--- -;- .. ;;.6t" _ .. -'- _.- . _. ++- -' .--J-..-

f4..r ',ii" "T'T , .. ".:," i"".-r r : :' / . i "i i"" ,-

~···':+-~"--r;r.' .••••.•••• ;, .. ;;'}.·'f=i~t~(:~t[i~, .;:,.·+"r~."+ ,rr'j: :.8:.:.+

"i 71" .,/ ,,-I.-V '[1 '1_ .+, 'W- .... ,J. ·1,-;·... ..

4"'::. Tt~t.r*:Lii.I··· / /(:=:~tr~T'"" ::;~~~]-cJ~~=!IJj+T ::::,'

H- f---.- i---t-7- .• --j--,. '; i ~~ - -- . -"'1' - . ': -- ~~IO'~ _ ... _.I~Lj.:-_=·.i. r1' ·i ~( ..... / :/·1··, .. ··• ' ~c=~~~'7 - .'- .. ·7~ ~ __ : __ ~_~ ~ . "'j"i' ...

.. ' .+ ).,.: .,. I:· ... i .... yi·t· 1 . ..... .. i.. .. iT' .. ; .. + ·····1;· .

~r·_··- '-r tJ-r T·r--· .1. I,.. ', ..... ! ...; .. ; i . , . ~ ~. i .- ;. i 'i: i ...

~:~ f+ +1 "_Iit T ~ .1. i: riIl ::':Ji __ =~:tj="]_,! " ......, ····i:-

1'·1' .. I ... , "'i . . : ... .. j ... 1""""\ .. i -'I ···i· . i ,...: l : ·1"·~·· , . I· .... 1·· .. +··+·· ... , ..; :. ,tt r.- .....

i ':'.' : . .:.. .. J"'H-- . __ .~'~3i .. ,.. .: I·· ~. O!_+ _ _ ~_. __ ~~ ~~,- '~L .. ~_ . .. ~:~o~ _: __ .. _. --:O¥'!~_f-' _. Ii .: ... ... _O~ I.~ __ ) _ .. "_ '!:O ,....·-!O·II{:J .... ; n. . L...! i i· .. .; .;. i'" ' . 1 : :. • • .. i I : C 1>' '" j . : ... j: .. : ...

4i- " I';~ iI' : i i : ' , :' -!-. -+ : -,-f--'-'- -- --'1 ~- -,- .... -1"--

.. "'1 ... : - _.\:...,. - !. I· : . : : ; .... ,: : I· i· : ; I "I'.'. .,.. . ...i..: i·

~ +-~. : : • ~_...J.._ r-i- -L -+- ,.......- -~-I- .. '- ~ I -i-· -.- -+- _ . .:... ..... -', "1 -- .. _. -' -- - _ .. ' ; ... --+-+-+-i-I-i--i--'- ;. --'--t-;-.'; +-',.-1_ .... -1 ...

\h ~ .. , ":1-" . ··1· . ,:.+ .. :)~~._ .... ) ... "'1 . -\ : .. :.! . ; 1,-, "; ... i. . I ; : .! ' . ! "')' .j.. ;' .. :... .~.. 'i-'~' -. :., " .. r···· ';;io'-'l<t.~ -"-f--r" ~-+.+- -t-.; t-;- _ -.,_. . '" ... t- .. " -i'- 'r - - ·f··-- .. ·· _·1 j.... ."""-j- 1-,- "T -. r'- '--r I .

~'~FF ~:: 2r-- Jt: ::.L-:j .... :.~ ~ J .. : -;- . : ... ... i. ~.,_~L . ." 1-·· -- ·-~_c-J·l .. ~- ; :. . J' . .... . ) r-L ._.L. ~L_i..- -- ... - r--';-i --L- ... :. rt i ~l."; 'oj'" ;: .. ! _ ;.::,~. : . j" 1. _.~ __ L e-L.~ . T' , i i _.L.1 .. ~_ ~ _.:.. " .... 1...._. :_~ 'J.:.. .L ill r r j' : .( Figure3-10bDrag Comparison of Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge

Flap Deflected +10°, Power- Off, Mach = .2

F- - HORIZ • ' i I i SYM TEST RUN MACH LEF TEF I • X P redicted 0 . 20 0 . 0 2 5.0 -10 .0 ..... J - : - "_ ......

[] Predicted 0.20 0.0 25.0 0.0 I i /k Predi c t e d 0 . 20_- _0 . 0 25 . 0 1 0 . 0 <> 32" / 4 2 0 . 202 ..... [_0 . 0 25 . 0 0 . O -4- " _"1- ' : / -! " i .....

.....i.....ARC-12-327 : I !

"i i E 205 ! _ i _'1 -- SREF = 384.00 ft.2 i.

I I : 1 ; I : I I m i _REF = 142.68 in .....: :-t E205

i' i. . .i_ _.1 I _

--- = --_- :-- - - 'j- '" i ..... i

_,I.- _zL . ___, ,. - -J '.... "

Figure3-11aLift and Moment Comparison of Predicted and Test Longitudinal Aerodynamic Characteristics of Baseline E205 Configuration with Wing Trailing-Edge Flap Deflected +25 ° , Power-Off , Mach = .2

• • •

• • •

S_ TE ST RUN MACH LE F TE F HO R IZ. ! : i ' i X P r e dlct e d 0 .2 0 0 . 0 2 5. 0 - i0 . 0 I ...... ; , !, .....

_ P redict e d 0.20 0.0 25.0 0.0 " i ' Pr e dicted 0. 2 0 0. 0 25.0 I0.0 ...... ! ....... i.: '-l - -i. I !

-! _ 32 7 4 2 0 . 20 0 . 0 25 . 0 0 . 0 i -[ i I !

ARC-12 - _ lI-"- -- -[ ' E 2 0 : _ i z _ ..

I .... , SRE F = 3 84. 00 ft. 2 I _ .

.... _ --- ! -- i-! _- - !, • .i: _i I_!:' - _-i-= ........ - t_._ =--_ ....

.: .[| - . . .i...... 1 .......:. : . . . . . _ .. . ; .

- - + - . - ;---- i ........... i- - - ...... i i

'- ' _ . _ -- _ - -!--

! 1

o ,V_ o . , o _ , /o.: , _ - _ I _ o ._+ _l if!!..... i: -- _

....... -...... :_- i........ : '_ ............... : :-i..... ---] .......... ! .-i ................... _...... _........... _ ........ _ - -, ---- _ ....... _ ........ _-- - -_- _l ................ ' _ i :l . : .... i i ...... ' ! .................

t • ] • ! : : . . . , _ , . .. , . ' . . _ , ....... , Figure3-11b Drag Comparison of Predicted and Test Longitudinal Aerodynamlc Characteristics of Baseline E205 Configuration with Wing Trailing-Edge (j'I N i t.O -0.4 -O.Z 0.1, 0.8 0 ~ 10 I~ 2'0 -1.0 -08 -o{, ·az 0.4 •.•• 1 ., I .\ I '!

T" (X de, I !

I i

. L. ! :J. :\

Figure 3-12 Full-Scale E205 Airplane Power-on, Trimmed Lift Curve- and Drag Polar-Envelopes from Wind Tunnel Data for M = .2, CTTOTAL = 1.81

• • •

Figure 3-13 . E2 0 5 Trimmed Cruise / Maneuver Drag Polars ( _ TE 10 o 50 CLT CT = X .

TESTDATA 0 ( _ TE 0 C2LE CD S / _ SM

® ® ®

L CM o I CD E CLE (+) (-) CDo CMT M = XX CT = XX OPTIMUM e ENVELOPES MACH ( _ C= 0° OPTIMUM ENVELOPE CT = X eE _ TRIM eTRIM CT= X o_ CLE _ " ---

_ . _,,. _., _ f _4 - - _ _ : > _ , _

_ . . 4 • / _ _ " _ _ TE _"_ . _ _ TE , _ _ .e.6 _ . _ TE ., , J_ \E . _ @ - " " i.0 _ TETRIM CLE MACH Lo , 4Lr d Lr LT= Lo+ A Lr+ A LR e=T T '-'A- _ " C2LE CD E - CDOE cg WHERE ,L o = POWER OFF AERODYNAMIC LIFT 4 Lr = LIFT INCREMENT DUETOSUPERCI R CULATION T L R = REACTION LIFT DUETOTHRUST X = SUPERCIRCULATION A R M MT= MoX d L_ XR A LR Xr XR = REACTION LIFT A R M CDc = CAMBER DRAG CDs = FLAPSCRUB DRAG Figure 3 - 14 VEO-Wing Trim Method for Maneuver

Q • •

• : • ...................................... t..... ' I " " ' I ................ ' ' ' " '_ * ...........

............................................ f , 1 1, , ":"l"'": ": : :""_ [ :. ', ' . : ' , ::: i !i i: :!:: : : - : .: . ; _ : : :_,[ 1: :::,,: :i' : ', _ Jl _ I , . I . ,, , . _ , : . _ , ,, . , , I , ! ......... , I I ....

• I ....... ,................ . _ _ , [ _

I1 I I I I l ,,, _ ,, , ! I !

! ............. I

Ii!!_ 1111111111 , , ,, I lt,!h , ,, , .,.* * ,_

" " ' ' ' " ' ' ' • " ' ' " ' ' ; ' " : " : " " " " ' _ " ' t ' ' __I_,'.;_FI_..._-_h-i_ I.-'-_, _,. b-'bT ! - '1 -_ , _-_.'M ac h = .6 , Cp = . 30 2 MAX A / B PO WER _:: IL_li:_J_l :_[[ILI_]_. g_Ll__.'_'_ ' :._-'_-. K a ch = • 9 , C / _ - . 159 MAX A / B P ( _E R _' I.":.!l_:.l_dMti_ilii _E Y Hqc:I :. I: illii "-I.Li I_-'1"...1.:.'* , , ....,..., ,, _ , :l .... , , , ,,i, , , , , ,.. ,, ,,, , . ,:L:'.qliiil!i_i [ fill il_: _1 : _1: _--:-:I_---I_I_ L I'LL_IL:I _ - -q_-.q-Ht:- - .+l--_l-"--.:: -:1__ -_ M at h = . 9 PO WER O FF_,H_t. _ • :. .: ................... II .... , ......

-_- ..... _- :-: ..... c _ .... !..-'Math = .6 , C = o171'"::' : ':;:" _ : -' _ _. :,,:.. =_.. ::,: :,: : 7177I_._.M _, .... , ......... ' ......... " " '" ' I_,I ilF_IIU_ _ iFl_li_lT;_l _ l_:Tli : lTi_ : liilTil_ 7 7F_l_71_F.i_oo_-., , ,- . o , o _:,:::,,,,:,::,,,_ _7--_..: - i ...... "' _ " _ 1 _ _ .................................

.:::: . :: :: . i:_;_::: :: ...:. : ._ : : : : - !i!!i :: i: ii : !.!! !!! i!l:! 1 ii!:it ! i!:!!_ii! !{!!

!:: i '. i : i_ :i:!!i!!:!:t ii:i!: _ . ": : - :_:i:: '. :i!: . i! _ !i . iiii!iil :i !:!il .i!'ii!i!!! i! V . i':!i_:!i! iii!liiiii!. . :i:i

:i_ _., .... _ :_"-_: __._ _ : - :_: : : _:_ i :_: _',: '_: __i__:-_ ::_ :iii: _

I iI3 1' !tl 1 , i ....... ....... '! :'_':_::: . . _ ": . , __ ii . - : , • !!i _ i !!i

:.... I _ . . . ! . ,b ..l:.:. : ..... :.LL_ !__ _-_._ ....

i :: :I: . !: . L-L .....

re 3-15 Power-on and Power-off Predicted Trimmed e s as a Function 0f"Ecuivalent" _.i::l: ....

Lift Co e ffici e nt , C IE , Mach Numb e r , and CZ ( from R e f e r e nc e 1) Figure 3-16a E205 Lift and Pitching Moment Curves at M=I.2 with Canard Deflection

, • • •

.. +. -,-+-. ; c~(.:~'~~vj ;Y:o'j ~ / / +L_-[- ~..I

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• • •

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Figure 3-27 Photo of R104 Wind Tunnel Model

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Figure 3-28a Variations in CANARD Location for E205 and RI04 Wind Tunnel Models O Figure 3 - 28b Variations in Strake Shape for E205 Wind Tunnel Model Q

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