Analysis of the wind tunnel test of a tilt rotor power force model

C R 137529 A V A I L A B L E T O T H E P U B L I C

FINAL REPORT

ANALYSIS OF

THE WIND TUNNEL TEST

OF A TILT ROTOR

POWER FORCE MODEL

(NRSA-CF-137529) AIBLvSIS Of THE WINE N76-18167 TUNSEL TZST O F 9 TILT R O T O R POW59 POZCE EOCEL Final Report (Fell Helicopter C o . ) 137 p B C 86.00 CSCL 01C Unclas

REPORT 301-099-004

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N A S A C O N T R A C T MAS 2-8084

BELL HELICOPTER COMPANY .,)* , . r e 6 . ..a .-.- * - 7 1 '*..S -. ^. . & we.*.

CR 137529 AVAILABLE TO THE PUBLIC ANALYSIS OF THE WIND TUNNEL TEST OF A TILT ROTOR W E R E D FORCE MODEL R. L . Marr D. G. Ford S . W. Ferguson Bell Eelicopter Company Report No. 301-099-004 June 1, 1974 Prepared Under Contract No. NAS2-808b by Bell Helicopter Company, A Textron Company Fort Worth, Texas for National Aeronautics and Space Administration Ames Research Center This data is furnished in accordance with the provisions of Contract NAS2-8084 Foreword This report is prepared by the Bell Helicopter Company, Fort Worth, Texas, for the National Aeronautics and Space Administra- tion, h e s Research Center, Moffett Field, California, under Contract NAS2-8084.

The Administrative Contacting Officer was Mr. Dennis Brown. The Technical Monitor was Mr. Kip Edenborough, Tilt Rotor Research Aircraft Project Office.

BEU use or dtscbrure a t dilr on *IS ~ * ) c IS subgcl b the rntr*tan om thc btlc php -ooumw* TABLE OF CONTENTS page LIST OF ILLUSTRATIONS LIST OF TABLES LIST OF SYMBOLS I. SUMMARY 11. INTRODUCTION A . PREVIOUS TESTING B. TECHNICAL BACKGROUND C. TUNNEL DESCRIPTION D. OBJECTIVES OF THIS ANALYSIS 111. DESCRIPTION OF THE MODEL A. CONSTRUCTION AND DESIGN PARAMETERS IV. DESCRIPTION OF TESTS IV- 1 V. DATA REDUCTION 'I. RESULTS OF TEST

A . ROLL STABILITY - IN GROUND EFFECT (IGE)

B . WING DOWNLOAD C. STATIC STABILITY CHARACTERISTICS VII. ANALYSIS OF RESULTS A . HOVER DOWNLOAD B. ROTOR WAKE ON HORIZONTAL STABILIZER C. ROTOR WAKE ON VERTICAL STABILIZER VIII. COKCLUSIONS

A. ROLL STABILITY - IGE

B. WING DOWNLOAD C. ROTOR KAKE ON THE DIPENNAGE IX. LIST OF REFERENCES APPENDIX A 1 T C T n S ' Tl.1.ITSTRATTfWS Powered Model an Rotary Sting bkt, Langley V/STOL Turnel IV- 2 Helicopter Configuration, Nacelle Inci- dence 900, OGE, Walls Up Conversion Configuration, Nacelle Inci- dence 600, OGE, kalls Down - Conversion Configuration, Nacelle Inci- dence 30°, OGE, Walls Down Airplane Configuration, Nacelle Inci- IV- 3 dence 00, OGE, Walls Down Force and Mament Sign Convention v-7 VI- 7 Rollinq Manent in Hover VI-8 Roll Stability Characteristics in Hover Roll Stability Sununary in Hover Wing Download in Hover Wing Download Variation With Flap Deflec- tion in Hover, OGE Wing Download in Forward Flight, OGE Lift Coefficient Versus Fusela e Angle of Attack, Nacelle Incidence 980, Airspeed 40 Knots VI- 14 Lift Coefficient Versus Fusela e Angle of Attack, Nacelle Incidence 980, Airspeed 60 Knots Lift Coefficient Versus Fuselag8 Angle of Attack, Nacelle Incidence 90 , Airspeed 80 Knots Lift Coefficient Versus Fuselage Angle of Attack, Nacelle Incidence 90°, Airspeed 100 Knots LIST OF IUUSTRATIONS ( C o n t i d )

a

VI- 17 Lift Coefficient Versus Fusela e Angle of Attack, Nacelle incidence 980, Airspeed 120 Knots Lift Coefficient Versus m e l a e Angle of Attack, Nacelle Incidence 790, Airspeed 40 Knots Lift Coefficient Versus Fwel e Angle of Attack, Nacelle Incidence 8 0 , Airspeed 80 Knots Lift Coefficient Versus Fusel e Angle of Attack, Nacelle Irrideoce %o, Airspeed 120 Knots Lift Coefficient Versus Fuselage Angle of Attack, Nacelle Incidence 60°, Airspeed 120 Knots Lift Coefficient V e r a Fusel e Angle of Attack, Nacelle Inci nce 80,

Airspeed 240 Knots ?

Lift Coefficient Versus Fusel e Angle of Attack, Nacelle Incidence 3 0 , Airspeed 120 Knots Lirc Coefficient Versus Fuselage Angle of Attack, Nacelle Incidence 30°, Airspeed 160 Knots Lift Coefficient Versus Fusela e Angle of Attack, Nacelle Incidence 0 , Airspeed 120 Knots Lift Coefficient Versus Fusela e Angle of Attack, Nacelle Incidence 0 , Airspeed 160 Knots Lift Coefficient Versus Fusel e Angle

4 3

of Attack, Nacelle Incidence 0 , Airspeed 160 Knots iii LIST OF ILLUSTRAl'lIINS (Continued) Number Pitching Montent Versus Fusela e Angle v1-22 of Attack, Nacelle Incidence 800, Airspeed 40 Knots Pitching Mauient Versus Fusela e Angle

f

of Attack, Nacelle Incidence 0 ° , Airspeed 60 Knots Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 90°, Airspeed 80 Knots Pitching Moment Versus Fuselage Angle of Attack, Nace 1 le Incidence go0, Airspeed 100 Knots Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 90°, Airspeed 120 Knots Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 7S0, Airspeed 40 Knots Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 7S0, Airspeed 80 Knots Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 75O, Airspeed 120 Knots Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 60°, Airspeed 120 Knots Pitching Moment Versus Fuselage Aagle of Attack, Nacelle Incidence 60°, Airspeed 140 Knots Pitching Moment Versus Fuselage Angle of A t tack, JSacelle Incidence 30°, Airspeed 120 Knots LIST OF ILLUSTRATIONS (Cantimaed) Page VI-39 Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 30°, Airspeed f 60 Knots Pitching Koment Versus Fusel e Angle

2 0

of Attack. Nacelle Incidence , Airspeed 120 Knots Pitching Moment Versus Fusel e Angle of Attack, Nacelle ~midence%, Airspeed 160 Knots Pitching Moment Versus Fusela e Angle of Attack, Nacelle Incidence 80, Airspeed 160 Knots Yawing Moment Coefficient Vgrsus Yaw Angle, Nacelle Incidence 90 , Airspeed 40 Knots Yawing Marnent Coefficient Versus Yaw Angle, Nacelle Incidence 90°, Airspeed 60 Knots Yawing Moment Coefficient Vgrsus Yaw Angle, Nacelle Incidence 90 , Airspeed 80 Knots Yawing Pfoment Coefficient V e r s u s Yaw Angle, Nacelle Incidence 7S0, Airspeed 80 Knots Yawing Moment Coefficient V e r s u s Yaw Angle, Nacelle Incidence 60°, Airspeed 120 Knots Yawing Moment Coefficient Versus Yaw Angle, Nacelle Incidence 30°, Airspeed 120 Knots Yawing Moment Coefficient Versus Yaw Angle, Eacelle Incidence o O , Airspeed 120 Knots LIST OF ILLUSTRATIONS (~ont inued) Yawing Moment Coefficient Versus Yaw VI- 50 Angle , Nacelle Incidence 0 ° , Airspeed 160 Knots Yawing Moment Coefficient Versus Yaw VI-51 Angle, Nacelle Incidence oO, Airspeed 180 Knots Rotor Power/Collective Pitch Variation VI-52 With Airspeed, Nacelle Incidence 9 0 , QGE Rotor Thrus t/Collective Pitch Variation VI-53 With Airspeed, Nacelle Incidence 9 0 ° , W E Rotor Power/Thrust V~riation With Airspeed, VI-54

N&celle Incidence 90 , OGE

Rotor Power/Collective Pitch Variation VI-55 With Height Above the Ground, Nacelle Incidence 9 0 ' Rotor Thrust/Collective Pitch Variation VI-56 With Height Above the Ground, Nacelle Incidence 90° Rotor Power/Collective Pitch Variation Vf -57 With Airspeed, Nacelle Incidence 750, OGE Rotor Thrust/Collective Pitch Variation VI-58 With Airspeed, Nacelle Incidence 75O, OGE Rotor Power/Collective Pitch Variation VI-59 With Airspeed, Nacelle Incidence 600, OGE Rotor Thrust/Collective Pitch Variation VI- 60 With Airs?eed, Nacelle Incidence 60°, OGE LIST OF ILLUSTRATIONS (continued) Page Number

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VI-5s Rotor Power/Collective Pitch Variation VI-61 With Airspeed, Nacelle Incidence 300, OGE Rotor Thrust/Collective Pitch Variation V I - 6 2 With Airspeed, Nacelle Incidence 300, OGE Rotor Power Variation With Airspeed and VI-63 Nacelle Incidence Rotor Thrust Variation With Airspeed and VI-64 Nacelle Incidence VI-65 Rotor Flapping Variation With Airspeed and Nacelle Incidence Wing Download Comparison in Hover VII-10 Wing Download Comparison in Hover for VII-11 Flap Deflection, OGE Wing Download Comparison in Hover for VII-12 Blade Twist, OGE Wing Download Camparison in Hover for VII-13 Blocked Area, OGE VII-14 Variation of Wing Drag Coefficient at a , = -90° with Reynolds.Fhnber Horizontal Stabilizer Aerodynamic VII-15

characteristic^ , Nacelle Incidence 90°

Horizontal Stabilizer Aerodynamic VII-16 Characteristics, Nacelle Incidence 75O Horizontal Stabilizer Aerodynamic VII-17 Characteristics, Nacelle Incidence 6 0 ' Horizontal Stabilizer Aerodynamic VII-18 Characteristics, Nacelle Incidence 3 0 ' kt w acxksurc r d bti on %I$ ~ l p r tr m subpct (o the resbtctlrwt on the tttk gqc WPLUOOCTeR-

d

LIST OF ILLUSTRATIONS (Continued) Page VII-10 Horizotltal Stabilizer tlerodynamic VII-19 Characteristics, Nacelle Incidence O0 VII-11 Horizontal Stabilizer Aerodynamic VII-20 Characteristics, Nacelle Incidence 0 ° , Flaps Up VII-12 Induced Velocity Ratio in Plane of the VII-21 Horizontal Stabilizer VII-13 Effect of Yaw Argle on Rotor Wake Upwash VII-22 at Horizontal Stabilizer, Nacelle Incidence 300 VII-14 Effect of Rotor Wake on Directional VII-23 Stability, Nacelle Incidence 750 VII-15 Effect of Fin Configuration on Yawing VII-24 Moment Coefficient viii 1 h or diiarurr d (LU on this is

I wekt Lt @ mtri&m o n Iht We Psr. I

LIST OF TABLES Page FULL SCALEBODEL SCALE PARAMETERS 111-4 MODEL INSTRUMENTATION 111-6

DATA REDUCTION v- 2

CONFIGURATION CODE

v- 5

CENTER OF GRAVITY POSITIONS V-6 HOVER DOWNLOAD PARAMETERS HOVER DOWNLOAD aeeA Use ar Qrclerun ot Wd on mcs plgv 15 -ewrrrvr subpct t o the m t r i r t m w, the title pale LIST OF SYMBOLS Symbol Description Units

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A Rotor disc area Axial force N (lbf) Horizontal stabilizer lift cu~-ve slope f /deg Fore and aft flapping angle deg Fore and aft c clic angle,

reference to t K e shaft

Buttline location B.L.

Wing span m (ft) Lateral flapping angle Aircraft axial force coefficient, AF/sSw Aircraft drag coefficient, Drag/qSw Profile drag coefficient of the wing Aircraft lift coefficient, L / ~ S , Slipstream lift coefficient, L/qsSw Aircraft stability axis rolling moment coefficient about the aircraft center of gravity, RM/qSwbw Aircraft stability axis pitching moment coefficient about the aircraft center

of gravity , E4/qSwr;w

Aircraft pitching moment coefficient, due to horizontal stabilizer in rotor wake LIST OF SYBtBOLS (Continued)

Description - Units

Aircraft pitching ttrolrrent coefficient , due

to horizontal s a b i l i z e r ircidence iz rotor vake S l i ~ s t r e m pitching noment coefficient about the m n g quarter chord A i r c tft pitching moment coef f S cient , due to elevator i n rotor wake Aircraft normal force coefficient, N F ! q S , Aircraft s t a b i l i t y axis yawing moraent coe2ficirnt about the a i r c r a f t center

of gravity , YM/qSwbw

Rotor powr coefficient, Q/PAV= R Aircraft body axis pitching moment

coefficient , ~ q S w c w

Rotor pitching aroment co2fficient Aircrizft body axis rolling moment coefficient, W ~ S W b W Rotcr rolling moment coef f icieut Slipskrecrm rolling moment coefficient

Rotor thrust toe ficient - helicopter

notation. T/PAVT 5

Aircraft body axis side force coefficient, SF/qS, Slipstream side force coefficient

- 13

lire a i d t s c h r r e d fA ce m t h t r -owrrm

8 iubprt Bt 1 W f6lfatt;itt an rn ti%@

LIST OF SYMBOLS (Continued) Description

0 d q s

Aircraft stability axis side force coefficient, SF/qSw Aircraft body axis yawing nroment

coefficient , YM/qSwbw

Slipstream yawing moment coefficient Flap chord King chord Rotor diameter Fuselage station gravity Height of rotor disc above the ground Height to diameter ratio WD In-ground-effect XGE Horizontal stabilizer incidence, i~ positive leading edge up d e t 3 Aacelle incidence (iN = 0 airplane mode) deg i~ Rotor sidewash factor K~ L Lift N (lbf) Horizontal tail a m I n (ft> &H Length of mast m (ft> 1, Mean aerodynamic chord I n (ft) MAC Model scale factor MSF NF Normal force N (lbf) 301-099-00!t xii LIST OF SYWXlLS (Contimed) Descri~txon Out-of-ground effect OGE N-ta (ft-lbf) Aircraft pitching rr#latent P# N-m ( f t-lbf) Rotor pitching moment

Pn8

Rotor torque (per rokor) Q ~ / r n ~ (lbf/ft2) Free stream dynamic pressure slipstre& dynamic pressure, ~ / m * (1bf/ftZ) 9s q + T/A Radius of rotor m (ft) Aircraft rolling rrmment N-m (in-lbf) Rotor rolling nmment N¶ (ft-lbf) measured on rotor balance Reynolds rnnnber 3otor speed rev/ciin Side force M (lbf) Area of horizontal stabilizer m2 ( f t 2 ) Area of wing m2 ( f t 2 ) Area of wing under rotor (blocked area for zero flaps) m2 ( f t 2 ) Rotor thrust (per rotor) N (lbf) Aircraft thrust coefficient, T/qSw Aircraft slipstream thrust coefficient, T/qpS, Full scale aircraft ,airspeed kt Horizontal t a i l volume coefficient Model scale airspeed m/sec (ft/sec) 301-099-004 x i i i LIST OF SYMBOLS (continued) Description Units

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V Velocity of the rotor wake at %I the horf zontal stabil izcr d s e c (ft/sec) m/sec (f t l s e c )

v~ ~ i p speed

W . L . Uaterline location W. Induced velocity at rotor disc d s e c (ft/sec) =R Induced velocity from the rotor * ' R , H in the plane of the horizontal stabilizer d s e c ( f t/sec) Lateral stick positicn m (in)

x~~

W Yawing moment N-m (ft-lbf) Fuselage angle of attack Angle of attack at horizontal stabilizer Wing angle of attack Flaperon delfection angle, positive down Elevator deflection, positive down Flap deflection angle, positive down Pitch flap coupling Rotor wake deflection at horizontal stsbilizer, positive down € Wing plus rotor wake deflection at T horrzontal stabilize?, positive down deg t Wing wake deflection at horizontal W/H stabilizer, positive down 'I L!ynamic pressure ratio of the *R horizontal stabilizer, due to rotor interference Dynamic pressure ratio at the horizontal stabilizer, due to wing-body-rotor interference xiv mELL 'se w crtsciM'~te at daU m &I% 1.r subfect til the r*StrI€tm Mf mc t i t # plgr

--

LIST OF SYHBOLS (Contirmed) Symbol Description Units

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Dynamic pressure ratio at the horizontal

stabif izer , due to wing-body interf erence

Dynrraic pressure ratio at the vertical stabilizer, due to wing-body-rotor interference Dynamic pressure ratio at the vertical stabilizer, due to wing-body interf erence 8 - Rotor blade collective pitch setting at tip dei3 TIP Advance ratio 2 2 Kinematic viscosity m /sec (ft /sec)

Density w m 3 (slugs/£t2)

Elevator effectiveness parameter Roll mgle, positive right wing down d e t 3 Fuselage yaw angle, positive nose right d e i 3 Rotor speed rev/sec Sidewash factor Two series of tests were made to determine performance, stability and control, and rotor wake interaction on the airframe, using a one-tenth scale powered force model of a tilt rotor aircraft Testing covered hover (IGEIOGE), helicopter, con- (Figure 1 ) .

version, and airplane flight configurations, These tests took place in the NASA-Langley V/STOL wind tunnel, The first test was in September 1972 and was terminated because of insufficient collective pitch actuator capability of the arodel to c q l e t e the high power and high speed airplane flight conditions, M i - fications were made to the cotlective pitch actuator and testing uas completed during the second tunnel entry in October 1973.

Wind tunnel testing was performed under W A Contract NAS1-11582, Forces and moments were recorded for the model from predetermined trim attitudes. Control positions were ad'usted to trim flight

(one-g lift, pitching moment and drag zero 3 within the uncorrected

test data balance accuracy, Pitch and yaw sweeps were made about the trim attitudes with the controls held at the triramed settings to determine the static stability characteristics, Tail on, tail off, rotors on, and rotors off configurations were tested to determine the rotor wake effects on the empennage.

Data obtained during this testing will be presented in a Information presented in this report will cover the NASA T M . ~ analysis of the test data. This analysis covers only informa- tion useful for tilt rotor aircraft (rotor wake effects) and does not include any analysis of the aerodynamic characteristics for the tilt rotor model tested. Results from this test will b

used to supplement information obtained from other model tests. 1

Documentation of this analysis is covered under NASA Contract NAS2-8084.

The principal results from this analysis are as follows: Rotor Interference on the Wing (1) Hover tests in-ground-effect showed wing download to be in agreement with previous model tests. Correla- tion was established when wing download is presented in terms of blade twist, thrust, blockage, and flap deflection. Reynolds number corrections are required to determine the full-scale aircraft wing download.

Wing download does not appear to decrease si nificantly for full span flap deflections greater than 4 0 degrees.

Flaperon deflection is more effective than flap deflec- tion in reducing download.

Rotor interference on the wing was found to be negli- gible above 40 knots for all conversion angles tested.

Rotor Wake on the Empennage (2) The interaction of the rotor wake with the horizontal stabilizer is to produce an upload during low speed helicopter flight and changes to dowrrwash during air- plane f l i ht. The d m s h measured for rotors on in airplane flight is that due to the wing wake and is - .

the same as measured for rotors off.

Rotor-induced velocity at the empennage was founa to correlate with previous model test data. The rotor wake effect was also shown ta become insignificant as the nacelles are tilted forward and as speed increases above 120 knots, (3) Rotor Flapping Lateral flapping in helicopter mode was higher than estimated using low disc loading and low twist rotor induced velocity correction factor. Flapping angles were in agreement with previous model test results and estimates using the modified induced velocity correction factor for high disc loading and high twist rotor.

Roll Stability (IGE) The roll instability during hover ICE measured during this testing was found to be in agreement with pre- vious test results for tilt rotor aircraft. The roll instability can be controlled with a ,small lateral stick input with the control power available for the models tested.

Use ~r otsclost.re ~t ddld 2 1 7 this p q e .\ eECL r u b p c t to the r e s l r ~ c t ~ o n c n the t l l l r C ( v R - Y 11. INTRODUCTION This report presents the analysis ol: a wind tunnel test of a one-tenth scale powered force model of a tilt rotor aircraft.

Testing was accomplished to determine the performance, stability and control, and rotor wake interaction on the airframe ring hover. helicopter, conversion, and airplane flight. rhe Bell Model C100-FIB is a powered aer dynamic scale model 01 the Bell

Model D270 tiit rotor aircraft .? A three-component rotor balance

was installed in each nacelle to allow separation of rotor forces and moments from those measured from the tunnel balance for the overall model. This capability allowed analysis for rotor air- frame interaction to be made in addition to that obtained fron previous model tests. Rotor wake effects obtained from this analysis are applicable to tilt rotor aircraft configurations.

The principal objectives of this report are to analyze the test data and campare it with other model test data and analytical methods. Results are summarized for application to the design of existing and future tilt rotor aircraft. The documentation was accomplished under XASA Contract NAS2-8084.

A. Previous Testing Most of the tilt rotor model testing by Bell Helicopter Company has been with a one-fifth scale model of the Bell Model 3 0 0 as reported in Reference 2. These tests were directed at determining the performance, stability and control, and aeroelsstic characteristics for the XV-15 Tilt Rotor Research Aircraft. Rotor wake interaction obtained with that model was lhited in scope due to tunnel capability (inadequate for high spsed airplane flight) and model capa- bility (no rotor balance). Therefore, results obtained from the test presented in this report will be used to supplement the information obtained from the fifth-scale model tests.

Technical Backpround Tilt rotor model testing was initiated to address problem areas encountered during the XV-3 flight test program.

These problem areas were primarily noted during hover and low speed helicopter flight and were related to the rotor wake inter- action on the airframe. A detailed discussion as to the approach to correct these problems is presented in Reference 2. This test investigated these problem areas.

6 E u Ure or d ~ r c ~ u r e ol data on th~, pa)c IS rubpct to the *eslr~cl~on on the ttlle ~€UCOPlERcommrvr C. Tunnel Description The wind tunnel test was accomplished in the NASA-LangLey V/STOL wind tunnel. The V/STOL tunnel has a 4.88-meter (16-foot) test section and can operate through a speed range of 6.1 to 76.2 meters per second (20 to 250 feet per second), with the walls down. With the tunnel walls raised, the test section is opened to eliminate wall effects and can be oper- ated through a speed range of 0 to 18.3 meters (60 feet per The model was mounted on a rotary sting support second).

with a six-component internal balance. Adjustments could be made for pitch, roll, and yaw at the desired h/D (height above ground to rotor diameter) to obtain ground effect information.

This capability allowed the model to be tested at all config- urations from helicopter to airplane flight.

D. Obiectives of This Analvsis This analysis h-s the following specific objectives: 1. Determine the rotor wake effect on the wing. This includes evaluating the effects of various design para- meters on wing download duricg h w e r and changes in wing download during forward flight .

2. Determine the rotor wake effects on the empennage in sufficient detail to supplement existing data ior use in the mathematical model for flight simulation of tilt

rotor aircraft . Rotor wake characteristics for two

types of empennage configurations will be shown.

3 . Compare measured rotor performance with estimated.

4. Ebaluate roll static stabiiity during hover in-ground- e i ' f ect.

5. Evaluate rotQr flapping characteristics.

These objectives were accomplished and the results are dis- cussed in the following sections.

6EU UIP or dtsclo~ure of data cn thts plqe tr HEUCOFlER-I 5ubpcl to the rerlr~ct~on on the t ~ l l e pap 111. DESCRIPTION OF THE MODEL The model tested, designated the Bell Model C100-FIB, is a oc2- tenth length scaled aerodynamically sinilar model of the Bell Model D-270 tilt rotor aircraft. The model was developed under Phase I of Air Force Contract F33615-69-C-1578. The model has two five-foot diameter rotors that are mechanically intercon- nected. They are driven by two fuselage-mounted TASK variable frequency motors having a continuous operating rating of 25,353 The span of the model between rotcr watts (34 horsepower) each.

centers is 1.95 meters (6.4 feet); with the rotors turning,the span is 3.47 meters (11.4 feet). Overall model length is 2.47 meters (8. f feet). The model was supported during the test on a fuselage-mounted six component balance rotary sting support.

A list of scale factors, full-scale and model-scale parameters for the rotor, wing, fuselage, and empennage are given in Table 111-1. These are the parameters used during data reduction and analysis.

A - Construction and Desien Parameters A description of the model components are as follows: 1. Fuselage The basic fuselage backbone, from the wing aft to the empennage. was a square steel tube with aluminum plate bulkheads to support removable fiberglass fairing shells.

Forward or the wing bulkhead, four aluminum longerons extend forward to support the motors, center gearbox mounting bulkheads, and the nose secticn. The cylin- drical fuselage section ahead of the wing is formed by two removable shells of curved alw~~inum plate.

2. Empennage Tbe vertical stabilizer has a steel spar and was atteched to the aft portion of the fuselage Sackbone, The drive motor and potentiometer for remote control and position indicator of the elevator ~ ~ 3 s housed at the base of the vertical stabilizer in the spar. A ,udder was not re- quired for this test. The hgrizontal stabilizer was mounted midway up the vertical stabilizer. The elevator could be remotely varied 220 degrees. Horizontal sta- bilizer incidence could be macually varied 2 5 degrees by changing fillet blocks &ich attached the horizontal stabilizer to the vertical stabilizer. The aerodynamic shape for the horizontal and vertical stabilizers was formed from wood panels.

Ure or r)~rclosure of ddtd on th15 m e 1 5

rcbjecl t o the rertr~ct~on on the tttle pwe I

I

3. Wing The basic wing structure was a hogged-out aluminuin chan- nel section which was closed to form a torque box by means of a bolted dn, lower surLace aluminum cover plate. At the root end of each wing box, a fork- shaped sieel root titting %as attached to the upper part of the center gezrbox, The rotor interconnect power shaft was throtigh the wing box and attached the center gearbox to the wing tip mounted rotors. Brackets were available to ad'ust the flaps (df) to 50 degrees down, and flaperons i b , ) to 20 degrees dowil. Taping of the flaperons was required to permit full span flap testing to 62.5 degrees to evaluate the effect of flap position on wing download during hover. The aerodynamic contours for the wing, flaps, and flaperons were shaped from wood panels attached to the metal spars.

4. Nacelles The nacelles were mounted oc the wing tips and housed the rotor controls, balance, and transmission. Non- structural fiberglass fairings were attached to alumi- num plate bulkheads which supported the transmission.

Conversion struts were available for manual adjustment of the nacelle from 90 degrees (helicopter) to O degrees (airplane) in fifteen-degree increments.

5. Rotors and Controls The five-foot diameter rotors were provided with re- motely controlled collective pitch and longitudinal nionocyclic pitch control for each rotor. The rotors have three blades per rotor. Each blade was mounted to a gimbal hub to permit rotor flapping. Flapping was restrained by hub springs located in the rotating sys- tem and was recorded uslng a strain gaged flexure ref- erenced to the rotor shaft. The rotor blades were dy- namically scaled in stiffness and mass distribution based on a model tip speed of 0.6 times full scale tip speed.

The model rotors were operated up to 1884 rpm which was representative of a full-scale hover tip speed of 251 meters per second (822 feet Der second). The blades were prbvided with 'strain gages for monitoring beam, chord, and torsion loads.

Cyclic pitch range was +12 degrees. The collective pitch range was from -11 to +33 degrees (measured at tip of blade). Cyclic and collective control positions were BELL H E U O O P T E R o w r w r r v subpc: to the r e r t r ~ c t ~ o n on the title paqe 5. (continued) inst~vmented for both rotors. The collective pitch could be changed simultaneous1~- f ~ r both rotors, or to the right rotor independently for trim capability.

Loads generated by the model collective and cyclic con- trol system did not cross the rotor balances, but were monitored from an instrumented pitch link for each rotor.

Rotor Drive System The rotor drive system consisted of twi, TASK motors mounted in the fuselage driving aft through flexible couplings into a coupling gearbox whic;. reduced the motor speed by a factor of three. Tl~e ; .tputs of the gearbox are coupled to the interconnect slafts with univer~al joi~ts. Wing tip gearboxes are provided for a furth?r reduction of shaft speed by a ratio of two to one. The wing tip gearboxes are :nodified Bell Fddel 47 helicopter, 90 degree, tail rotor gearboxes. Each rotor shaft was strain gaged to sense rotor torque.

The rotor shafts were extended to carry an instrumented slip ring to measure rotor blade loads, flapping, %tc., and a tachometer/azimuth wheel for driving a magnetic pulse pickup.

7. Rotor Balance A three-component rotor balance was installed in each nacelle. The primary measurement was rotor axial force for use in evaluating the rotor/airframe lift distribu- Longitudinal and lateral moments were measured tion data.

and included in the balance equations to improve the accuracy of the axial force aeasurernent in addition to providing supplemental rotor data. Dial thrust and torque bridgzs were incorporated in each balance to permit the tunnel test to continue if the signal from one bridge was lost.

The model was instrumented to measure the parameters as listed in Table 1 1 1 - 2 . This data was presented on oscillograph recorders and on the V/STOL tunnel data reccrder system for monitoring during the test. Model motor temperatures were monitored using a Brown tempera- ture recorder.

-LL Use or d~rclosure of adtd sn ihts HeUCQCTeR c o ~ w r v r sublect to the restrictinn an trz : tic TABLE 111-1. FULL SCALE~~~ODEI. SCALE PARAMETERS Scale Factors - Model/Full Scale Lerg t.h Velocity Force "ower - .

Design Parameters Full Scale I Aircraft: I C.G. Location (Mid) F.S. @ i~ = 900 W . L .

F.S. @ i~ = O0 W.L.

Design Gross Weight, N (lbf) Rotors : Blades Per Rotor 15.24 (50) Diameter. m (ft) 1 . .524 (5.0) Blade Chord, cm (in) 101.6 (40) 10.16 (4.0) Blade r w i s t , deg 25 25 RPM Helicopter 31 4 1894 Conversion 268 Air ' ze 228 Mast gth, m (ft) 3.106 (1C.19) .3106 (1.019) Hub : ~ g , m-N/d~.g (ft-lbf/deg) 1242 (916) .447 (.33) 63, big -25 -25 Conversion Axis F.S. 526 52.6 W.L. 195.4 19.54 B.L. 385.5 38.55 Win-.

2' Span (Rptor $enterline), m (ft) 19.58 (64.25) 1.959 Area, m L (ft ) 65.59 (736) Aspect Ratio 5.85 5.85 MAC, m (ft) 3.38 (11.1) 313 (1.11) Location of 1/4 MAC F.S. 514 5 1 . .4 W.L. 195.4 19.54 B.L. 180.6 18.06 Leading Edge Sweep, deg -6 -6 Dihedral, dee 2 2 TABLE 111-1. F U U SCALE/MODEL SCALE PARAMETERS (Continued) Model! Scale F u l l Scale Flape rons : 33.5 (13.2) Span Per Side, a n ( i n ) 335 (132) Chord/Wing Chord -275 ,275 Flaps : Span Per Side, ca ( i n ) 381 (150) 38.1 (15.0) Cho-ing Chord -275 ,275 Fuselage : Length, cm ( i n ) 2261 (890) 226.1 (89.C D i a n e t e r , cni ( i n ) 304.8 (120) 30.48 (12.0

ttorizontal S t a b i l i z e r : -

Span, em ( i n 23.23 lol6 (250 '400{ A - , m2 (ft ) Aspect Ratio 4.45 4 Chord, deg 15

MAC, rn ( f t 2.30 (7.56) -230 (-756) j

Location of 1/4 MAC

F,S. 1004,5 100.45 !

U,L, 291.1 29.11 Elevator:

I

.0474 (.51) Area, m2 (ft2) 6-74 (51) ChordjStabilizer Chord -265 ,265 Verttcal S t a b i l i z e r : 548.6 (216) 54.86 (21.6)

SP, q (ini

A r e a , m ( f t ) 19.0L (205) .1904 (2.05) Aspect Ratio 1.6 1.6 4 Chord, deg 32 3 2 MAC, m ( f t 3.53 (11.58) .353 (1.158) Location of 1/4 MAC F,S. 955 95.5 W . L . 201.1 29.11 Rudder : 2 2 Area, m ( f t ) 4.08 (43.9) . -

-

m U$e OI af~LOfure o f OI, ma prpc rr -arvrrWr submt to Mc reslrrton on the talk p q e TABLE 111-2. MODEL INSTRIRENTATION-3 - - - Blade beamwise loads (27.3% R) Blade chordwise ? 3ads ( 27.3% R) Blade torsion loads (27.3% R) Blade flapping Pitct- link loads Rotor torque Rotor speed and azimuth Rotor axial force Rotor pitching moment Rotor yawing moment Collective pitch position Cyclic pitch position Elevator position Wing tip gearbox temperatures Interconnect drive shaft bearing temperatures Center gearbox temperatures Motor temperatures Motor frequency control and amperage *Rotor, blade, and motor parameters were recorded for both left and right rotors.

IV. DESCRIPTION OF TESTS Testing was accomplished in the NASA-Langley V/STOL wind tunnel during two tunnel entries. The first entry, V/STOL test nunber 31, was £ran A ust 22, 1972, through September 13, 1972. The second entry, '7 V STOL test number 69, was from October 11, 1973, through November 9, 1973. Both tests were accomplished to ful- fill the same test plan; therefore, run nmbers were made con- tinuous for both tests. Total occupancy time was 660 hours.

Rotors-on testing, rotors turning, accounted for 70 hours of this Rotors-off tzsting ac- time resulting in a 18.5% utilization.

counted for only 42 hours. A total of 359 runs were made for a run average of 0.82 nm per hour.

The model was mounted on the V/STOL tunnel rotary sting support system with an internal, six component strain gage balance to Fuselage pitc-h attitude record aircraft force and moment data.

was generally varied from -16 to +2d degrees and yaw angles were Both hover a r i d forwgrd flight varied from -2 to +f 6 degrees.

was investigated. Yaw sweeps during the V/STOL test 31 were made with and without the horizontal stabilizer.

Dynamic characteristics of the model were such that removal of the em- pennage mass would cause the model to vibrate. Onlv pitch char- hrring the sec~nd tunnel acteristics with yaw were obtained.

entry, a simulated empennage nass was inserted in the tail cone fairing to allow complete empennage off testing.

Initial control positions and trim aircraft attitujes were determined prior to testing using the Bell Helicopter Company computer program C81. Static stability data were obtained dur- ing pitch and yaw sweeps from the trim conditions. Control In order to ob- settings were held cgnstant during the sweep.

tain wake effects on the horizontal stabilizer, both elevator sweeps and horizontal stabilizer incidence sweeps were made.

Tests were also accomplished with the empennage and/or the rotors removed. The model is shown mounted on ' h e sting in Figures IV-1 thraugh IV-4 for the various configurations tested.

The wing download was measured during hover at n/D ratios from Various combinations of flap and flaperon 0.525 through 1.825.

settings were tested at h/D of 1.825 to determf-ne the effect of £la? settings on wing download. Settings testq2d were 0, 20, 50, and 62.5 degree, all full span flap settings.

The rotary sting permitted the model to be set at various roll angles and h/D ratios. Rolling moment was measured to determine the in£ luence of the wing/rotor interaction cn roll stability.

Several tunnel wall configurations were tested to determine the wall interference effects on the model perfomtkce. These con- figurations were with the 1) walls up (open test section), 2) walls down (closed test section), and 3) walls do~m/slots open (slots in walls, floor, and ceiling opened).

V . DATA REDUCTION Force and moment data ~neasured or) L i ~ c y wind tunnel sis-component internal balance and rotor ba1tr11cc.s wcurc. reduced using a NASA- Langley data reduction prograni. (:o r-rccL ions were included for tunnel wall effects and interference effects of the sting support.

R e tunnel test data was tabulated in a format as listed in Also included in Table V-1 is a comparison of the Table V-1.

symbols used in the tabulated test data presented in Reference 1 and those used in this report. This is intended to give the user a means of correlating the data presented in the two reports.

Control positions and test conditions (airspeed, angle of attack, sideslip, e t c . ) are also listsd. The force and moment sign convention used for the rotor and airframe is shown in Figure V-1. The configuration code used is listed in Table V-2. A run schedule summary is given in Appendix A .

The data reference center for the internal sting balance was station line 52.6, water line 11.0, and butt line 0 . 0 . Data was corrected to a center-of-gravity location equivalent to a mid-cg as lis ed in Table V-3. Maximum Reynolds number tested

&

was 1.28 x 10 referenced to the wing chord of -338 meters ( 1 . 1 1 f e e t ) .

Airspeed presented throuout the text of this report, in the figures and the Appendix, is given as equivalent full-scale

airspeed (V . For the scale of the model tested, the

equivalent f6f1-scale value in knots is nearly equal to the model scale test airspeed (VM S ) in feet per second as shorn . .

below: eECL ltre w Q~sclosure i%!d Or tqfs we 1 5

- -v

subfit 10 the res~ritt~on dn trie 1 114 TABLE V-1. DATA REDUCTION Tabulated Data Computer Information Kotation As In Ref. 1.

Conf 1 duration I * A, Tunnel Setting V, FS 'F.s.

VTUK ' M . s .

P

i

Q ' 9 QS B. Model Attitude I ALC i

1 BETA -4

I PHI !

i

C . Rotor Controls, 1 THL, THR ' 8

f i I T I P left and right rotor i BlL, B1R

i

i ELEV ! 6 !

e i RPM R P?I

: j L

t

11. Airframe Aerodynamics 1 i

I

A . Lift ' SF

i

I

CNF I 'XF

I , CLS

I CLs CLVT i C B. Drag 1 AF AF CAF j CAF

I

I CD

I i

I CXS

!

I i i

C. P i t c h i n g Moment ' i Pn

I

I CFM

I

CM

! I

301-099-004 V-2 m b e or Qscwrure d dit, on thrs $ 5 W-aovrrrVr r u b m ID the restrxtm on the tttk pqe.

TABLE V-1. DATA REDUCTION (Continced) Information 11. Airframe Aerodynamics C. Pitching Moment (Continzed) Dm Rolling Moment RM CRM CRS Em Yawing Moment MI CYW CYPl CYMS F. Side Force SF CSF CY CSFS 111. Rotor A2rodynamics A. Thrust TAV, TL, TR average, l e f t and CTAV, C T I , CTR r i g h t r o t o r

TCT, TCL, TCR 1 Tc

S CTAAV, CTAL, Tc CT -\R I B. Torque

QAV, QL, QR ' Q

average, l e f t and CQAV, CQL, CQR r i g h t r o t o r C~ Iist or Ctstlorure .)I :dtd T, !hi< m e ' ElELL rubwl lo the rcrtrhctton 313 tnc Idle w e

--

TABLE V-1. DATA REDUCTION (Continued) I Tabulated Data f Symbol As Cornput er Information Defined Notation As In Ref. 1 In Report 111. Rotor Aerodynamics C , Moments

m, PMR

* R l e f t and right CPML, CPMR rotor Cp%

m, RMR

R " ~ CRML, CRMR C R M ~ I TABLE V-2. CONFIGURATION CODE --- - - * ~irframe/Kotor: SL:il,i 1 izcr Code Flaps ~ o r i z o n ertical Rotors

-

1 50/20 ON ON ON 2 50/20 OFF ON ON 3 50/20 ON ON OFF OFF ON OFF 4 50/20 ON ON

5 o/o ON

OFF ON ON

6 o/o

7 010 ON ON OFF

Horizontal Stabilizer: ---

Code Incidence

-

0 Tail off (Horizontal Only)

- 5 "

2 0 3 50 Empennage Off Nacelle Setting: Code

- Nacelle Incidence

0 0 2 30 4 60 5 75 6 90 (Helicopter) Code as show. in the run schedule'listed in Appendix is written: air£ ram* empennage-nacelle TABLE V-3. CENTER- 01.'-GRAVITY POSITIONS '.

Nacelle Incidence (deg) F.S . W.L.

52.9 19.9 90 (Helicopter) 75 52.6 19.5 60 52.3 19.0 30 51.7 18 . O 0 51.0 17.1 Use or d~lciortt~e 01 datd 01: Itti\ we I\ 6€U subpct lo tne rertrtctmn on the t ~ t l ? age V R - VI. RESULTS OF TEST

A. Roll Stability - In Ground Effect

Roll stability was measured in and out of rotlrtd effect by rolling the model at h/D ratios of .53, . 6 $ , .83, 1.00, and 1.84. Aircraft rolling moment measured is given in Figure VI-1.

Rolf stability versus height above the ground is summarized by Figure VI-2 in terms of the amount of lateral stick re- quired to maintain trim and rolling moment per degree roll angle between 22 degrees of roll. The amount of lateral stlck required to maintain trim was determined by dividing the model data roll control power into the rolling mcnnent per degree roll measured during the test. As iridicated, the model was f o ~ ~ ~ d to have positive roll stability near touchdown, h/D c.60. Between h/D = -60 and 1.67, the model showed negative roll stability, The maximtrzt instability occurred at a proximately h/D = .85. Above h/D = 1.67, or

WE, the mode f' was again stable. Also illustrated is the

small amount of laterd control required to trim at the maxi- mum instability height. As an example, for a ten-degree wing round-effect, the lateral stick for trim would be 1.12 cm .44 inch).

drop Roll characteristics obtained during this test were found to be in agreement with those determined during sinilar tests on other tilt rotor models for small roll angles. The non- linear roll effect shown above 5 degrees roll were not appar- ent during the other model test (~eference 2 ) .

A comparison of the results of these tests are given in Figure VI-3. With the control power available on current tilt rotor aircraft and with SCAS, the roll instability is not expected to preseat the problem as it did for the XV-3. These levels of insta- bility have been incorporated into the tilt rotor simulation e r c have received no unfavorable pilot comment concerning 1 .wer handling qualities .

Wirlp; Download Wing dobnload was obtained during hover in and out 02 ground effect by making collective sweeps and comparing rotor thrust required to hober at an equivalent full-scale gross weight. The collective pitch required to hover was in- creased by approximately two degrses from that required to hover IGE. The wing download in percent rotor ~hrust at the height to diameter ratios tested is sunnnarized in Figure VI-4.

For a full-s?an flap configuration, the download varied from 9.5 percent OGE to -6.5 percent (upload) ICE.

Ratsing the flaperons to a conversion flap/flaperon settin increased the download, OGE, by 3 percent for a total of 12. 5 percent.

Use of dtscbsure of data on t h ~ s subpct lo the ratrtctloq E" the !:tie m e , B . (continued) The effect of flap deflection on wing download for full-spax flap OGE is shown in Figure VI-5. (The faired curve shown is based on compdrison of results of other model tests as discussed in Section V T I . ) Test results indicate very little improvement in relievi,lg wing download for flap deflections above 50 degrees. For cornparison purposes the test point for flaperon setting of 20 degrees was plotted as 20-degree flap deflection. It is nearly in agreement with the full span flap results indicating that most of the reduction i : ~ download occurs from Llaperon deflection or, in general, from deflecting a sur- face which is in the projected plane of the rotor.

Rotor/wing interference was determined for forward flight at var- ious nacelle incidence angles. The rotor/wing lift sharing with nacelle incidence is shown in Figure VI-6. Comparison is shown between rotors on and rotors off to illustrate the rotor wake effect ot, the wing lift. Airfranre lift was obtained by taking the difference between the lift recorded on the main balance (airframe plus rotor) and subtracting the rcLor thrust measured on the rotor balances. Comparison with the rotors- off tests indicate that the wing lift is not influenced by the rotor wake above 40 knots, and the wing dowrlload in h ~ v e r changes to an upl-oad with forward flight.

As shown, the wing begins to contribute lift at an airspeed of 35 knots, whereas linearized rotcr wake theory would predict the wing to be immersed in the rotor wake at that low air- speed.

Flow visualization tests on another model2 were in agreement with these test results. As observed during that test, the rotor induced a strong upwash at the wing leading e2ge and on the inboard wing section at speeds as low as 20 klots. A t 30 knots, the rotor wake was nearly completely off the wing.

The nonuniform induced velocity distribution of the tilt, rotor is considered to be the reason for the rotor wake moving off the wing at these low aFrspeeds. Because of this effect, the aircraft lift was higher than originally esti- mated prior to the first tu-lnel ently. Further discussioll on differences betwsen estimated and test can be found in Section V T .C.

Static Stabili+\ Characteristics -- Stability data are shown in terms of lift, pitchlng moment, and yawing moment coefficients. Force and nioment data were obtained for rotors and/or empennage oa and o f f to e~raldatt rotor wake effects. Pitch and yaw sweeps were made at~ot~t predetermined trim attitudes for level flight. T e s L c o ~ ~ ( l i - tions are shown in Figures VI-7 through V I - 4 5 for s p e e d range of 40. to '^ri knots and nacelle incidence range o t 90 I LX- j: *xe.:rt r?! w w mn 1 9 ~ 1s

I r&*ct to the r&rclmn m thr trtk sac I

(Continued) (helicopter) to 0 degrees. Information shown is for combina- tions of test data that were used in the analysis or the rotor wake on the empennage.

Two types of procedures were followed in determining trim settings m d obtaining wake data for the empennage. During the first tunnel entry, trim attitudes and control positions were set from predetermined estimates for rotor wake effects As noted in Section VI.B., i e was on the wing and empennage, found that the rotor wake on the wing was different than estimated. This was also found to be the case for the empen- rage (to be shown later). The rotor performance versus collective pitch setting was also found to be different than estimated, As the result ~f all these differences, the air- craft was not tested in a canpletely trimed configuration during the first tunnel entry. Because of the improved lift of the ving over that estimated,the model was trimed for Level flight at one g, although not necessarily at trimmed angle of attack or cyclic control position. These correc- tions to the rotor wake characteristics were made in the theoretical estimates prior to the second tunnel entry. The procedure followed during the second tunnel entry was to set the model at a specified trim attitude and trim both lift and drag by adjusting the controls to o3tain trim thrust and power. ' R e elevator was coordinated with cyclic position to Rolling and yawing moments trim aircraft pitching moment.

were monitored to insure lateral-directional trim also.

Rotor torque was found to be a good indicator for control setting repeatability when repeating a test condition going froat tail on to off, and making incidence sweeps and yaw sweeps. During the sweeps from trim, the controls wkre not changed.

Both elevator sweeps and horirontal stabilizer incidence sweeps were used to evaluate rotor wake characteristics on the horizontal stabilizer. During the first t-~r~nel entq-, only elevator sweeps were made. Incidence sweeps were n l a d e during the second tunnel entry.

1 . Lift Coefficient A comparison between rotors on and rotors off lift coef- ficients for the range of airspeed and nacelle incidence an les tested is shown in Figures VI-7 through VI-21.

Li f t coefficient is presented versus fuselage angle of

attack and is referenced to wing area and free stream Comparison was also made between the dynamic pressure.

two tunnel entrles for several airspezds. The second VI- 3 1. (continued) entry, V/STOL t e s t 69, sanetimes had higher l i f t coef- f i c i e n t r o t o r s on t h a n t i \ r . I'i I-:.c t l t ~ t r v !it.c-trttsr t i l t * d i f fcrcnt t rittt procc'tlt~rc* ;I:- , I i sc~tsst.ti :rl$trvt-, These f iqures illustrate L ilc. iuttout~t of 1 iit sharing be t w t e n the roLor and airframe as speed increases and the nacelles are t i l t e d forward, Figure V I - 6 sumnarizes the l i f t sharing i n tenns of percent wing l i f t t o t o t a l l i f t for the condi- t i o n s tested. As shown, a t speeds above 120 knots and nacelle incidence angles 60 degrees, the r o t o r s do not pro- vide much additional l i f t t o the a i r c r a f t .

2 . Fitchine Moment Coefficient Pitching moment c h a r a c t e r i s t i c s f o r tne same range of airspeeds and nacelle incidence enqles a r e s h a m i n Figures VZ-22 through VI-36. Again, comparisons are shown between r o t o r s on, r o t o r s o f f , and f o r the two tunnel e n t r i e s . Rotor-off t e s t s were not made for all speeds t e s t e d for rotctrs on. Comparative p l o t s were made f o r r o t o r s off with airspeed t o detennine the e f - f e c t s oc pitching moment f o r the range of Reynolds number t e s t e d . The f a i r e d l i n e s shown a r e f o r r o t o r s off and are the r e s u l t of the comparison. I t w a s found t h a t f o r the speed range t e s t e d , Reynolds number did not have a large When rctors-off e f f e c t on rotors-off pitching moment.

runs a r e made a t the same conditions as t e s t e d f o r r o t o r s on, the data analyses is e a s i e r and eliminates any Reynolds ~ m - b e r e f f e c t s t h a t may e x i s t . Although soKe small differences existed, it is f e l t t h a t the analysls of the r o t o r wake is s t i l l valid. As shown, the r o t o r wake produces a nose-down pitchiilg moment during low speed helicopter f l i g h t and i s e f f e c t i v e l y reduced to that of the wing wake above 120 knots. This change is caused from the wake changing from an upwzsh during lov speeds Trends are s i m i l a r t o t o a downwash a t high speeds.

t h ~ s e obtained from the powered a e r o e l a s t i c mcdel t e s t 2 (see Section VLI f o r a more d e t a i l e d discussion of r o t o r wake e f f e c t s ) . The a i r c r a f t was s t a b l e with empennage on f o r the speeds t e s t e d f o r both r o t o r s on and o f f .

The p i t c h up a t around 8 degrees angle of a t t a c k f o r 75 degrees nacelle irtcidence a t 40 and SO knots (Figure vI-26) i s s i m i l a r t o t h a t observed on the a e r o e l a s t i c model. The pitch angle range kested f o r nacelle incidence a t 90 degrees was not large enough t o detennine if the p i t c h up occurred a t thpt nacelle incideilce.

Conlparisons made between the two model t e s t s indicated the same trends would occur at 90 degrees. To properiy analyze t h i s ef- f e c t would require the velocity d i s ~ r i b u t i o n of the r o t o r a s the model changes angle of a t t a c k .

3 . Yawing Moment Cczfficient Directional stability for the airspeeds and nacelle in- cidence angles tested is shown in Figures VI-37 through VI-45. Yaw sweeps made during this testi were limited

due to a model dynan~ic problem (Section IV ? and the time

required to make the sweep. Therefore, a complete set of rotors-on and rotors-off data combinations were not obtained for each airspeed as tested for longitudinal stability. Where rotors-off runs are not available, runs from other airspeeds are shown for comparison.

During low speed helicopter flight, directional stability was shown to LE improved with rotors on over that for rotors This was apparently due to the stabilizing ettect off.

of the rotors since with the empennage off, the aircraft was directionally stable. As airspeed increased, the rotor wake effects became less effective. This empennage configuration has iower directional stability for side- slip angles less than 8 degrees than that above 5 degrees.

These are the characteristics of the fuselage/emperige configuration tested and have been shown also during other model tests. At small sideslip angles the wake from the ~ingffuselage intersection reduces the dynamic pressure at the base of the f i n . At higher sideslip angles, the fin is in the free stream which proviaes increased sts- bility.

4. Rotor Performance Rotor performance was measured during hover in and out of ground effect. helicopter. conversion, and airplane flirht.

Data presented in Figures VI-46 through 1 ' 1 - 5 6 show the average power ccefficient and thrust coefficient of both rotors for the airspeeds tested during a collective pitch sweep.

Fieures V I 4 6 and V I - 4 7 are a comparison of the effect of tunnel wall interference on rotor erfomance. With the tunnel walls up (open test section ! ? there was a slight increase in power for the same collective pitch setting.

Thrust incredsed for 20 knots and decreased at 4 0 *knots for the same collective pitch setting. Figure VI-49 shows rotor power coefficient in ground effect compared with the faired curve out of graund effect from Figure VI-46.

Ground effect was shown not to have an influence on rotor power.

Some ground effect on thrust was evident in Figure V I - 5 0 .

Veq- little change occurred until the model was at a h/D = .53 or near touchdown.

Rotor performance is summarized at trim attitude for the airs eed and nacelle incidence angles tested in Figures VI-5 7 and VI-58.

The test data are compared with estimated performance using the digital flight simulation program.

Pretest estimates were made using linearized rotor wake n ~ < t - I &=%-,;re .=' &?A , r , rtlii ., t u r n ? b W r ~ l r x * w &lir the 11iic

--

4 . (Continued) theory. Post-test estinr;ltcs are made using rotor wake effects based on tunnel Lest results. 'L'he primary dif- ferences between these t w i ) ::tethods are the trim angle of attack and w i n g downloading. i'retest estimates were used to set the model trim attitude; therefore, cmpari- so- made with post-test estimates are not necessarily at tbe same angle of attack.

Preliminary analysis made between tunnel entries indi- cated differences between estimated and test. As a re- sult. improvementswere made to the rotor data tables in one of the digital simulation programs (~81). The tilt rotor simulation program used in post-test estimates has been shuwn to be in agreement with the full-scale wind tunnel rotor test. In order to use the tilt rotor digi- tal simulation program, a chan$e was required to make fuse- lags drag and blade twist at the three quarter radius input parameters, As shokm in Figures VI-57 and VI-58, the post-test estimates for pgwer coefficients are in closer agreement with test than pretest estimates, Both estimates for thrust coeffici~nts were i n close agreement with test, Reynolds number effects on the airframe and rGior k - e r e not accounted for in the estimates, TkLse effects were accounted for in the analysis of dacs from Reference 2 and indicated that closer agreement could be obtained between estimated and test. Because of tine, t%is =as not included in this analysis since these effects were previously established, 5. Rotor Flapping Rotor flapping angles measured at trim attitude are shohn in Figure &-1-59 for the airspeed and ~acelle incidence ansles tested. Total flapping a ~ g l e measured was similar in magnitude to that measured during the aeroelastic model test. X comparison is shown with post-test esti- mated forelaf t and lateral flapping angles.

Rotor lateral flapping for low speed helicopter, obtained during both powered model tests, was hi~her than original'y estimated. This was found to be the result of the induced velocity representation being used.2 Pretest prediction methods used a triangular distribution of induced velocity which includes a factor to modify t h : triangular distribu- tion for fordard flignt. The factor was derived for low disc loading, low twist rotors, and has provided reasonable correlation with such rotors. This factor was revised to reflect the higher d i s c loading and twist for the tilt rotor resulting in better correlation between estimated and test.

These factors are sumdrized in Reference 5 .

Rolling Moment in Hover, Figure V I - l .

Figure V I - L . 2011 Stability Characteristics i.n Hover, ,004 eE8-L V s r -17 O n i &urz rd ddtd .^ tt; 5 ; ~ ) r .x nEUCOPTER c o - e a w v ril:pft te the r ~ \ t r t11:n j t t i r i p m e

-.12 .-. W -.W o ,a4 -08 - / 2

LATERAL S T I C K D,SPLI)CEMENT 7D TWm, xu/#

i h / c n ~ s / D E G ~ E E OF ROLL Figure VI-3. Roll Stability Summary in Hover Wing Download in Hover Figure V I - 4 .

Figure VI-5. Wing Download Variation With Flap Deflection in Hover, OGE.

BELL HEUCOPTER c : - v

0 2 0 W 60 80 100 /20 /+O

AIRSPEED, VES -KNOTS Figure VI-6. Wing Download in Forward F l i g h t , OGE.

Figure VI-7. Lift Coefficient Versus Fuselage Angle of Attack, Nacelle Incidence 9 0 0 , Airspeed 40 Knots.

Lift Coefficient Vcrsus Fuselage Figure VI-8.

Angle of Attack, Sacelle Incidence 900, Airspeed 60 Knots.

Figure VI-9.

Lift Coefficient Versus Fuselage An l e of Attack, Nacelle Incidence 909, Airspeed 80 Knots.

B E U I ,v ,, ,,,,, :8->,,,, <. 7 4 ! ,- > ,.rf 1 HEUCOPTER C O M P ~ N V \,1:11+1 1 1.1 !!it. r r \ l t . , I t:. .. ' . * ' . # , A , + - FUSELAGE A M i I S O F A T A C K , & F + E G Figure VI-10. Lift Coefficient Versus Fuselage Angle of Attack, Nacelle Incidence 900, Airspeed 100 Knots.

x F . . : >t ~ \ , : - + . 4:: .. r j + \ 6€U .:.wt i: ! I . .;." -: ' ' , . ' + r 2s..

HELICOPTER i - V .

Q

F i ~ ~ r e VI-11. Lift Coefficient Versus Fuselage An l e of Attack, Sacelle Incidence 90%, Airspeed 120 Knots.

BELL , . . .. . 5 . .

. -

. . ! . . .,,:-

HELICOPTER cm-\r r4L'.5EiJ)GE R I V G E / l r A C k ; d F aZG F i g u r e Y I - 1 2 . L i f t C o e f f i c i e n t i'ersus Fuselage l e o f Attack, Xacellc Incidence ? y g . Airspeed 40 Knnts.

Figure VI-13. Lift Coefficient Versus Fuselage Angle of Attack, Sacelle Incid$nce 7 5 0 . Airspeed 80 Knots.

-30 -20 -10 C /U 23 30 k 2 ; 5 a ~ s c ,~I\;GLE CIF ATAcK,P(F--vEGC F i r e - 1 . L i f t C o e l f icient- Yersus Fuselage A n ~ l e of .jttackl Xacelle Incidence 75', Airspeed 123 Knots.

Figure \,-1-15. Lift Coefficient Versus Fuselage Angle of Attack, Sacelle Incidence 600, Airspeed 120 Knots.

eELL HELICOPTER rm-rt Figure VI-16. Lift Coefficient iersus Fuselage An-12 o f Attack, Saceilc Incidence 608, Airspeed 1&9 Knots.

Figure 1-1-17. Lift Coefficient Versus Fuselage An le of Attack, Sacelle Incidence 305, Airspeed 120 Knots.

301-099-004 EELL . I . : - - - . .. ? I . . I , . ... ; * HELICOPTER clcww=w..r F u r - 1 . l i f t C o e f f i c i e n t \:ersus Fuselage Angle of A t t a c k , Nacelle Incidence 300, Airspeed 16C Knots.

Lift Coefficient Versus Fuselage Figure VI-1.9.

Angle of Attack, Sacelle Incidence 00, Airspeed 120 Knots.

-30 -20 - 1 0 0 1 0 20 30

FUSELAGE ANGG O F A'/ 7AC K, d k - - Figure 1'1-20. L i f t Coefficient Versus F u s e l a g e Angle of Attack, Sacelle Incidence 0°, Airspeed 160 Knots.

Figure VI-21. L i f t Coefficient Versus Fuselage Angle of Attack, Sacelle ~ n c i d e n c e o O , Airspeed 160 Knots.

BELL HFLICOPTER ~ - C l M f q h >

-3 -20 -/o 0 /n 20 3

FO S E L A 6 E AhlG/!E O F ATTACK, dF *- DEG' F i g u r e V ' I - 2 2 . Pitching Yoment V e r s u s Fuselage tln l e of Attack, Tacelle Incidence 906, Airspeed 40 Knots.

BELL HELICOPTER ~ . O M F ' ~ N V Figure VI-23. Pitching Moment Versus Fuselage Angle of Attack, Nacelle Incidence 900, Airspeed 60 Knots.

BELL HELICOPTER < ornf-.%%t Figure \'I-24. P i t c l ~ i r l g 'loc-ient i-ersus i uselagc!

Angie o f I t t a c k , Sacel l c i n c i d e n c e 900, P i r s p a e d 'Q Knots.

FEELAGE A . IGLE DF A; TAPX,+ - VEG

Figure VI-25. Pitching ?Imec+ \-zrsus Fuselage Angle of Attack, Nacelle Incidence 900, Airspeed 100 Knocs.

eELL HELICOPTER c-c-hr -30 - i O - /O 0 I0 20 30 F1,SELJIGt2- 66- AT=,+ . - E G Figure C'I-26. pitch in^ >Ioment Versus Fuselage An l e of Attack. XacelLe incidence 908. Airspeed 1 ? C Knots.

Figure VI-27. P i t c h i n Moment Versus Fuselage

An le c I A t : a c k , Nacelle Incidence

7 5 8 , Airspeed 40 Knots.

i r e I - . P i t c h i n g Xoment \-ersus i-uselage .Angle of .+Ittack. S a c e l l e Inciderace 7 j 0 . Airspeed > Q Knots.

Ficure I-1-29. Pitching ?!orient I'ersus Fuselage ?ingle 01 :\ttack. Sacelle Incidznce 750. Airspeed 12P knots.

. - F i g u r e VI-30. f itching ?iontnt 1-ersus i uselege .-ingln of :ittack, Sacell€ Incidence boo. : \ i r s p s e d 129 Knots.

FUSELAGE AAXLE OF- 4T;rAC:K Figure - VI-31. Pitching lIoment Versus Fuselage .-In-12 o f Attack, Sacelle Incidence 608. Airspeed 140 Knots.

BELL HEUCOPTER c - r > n r u % t --. \ ,

. l g u r e 1-1- 32. P itchi:;; :,or.cnt i'ersus ! usc!a;e

.-Jn;ic oi ;t tack. '.ace 1 l r i n c i d e n c e 3ii'. Airspeed l L ~ f k n o t s .

BELL HELICOPTER COMF-+-.kt.

Figure VI-33. Pitching ?foment V e r s ~ s Fuselage ,-Ingle of !Ittack, Tacelle Incidence 300, Airspeed 160 Knots.

B E U HELICOPTER coms=4%r

Q

i r e 1 - 3 4 . t ' i t c h i n g xlorent i'ersils Fuselage :l.ngle o r :ittack, Yacel l e Incidence 00, Airspeed 120 Knots.

eELL ,.. . . ' :+J ,c v HELICOPTER C : - P A % V . . l . , , . l , l , . - ' : - -,,

Q

Figure VI-35. P i t c h i n g Eioment Versus Fuselage :\nglo of A t t a c k , Xacelle Incidence 00. Airspeed 160 Knots.

B E L L H E L IC:OPTER <.C>MF.~&V Figure VI-36. P i t c h i n g :.io.~ent Versus Fuselage Angle of :Ittack, Y a c e l l ~ Incidence O O . Airspeed 160 Knots.

Y A W A N G L E , Y -- DEG

Vswing iLIornent Coefficient \-ersus Figure VI-37.

Y s w Angle, Kacelle Incidence ?no, Airspeed 40 Knots.

-4 O + 8 1 2 1 6 20

'YAW A N G L E ; y ) DECi Figure I'I- 38. Yawing PIornent C o e i i i c i e n t I'ersus Ya7*~ Angle, :iacelle Tncidcnce ?!?O, Airspeed 60 Knots.

mELL . . . .

. . . ..

. f *.- .

HEUCWTER co-%r . . .

i. icure i'i- 3". 1-awin, ?!or. ent Coef i i c i e h t 1 - G ~ S U S yaw ?.ngfe. Sacelle Incidence q C Q .

Airspeed -(> Knr 5s.

BELL

HEUCOPTER . --*-ar+

\---'-: 7 . .

Figure YI-- . ..- o- c < (>efCiciex: !-erstis

- - Y z w ' * r . c f c . ' , ; ? c E ! ~ E Ificitit:-:ce :".

r - k n o t s .

1,s I I H t l IcXX3TER rwmr

Q

. -

F i e I - . 'fawing ?lment C . o e i i i c i a n t i e r s u s l*aw * l g ~ l e , Sacelle Incidence 6 0 O , ~ i r s p ? e d 1 2 0 Knots.

Figure '.-1-&2. )-a* . % .- i n q ':o:::ent c-oc i l i c i c n ~ i'ersus Yak- ;in; l e . \;ace 1 1 o Tnc i riencc 2. I @ .

rji r s p e ~ d 12n k n o t s .

B E U .. . . . I . . . * -

H E L C-OP-R ~ : c w u s * . ,:, , * * a?. *m.- % : • * ' * - . . A * -

YAW ANGLE, cjl'- 0&6 F i ~ u r e - \-I-A3. Yatcing Yorent CoefFicient i'ersas ?-aw .An&le, b c e l l e Incidence 9 , Airspeed ? 21) Knots.

6€I.L HELICOPTER c rm*-*xr /- i / ' J 2

' -

.* Y YAN ANGLE, @ -- DEG . - * - F ' r e i ' - . 1 aiiin; "Iorient Coe i f i c i e n t i e r s u s - - c; c :zw .An;:e, Sac,ll~ I ~ c i d c n c e . , I ? i r s p e ~ c i 1 6 4 Knots.

F i g u r e '\-I-.i5. i-a~.:ing Yoment Coeif i c i e n t Sersus 1-aw .-2nsle, Sacelle Incldenc? 0°, .Airspeed 1 F C : ; z : t s .

Figure VI-46. Rotor Power/Collective P i t c h Y a r i h t i o n Kith A i r s p r e d , Xace:le incidence 90°, OGE.

. . . . . . . . . .

. . . . . . . . . . . . . ..A. + . . . . , . , . .

.... :. :: I . , . . . . . . . . . . . . . . L . . . . . . * . .

--+--7v-v.- T T T C C C C C C C - ---.. .: ...... . . . rT.t -.-:. .-.-

. . .

... 8 . . . . . . . d . . . . .

+ . . . . . ..... . . . . . i ' i i . : 8 .

t . . . - - -, . , - . . a . : : ! : ; - * . : : . , . ' !

.+. L . . . . . . . . . . .. . . , . . * . , . , . i - . . . 4 . : ' ....

. * _ . . - - : : : : . : : : : ; . i i : . . . . . I

. . . . . . ..- .* . , . . , . , . .

, .

. . * . . I !

. . .

-I-?.-_t-2---: .-.- L-.:;~i -... L-.A-&...'..-- I-.L--*.. .:. . . . * .... r . d ..

. . . . i . . . . . . . . . . . . . . . . . . . * . . . . . . . . . . . . r . * ; : : : I : : : . - , . . . . . . . . . - + . . , . .

, . , . . . i : : : : . . . . . . ' . < & . * . . . . . . . . . . . . . ' . . . . . . . , i . . . .

. . . . . . . . * . . . . . . . , . . : . .

. . . . . . . * : : : 1 _ I . . , : : !

. . . . . . . . + a , . . .

* . * . . * * . -; L .--.-.- i ---- C-I-..;.. .; 1..2.--;--.1 . . . . . . . . . . . ........ ....

.... / . t .. , . . . . . . * . . . . . . ' . .

,. + , . i . . i . ' . . . _ . . . ... . * . . 4 . L . .

. . . , . . . . . . , . ' . . ' :.. ! :. t , . :- : ' a , . . . . * - i 6 . ' . . . . . ' . * . . -

-

- . . . . . . . .

; . ; . . ' * . . . - . . . . . . . . . . . . .

. .

6 , . . . . . . . . . a . .

, . . . . . . . . . . . . . . . . I . . , . . . , -*-+-- 'L- .--.-.I-I _ . . * -.--I____- .-.-+.--*.--* r--.+.-- .-.+. ~ . . * . a , * . . , . + . . . . . . . i . . ! : : : . : I . i . . . . , .

.-. . . . i i : - . . , - : ; . , i I . : 1 . . : : . : r . t . .: ... . . . . . .

. ,

- . ' . - - *

. .

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

. . . . . . . . . . * . . , . . . .

. . . . . . . . . ' . .

. . . . . . . . . . , . . . A . . . . . . . . . . . . . . . .... . . . . ? '.. *. -* ...-.

. -. . 6 .-.-... * .................... * - - . . . . . . . . . . . . .

. . . 7 . .

. . . . . . . .

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

.

- _ --.-_ --- - .--- &--L_&--- -

. . .

. . .

. * . ' . -~ ......... * . . ................... - ................ - ...................... - - . * .............

. . . . .

. . * . . .

. .

. . .

-9 -2 o 2 4 6 8

r I - Rotor .-hrust/Collecti\-e F i t c h Yariation VLth . k i r < p e e d . Sacelle Incidence l C O . OGE.

eEU L . I C ~r . l i ~ t i f s t ~ r t 0% %+I= ' r i'. \ pr+r S \ n€UCOeTER GCWWWUT h b p * l to the r;%tf~,lto- jn *rr Itti* we Figure VT-48. Rotor Power/Thxust Variation With Airspeed, Nacelle Incidence 90°, WE.

B E U 9 . , , ..

.

HEUCOPTER COM-~U*. . . : K C ! t ' r , , < , I . t , , :-. t

&.

Figure VI-49. Rotor Power/Collective P i t . c n V a r i a t i o n With Height Above the Ground, Nacelle Incidence 9 0 ' .

BELL HELICOPTER C C I M P ~ ~ * Figure VI-50. Y.ot?r Thrust/Collective Pitch V a r i a t i o n I<ith Height Above the Ground. S a ~ e l l e Incidence 40°.

Rotor Por~er,'Collectlve Pitch Variation Figure V I - 51.

Kith Airspeed, Nacelle Incidence 7 5 0 , OGE .

R o t o r T h r u s t - / ~ o l l e c t i v e P i t c h \.-ariation F i g u r e V 1 - 5 2 .

i i i t h : l i r s p e e d . Sac2lle Incidence 75').

OG i s . .

Rotor Power/Collective Pitch Variation Figure VI-53.

With Airspeed, Nacelle Incidence 60°, GZE .

Rotor ~hrust/Collective Pitch Variagion Figure V I - 5 4 .

With Airspeed, Kacelle Incidence 60 , OGE .

Figure VI-55. Rotor Power/Collective Pitch Variation With Airspeed, Sacelle Incidence 30°, WE.

Rotor Thrust/Collective Pitch Varia&ion Figure VI-56.

With Airspeed. Sacelle Incidence 30 , OGE .

Figure V I - 5 7 . Rotor Power Variation With Airspeed and Nacelle Incidence.

40 6 3 80 la A I R S P E E D , Vks .- K1-S Figure VI-58. Rotor Thrust Variation With Airspeed and Nacelle Incidence.

AIRSPEED, &s. .- KTS Figure VI-59. Rotor Flapping Variation With Airspeed and Nacelle Incidence.

llrr ~ i r t(Iluttnlrrr kit IWJ $111 I l l , \ iqt' I \ mEl-L 5ubp~t k V# r n t r i c t w n ti11 the tttlt. imp

- -V

VII. ANALYSIS OF RESULTS The information presented in this section includes analysis of test data related to determining the effects of the rotor wake The result of on the airframe applicable to tilt rotor aircraft.

this analysis will be used in design of future tilt rotor air- craft and to update the tilt rotor flight simulation mathematical model. Results obtained during this analysis will be used to extend rotor wake eEfec5 information obtained during the powered aeroelastic model test.

A. Hover Download As noted in Section VI.A., the wing download was measured in and out of ground effect and at several flap settings. These test results were compared with other powered tilt rotor model tests29697 as shown in Figures VII-1 and VII-2.

As indicated, this test and Reference 6 are in close agreement while the download for the zeroelastic model2 is much higher.

Also, the other two model tests show the same trend in down- load variation with flap deflection in that download is not significantly reduced for flap/flaperon deflections greater than 50 degrees.

From these two figures it appears that for a flaperon setting it is possible to have anything from near 50 degrees and OGE, Several parameters were 5 . 5 percent to 13 percent download.

found to be influencing these differences and to have an ef- fect on determining hover wing download. The parameters con- sidered for the various models tested are listed in Table VII-1.

Of these, blade twist and percent of wing area under the rotor were found to be the most significant parameter in detsnnining download.

The effect of blade twist is shown in Figure \?XI-3. A 3.5 percent difference in download was measured between one model with 25 degrees blade twist and another model with 40.9 degrees blade twist. The change in blade twist was also in conjunction with an increase in blocked area under the rotor as shown in Figure VII-4. From the trends shown in these two figures, it is difficult to determine which parameter con- tributed to the increased download between the two models tested or percentage each c~ntributed. In extrapolating the test results, it appears that the blade twist change may account for about .8 percent and the blockage for 2.2 percent.

The remaining 0.5 percent is due to different flap settings.

The differences in the models tested make it difficult to establish a general-ized design chart, but these figures can be used to give the desi ner some insight into trends that flaps,

f i blade twist, and bloc ed area under the rotor have on wing

download.

use i n dnclorurr d d & on Ibis ~ q c 15 m subject m the rntrwtion on the title m.

w-aaoucw* e-

A. (Continued) The test data would indicate a download on the order of 13 percent for a tilt rotor with 40.9 degrees blade twist, This is higher percent blockage, and flap setting of 75/45.

than previously estimated for this configuration. This differe ce is attributed to a Reynolds number effect.

v

Hoerner shows that the drag of cross-sectional shapes simi- lar to a wing. with flaps deflected. at an angle of attack of This -90 degrees is highly dependent on Reynolds number.

effect is shown in Figure VII-5 for the variation of dra

f

coefficient of the cross-section with Reynolds number re -

erenced to wing chord and mean induced velocity from the rotor. Calculation required to correct model scale down- As noted on load to full scale is given in Table VII-2.

Figure VII-5, the drag coefficient was reduced by approxi- mately 50 percent due to Reynolds number changing from model scale to full scale. Therefore, a 13 percent download mea- sured for the model would be approximately 6 . 5 percent for Other configurations would be the full scale aircraft.

similarily reduced.

B. Rotor Wake on Horizontal Stabilizer The wake characteristics at the horizontal stabilizer were determined using pitching moment data from tail-off, tail-on, and incidence runs. For rotors-off configuration, the wake is the wing downwash angle ( t W H ) .

Rotors on, the wake is

the total wake angle ( c ) whic includes the wing downwash b

T; angle plus the rotor wa e (cR/~). AS noted earlier, two

methods of obtaining the wake effects were used during these tests. During the first tutncl entry, only elevator sweeps were made; whereas, during the second entry, incidence sweeps were used. The following are the equations used to determine the wake for rotors on and off including the equation for elevator and incidence.

Knowing, EIEU Me or dtschurt of drtl on this tr H w R - r subfgct t o the restrtctm on the $I!# plgc B. (~ont inued) then, giving, E = a F - ( Y H + i H + T T e ' e where, For rotors-off configurations, the same equations can be used by setting ~ R / H and q~~ to zero and replacing " and qHT with 'W/H and 'H* respectively.

The horizontal stabilizer lift curve slope was estimatedg at the test Reynolds number to allow determinGion of the dynamic pressure ratio (qH ) at the horizontal stabilizer.

T Knowing

-

SH'H 1 . 3 1 V H = r =

wcw

aH = .054/degree then, Use or disclosure d ~ t i on N i l ~ plpc is subpct to the r e r l r i ~ t ~ w , w, the !ttIc pT)e -a -v

@ " - -

B. (continued) Frm the test data, elevatar effectivene~s ( I , ) was dcter- mined from These equations were programmed to accept wind tunnel test pitching moment data. The rssulting dynamic pressure ratio and wake angle are shown in Figures VTI-6 through VII-11 for both rotor on and rotor off at each nacelle incidence ailgle tested. These figures show that for nacelle incidencz angles of 90 and 75 degrees the rotors produce an upwash and an increase in dynamic pressure. These are similar char c-

teristics as obtained during the aerc2lastic model test. 4

A constant dynamic pressure ratio was used at 60, 80, and 100 knots because only elevator sweeps were m d e at these speeds.

As mentioned above, both incidences and ele~acor sweeps were shown to be methods of determining wake characteristics on the horizontal stabilizer. As the result of t:~ese two tests, the comparison between the two methods is valid only for the angle of attack that the elevator sweep was made or for con- figurations ic which dynvnic pressure was not expected to vary much with angle of attack. This vas primarily the rea- soti for making incidence sweeps d~:-ing the second tunnel entry. Although incidence sweeps take longer than elevator sweeps, the test data are much more useful for analysis.

The measured wake angle at nacellc incidence angle of 75 degrees and 120 knots is considered questionable. At this a-rspeed, the other nacelle incidence angles show closer cgreement with rotors-off downwash. Yo error was found in re- view ; £ the analysis, but this condition is felt to be in- consist-ent with trends from this and other model tests.

The upwash from the rotcrs c-ontinually decreases a : the nacelles are tilted forward from 60 degrees until the total wake angle is nearly the same as the rotors-off wing down- wash va11-e. Some scatter exists i~1 comparing dynamic pres- sure ratic, but in general becomes that of the rotors-off value also.

As notei earlier, tttl total wake angle (cT) at the horizontal stabilizer is made up of the wing downwash ( ~ w / H ) and the rotor wake ( E ~ / ~ ) . The rotor wake is represented in the tilt Ole or dtsclosurc ol data on tkts plpe IS 6EU subpct to thu restrictton on the tctk p*lc WE-Roomr*rrr B. (continued) rotor simulation math model in terms of the rotor induced velocity. It is computed from the ratio of rotor induced velocity in the plane of the horizontal stabilizer to the rotor induced velocity at the rotor disc ( w ~ R / ~ / w ~ ~ ) times These two terms the mean rotor induced velocity ( w i R ) .

give the velocity of the rotor wake at tne horizontal sta- In order to obtain these parameters the bilizer (vH~).

following equat:on was used.

= ) w. IR = sMmsm tan R/H

Theref ore, tan ' ~ / d ' i ~ (13) Rotor induced velocity (MiR) used was that calculated from the tilt rotor simulation math model corrected to model scale. The induced velocity ratio determined from this test and the aeroelastic model test are summarized in Figure VII-12. Both tests indicate the same trend with airspeed and show that the rotor wake in the plane of the horizontal is effectively eliminated above 120 knots for all nacelle incidence angles. With this type of wake reaction on the horizontal stabilizer, trim aircraft attitude becomes more nose down and longitudinal stick gradients are shallower than originally predicted with linearized rotor wake theory.

It should be noted that the downwash velocity (VH~) is merely a convenient way to represent the rotor wake effects on the horizontal stabilizer and does not represent the actual wake from the rotor. This is illus~rated by the fact that the vector sum of the free stream velocity ( V M * ~ , )

a ~ d the downwash velocity (vHR) is not the total velocity

indicated by the total dynamic pressure ratio. Furthermore, the variation of downwash velocity with airspeed implies that it increases with airspeed, which is opposite to the momen- tum theory of rotor induced velocity. With the test data available at this time, it is not possible to separate the effect of the rotor wake on the wing downwash. Wing lift on the inboard section is considerably different with rotors on than with rotors off, and would change the contribution of the wing lift on the wing wake at the empennage. However, 301-099-004 V f 1-5 USE or d~,closure of data on t h ~ s page Ir I m L L c c t h ~ r t to the trslr~ctron on the lrlle ~ J P HEUOOPTER -V B. (continued) c~rnbining the dynarriic pressure ratio (VHT) and downwash (ET) does give the correct empennage lift. Additional testing in this area would be required to obtain pressure distribution and local flow velocity data to properly describe these ef- fects. Sideslip was found to reduce the magnitude of the up- wash on the horizontal stabilizer as shown in Figure VII-13.

The downwash velocity was determined by assuming the total dy- namic pressure ratio to be equal to its value at zero side- slip. The reduction in upwash velocity causes a nose-up pitching moment when the aircraft is sideslipped. Compari- son is also shown with the aeroelastic model test. Wake effects on the two empennage configurations are nearly the same. Free-flight testing of the aeroelastic model and the rotor simulation tests have shown that only a small amount of longitudinal cyclic stick is required to correct this pitch up and to maintain pitch attitude with yaw.

C. Rotor Wake on Vertical Stabilizer Rotor wake effects on the vertical stabilizer were partially evaluated in terns of a parameter defined as the rotor side- This is defined as the ratio of the ver- wash factor ( K g ) .

tical stabilizer yawing moment rotors on to rotors off and indicates the change in dynamic pressure ratio and sidewash due to the rotors. Kp is defined as follows: Rotor sidewash factor ior the conventional vertical fin tested The was compared with an H- ;ail2 as shown in Figure VII-14.

airspeed shown was the only case in .c~hich data wer= available for both rotors on azd rotors off. T h e conventional fin is shown to have mere directiozzl stsbliity during low speed beli- copter flight with rotors on than he H-tail configuration.

In com2arlng other configurations tested for the conventional vertical fin with the H-tail, several other items were noted.

Rotors off, the directional stability of the conventicnal f i r ) Is reduced for sideslip less than 8 degrees whereas the H-tail has more linear stability characteristics with sideslip. This is generally the c.ase throughout the speed range tested. It would indicate that the effective-less of the coi~ventional fin is reduced at small sideslip angl,~s due to the wing/fuselage wake. At low sideslip angles, the H-tail is outside this wake.

At the higher sideslip angles, the conventional fin passes outside the wake to increase stability. With rotors on, low ( Continued) speed helicopter flight, the rotor wake effect is also dif- ferent between the two vertical tail configurations. The W-tail fins were shown2 to I,e near the center of the vortices of the rotors which reduces clircckicwal stability due to reduced dynamic pressure. As the aircraft is sideslipped, the vortex cores shift with respect to t . . e fuselage center- l int to increase the e f f ectivencss of the vertical fins.

The ,onventional iin is located outside of these vortices in an ai-ea which causes it to have increased stability over that of the H-tail. As airspeed increases and the nacelles are tilted forward, the rotor wake effect decreases and the stability characteristics are similar to that for rotors off.

A canparison of the two fin configurations is shown in Figure VII-15 at zero fuselage angle or attack. A t first appearance, both fin ~o~~figurations show close to the same level of stability. Also shown are the results of addition- al testing at small angles using the Reference 2 model with a conventional fin. An additional reduction in stability was found for the conventiorlal fin between 52 degrees. The level of stability provided by this conventional fin con- figuration was found to be inadequate. Increasing the fin size to give the same level of stabilitv as the H-tail would have required a very large fin. The difference in the two configurations became more apparent at angle of attack when the conventional fin becomes more immersed in the win?/ fuselage wake. It low-speed characteristics were the only consideration, the conventional fin wocld appear to be better, but most of the low speed directional stability is provided b y the rotors and SCAS. (Tilt rotor simulation tests, SCXS- off, have shown the handling qualities of the aircraft to be adequate and controllable with the reduced stability of the H-tail during low speed helicopter flight.) In high speed flight, where the fin is more effective and provides the major portion of the directional stability of the aircraft, the H-tail configuration with its linear characteristics here found to be more desirable.

- TABLE VII-1. i4OVER DOWhnOAD PARAMETERS (FULL SCAr,E) TR-71-626 01- 0 -999-002 200-094-2707 T e s t 31/39 I I W*ng ~rea(l), 54.07 15.61 10.78 65.59 m (ft2) (582) (116) (706) (168) Wing Chord, m(ft) 3.38 2.62 1.59 1.14 (11.1) (5.22) (3.75) (8.6) Disc Area, mZ(ft2) 182.46 220.7 45.61 38.55 (1964) (2376) (491) (415) 37 . O 40.9 Blade Twist, deg 1 25 10/20 7.62 Rotor 3im, m(ft) 15.24 16.76 7.62 ( 2 5 ) ( 5 5 ) ( 2 5 ) (50) .65 .75 SW'/SW(~) .62 ,705

.275 .22,/.44 . 25 - 2 4

cf/% ,204 ,161 .I87 ,125 tm/D (1) Measured between centerline of rotors, SW Wing arc a under rotor disc, SW' ( 2 )

L

TR-71-626 301-099-002~ 200-094-2707 Test 31/69 t 1/10 Model Scale 1/20 1/5 1/4 .0085 -111 -0042 .I147 CT VT, m/sec (fps) 228.6 169.77 Model Scale 150.26 100.58

(750) (557) (493) 1 (330)

Full Scale 225.55 228.6 212.14 1 2;;?;; (740) (750) 0396) wiR('), m/sec (fps) Model Scale Full Scale ~ ~ ( 2 ) Model Scale 4.8~105 3.04~105 8.0~106 3.36~106 Full Scale (1) Rotor induced velocity cmputed using the d i g i t a l tilt r o t o r simulation program

1 . 5 * W ~ R * CW

(2) R x = v Figure VII-1. Wing Download Comparison in Hover.

Figure V I I - 2 . Download Comparison in Hover f o r - ~ l a ~ Deflection. K , E .

BELL ,, ' . . : : - . . - HEUCOPTER c r a w - i r + . t - .. ,, .. . .-- * . _ : % - Figure \-11-3. Wing Download Comparison i n Hover for Rlade T w i s t , OCE.

Wing Download Comparison in Hover Figure VII-4.

for Blocked Area, OGE.

! i---&T.! I $ ! r .---+ I I

, . > I 1 z

L ,&L I

REF 8, FIG 13, SEC 3-9 (BASED ON CROSS-SECTIONAL SHAPES WITH ROUNDED EDGES AND FLAT SIDES)

--

. ; ' / .

1 : : ( : j : : , . .

8 . $ i !

i 1 : ! .

I , . IK HOVER i I REYNOLDS NUMBER, Rx Figure VII-5.

Variation ~f Wing Drag Coefficient at a = -90' with Reynolds Number.

W Figure VII-6. Horizontal Stabilizer Aerodynamic Characteristics. Nacelle Incidence 90°.

Figure VII-7. Horizontal Stabilizer Aerodynamic Characteristics, Sacelle Incidence 7 S 0 .

B E U HEUCOPTER c(3rrr~4ruv

Q

-6 -9 o 4 8 1 2

FUSELAGE ANGLE O F A T d C K , d F c 'DElf Figure VII-8. Horizontal Stabilizer Aerodynamic Characteristics, Nacelle Incidence 60°.

BELL HELICOPTER C O M P 1 \ N V Figure VII-10 Horizontal Stabilizer Aerodynamic Characteristics, Nacelle Incidence 0'.

E E U ;.!5t fj, a ~ s ~ o $ g ~ r > L'I tt.11.4 .-G ! I > , \ ju:,t, HELICOPTER COMWNY \LIL;P,I If1 ~ P ~ I I ' L ~ I G , ~ , J l l It'? lillr Figure VII-11. Horizontal Stabilizer Aerodynamic Characteristics, Nacelle Incidence 0°, Flaps Up.

301-099-004 VII- 20

w 60 80 /OO

AIRSPEED, V 6 s 8.- &rs

Figure VII-12. Induced V e l o c i t y Ra-io i r Plane of the HorizonCal Stab'l? ?e..r.

-- -. ~ - - --

-"I c 7' 63 1 2 /&

X4N ANGLE, (u- E G

F i g u r e i - 1 1 - 1 3 . E f f e s L of yaw Angle on Horor Lake Up.cash at Horizontal Stabilizer, Nacelle Incidence 900.

301-090-004 Figure VII-14. Effect of Rotor Wake on Directional Stability, Nacelle Incidence 750.

Figure VII-15. Effect of Fin Configuration on Yawing Norsent Coefficient .

The following conclusions pre made from analysis of the test data :

A, Roll Stability - IGE

Static roll stability characteristics obtained during hover tests in ground effect were in agreement at small roll angles A roll in- with that determined during other model tests.

stability is obtained as the aircraft approaches touchdown between h/D = 1.67 and -60 with the max- instability occurring at h/D = -85. This instability does not appear to present a significant problem for current generation tilt rotor aircraft with the improved control power avail- able and addition of SCAS.

B . Wing Dovnloaci Characteristics of wing download were determined during hover in and out of ground effec~with flap deflection,and in for- ward flight. Hover download in addition to being a functfon of h/D and flap deflection was found to be influenced by blade twist, percent blockage area of the wing under the rotor, and Reynolds rumber.

Very little improvement in relieving wing download during hover was obtained for flap deflections above 50 de rees.

Results indicate that most of the reduction in down f oad is

due to flaperon deflection rather than flap deflection. Down- load was also found to increase with increasing blade twist and the blocked area under the rotor. Hover download measured during model tests are subject to Reynolds number effect. For the configurations tested, the download for the full-scale air- craft would be approximately 50 percent l o > : e r than the model scale value. The full-scale aircraft is estimated to have a 6.5 percent download OGE, Rotor/wing lift sharing during forward flight was determined.

Results indicated that the wing lift is not influenced by the rotor wake above 43 knots and that the wing begins to contribute lift at an airspeed of 35 knots. Above 120 knots and nacelle incidence angle of less than 60 degrees, the rotor wake was found to not contribute much to wing lift.

Rotor Kake on the Empennage C.

During low speed helicopter and conversion flight the inter- action between the rotor wake and the horizontal stabilizer C. (~ont inued) is such to produce a nose down pitching moment. This results from a net upwash effect and increased dynamic pressure on the horizontal stabilizer from the rotors. Rotor wake ef- fects are essentially eliminated at airspeeds above 120 knots.

Rotor wake effects on the vertical fin during low speed heli- copter were found to be less apparent on the conventional ver- This is not compatible with tical fin than for the H-tail.

the low directional stability characteristics observed for the XV-3 which also had a conventional vertical fin. Addi- tional empennage testing usicg the same model would be de- sirable to describe the rotor wake effect for different em- pennage configurations. Directional stability of the H-tail during high speed flight is more linear with yaw angle than the conventional fin.

As the result of these tests and analyses, the following iterns have been incorporated in the tilt rotor simulation math model: 1 . Tm roved the wing download variation with flaps, ground

P

ef ect, and airspeed; 2 .

Modified the induced velocity variation at the empennage; and 3.

Modified the roll stability in-ground-effect.

h e M d t r c b ~ u r e ol a t & an this p*)e ir aEu- rubpcl to the rntrtcilon (M the ttt& pyc

--

TX. LIST OF REFERENCES Wilson, . I . : "NASA-l.an&ley Kc.st.arch Center V/STOL Wind Tunnel 'Test Kesults or I'owered I.'orce 'Tilt Rotor Aircraft," NASA- Langley 'TEi, to be published.

BHC Report 301-099-002, "V/STOL Tilt Rotor Study - Volume VI, Hover, Low Speed and Conversion Tests of a Tilt Rotor Aeroelastic Model," NASA CR 114615, May 15, 1973.

RHC Report 299-159-084, "Proposal for Wind Tunnel Tests of a Powered Tilt Rotor Force Model," March 1972, BHC Report 301-999-001, "V/STOL Tilt Rotor Study - Volume V, A Mathematical Model for Real Time Flight Simulation of the Bell Model 301 Tilt Rotor Research Aircraft," NASA CR 114614, April 13, 1973.

Plarr, R . : "Handling Quaiities Evaluation of the XV-15 Tilt Rotor Aircraft," American Helicopter Society Preprint No, 840, Yay 1974.

"Volume V, Wind Tunnel Test of a Powered Tilt Rotor Anon: Performance Model," AFFDL-TR-71-62, October 1971.

BHC Report 200-094-270, "Results of the Wind Tunnel Tests of the Quarter-Scale Semi-span Model of the Bell XV-3 Tilting-Rotor Convertiplane," September 18, 1958.

Hoerner, S. F . : Fluid-Dynamic Drag, 1965.

USAF Stability and Control Datcom, Air Force Flight Anon: Dynamics Laboratory, &right-Patterson Air Force Base, October 1960 (Rev. February 1972).

McKee, J. W . : "Experimental Investigation ok the Drag of Flat Plates and Cylinders i r . the Slip Stream of a Hovering Rotor," NACA ITu' 4239, April 1358.

!I.* or dtsclo;ure cf M a O n MIS pap lr eELL -oormrvr \u. ml t o the rcrtrtc:m M the litk plpc APPENDIX A Run Schedule Summary ( A description of the configuration code is given in Table V - 2 , page V - 5 . ) m E l L Use or d~rciorure of dlU on MIS j w p ir rubpct to the rntr~ctton on the trtk p q e -coumWr RUN SCHEDULE SUMMARY MODEL C100-F1B LANGLEY V/STOL TEST 31 C CONFIG.

RUN i~ VF.S RPM h/D NO.

NO.

(DEG) (KTS) 1-2-6 90 1884 0 Walls Up - Coll.

10 Sweep

0 - Cyclic

20 - Coll.

2 0 - a

- 4'

15 Void 16 40 - Coll.

17 ii 18

I - a Void

+

19 Walls Up - +

20 20 Walls Dn - Coll.

40 - Coll.

::

I - a

23 - Cyclic

- A81

25 - AB1

26 - #

2 7 'I

- be

1-2-6

2-0-6 - ET

- 4 '

V 3 1 40 - Coll.

3 2 80 - Coll.

33 Void

- 4J

90 35 Void

I

O 2-0-6 36 1-2-6 3 7

I - i

40 80

- #

- Coll.

41 120

- Coll. 120

43 - Coll. 100

- 6e

45 - Cyclic

I

46 - A6

- c r

47 100 I Cyclic 48 1 - '18~84

- V

49 15 40,80,100.120 1-2-5 1604

4il,80,,.00,120

- - V Coil.

50 51

- (Y

5 3 6 e

54 - Cyclic

5 5 v - A6

1 9 / 3 4

56 40 Walls Dn - AB1 1654

15 Sweep 1-2-5 i A Ult w dtxhnurc ol data a MIS plpc IS eELL sub@ (D MC ~ c I ~ K I M MI thC 11th w.

n-oovwm RUN SCHEDULE SUMMARY MODEL C100-FlB LANGLEY V/STOL TEST 31 ( Continued) CONFIG.

RUN i~ " F . S RPM h / D NO.

NO . (DEG)

(KTS j

57 Walls Dn - Coll. Sweep 80 15

- a

6e

- Cyclic

- Coll.

62 - Q!

- 6e

64 - Cyclic

- A9

66 120 Walls D n - AB1 15 1-2-5

1-2-6 67 1.84 0 0 Walls Up - Coll.

68 - Cyclic

- M l

70 - 9

- AB

71 1.84 72 1.00 - Coll.

- a

74 - Cyclic

75 - 9

6e

77 - A0

1-00 78 - Coll.

.83

. 8 3 - 9

80 .67 - Coll.

81 .67 Void

82 .67 - 8

83 .53 - Coll.

84 .53 - Cyclic

85 .53 - Cyclic

.53 V 5

86 - 8

t

1-2-6 87 0

Walls Up - de

88 100 2-0-6 Walls Dn - Coil.

89 - CI

100 0 2-0-5

90 - Coll. 40 15 2-0-5

91 - CY 40 15 2-0-5

92 - Coll. 80 15 2-0-5

93 - a ! 80

94 - Coll. 120

95 - o !

97 - Cyclic

98 - A0

99 Walls Dn - AB1 Sweep 30 1-2-4 1604

m€u- Use or d~rclorurc d &ta on th~s w e i f subject b t 4 rntrwtton on the lttle paye WEUOOPtER-r RUN SCHEDULE SUMMARY MODEL C100-FIB LANGLEY V/STOL T E S T 31 (Continued) COMFIG RUN i~ VF.S h/D NO.

NO.

(DEG)

(KTSJ

1-2-4 1604 30 140 Dn - Coll.

100 Sweep Walls

30 140 1-2-4 - a

1-2-4 30 140 102 6e 60 1-2-2 120 103 - Coll.

- a

- 6e

I I

- Cyclic

60 1-2-2 2-0-4 30 108 I 3 1 .

30 120 2-0-4 - Q, 2-0-4 30 140 110 - Coll.

- OL

30 1 4 0 111 2-0-4 2-0-2 60 120 112 - Coll.

60 120 2-0-2 - Coll.

60 2-0-2 120 114 - Coll.

2-0-2 1604 - QI

60 120 115 80 1-2-5 1884 - Coil.

15 116

- CY

80 117 1- 2- 5 80 118 6e 120 119 - Coll.

- a

120 120 ii'

I

80 121 - Coll.

- CY

80 122 0 2 0 - Coll.

- CY

2 0 124 40 125 - Coll.

40 - (Y

2 0 127 7 - 4~

40 2-0-6 1884 Walls Dn - $ S w L p

I End of Powered Test 1-2-6 0

0 129 Walls Up - 8 Sweep

Rotors Off Test

80 3-2-6 W~lls Dn - a

- a

120 121

- LY

I

120 160 1 1 ; Sweep

160 - $ (QI = 5")

136 - iL (QI = lo0)

- JI ( a = -lo0)

- de Sweep

-

1 139

6e

. I 1 3-2-6

0 160 140 0 Walls Dn - a Sweep

6Eu- Ow or d15clo\11re ul ddla on t h ~ r ~ a g c 1% W E U C O P T E R -v iubpct lo the r n t r ~ c t m on the title pdgp RUN SCAEDULE SUMMARY MODEL C100-FIB LANGLEY V/STOL TEST 31 (Continued) i T CONFIG.

RUN i~ VF.S h / ~ RPM NO.

NO.

(DEG) (KTS j

160 141 3-2-6 0 Dn - J1 Swee

Walls 3-2-6

160 142 - 0 ( a = -5'7

120 3-2-6 - 6 , Sweep

i 160 144 3- 3- 6

- be

0 160 145 3-3-6 - a !

- Q

15 120 146 3-2-5

160 147 3-2-5 - 4

15 160 3-2-5

148 - 4 Sweeg

15 160 149 3-2-5 - J1 (a = -10 j

30 120 i 50 3-2-L - a Sweep

3- 2-4 - Q

30 160 151 '

- 4

30 160 152 3- 2-4

- CY

45 120 153 3- 2- 3

- Q

45 160 154 3-2-3

45 160 155 3- 2- 3 - 9

- Q

60 120 156 3- 2- 2

- Q

60 160 157 3- 2- 2

- 9

60 158 3- 2- 2

- a

159 3-2-0 160 3-2-0

- 6e

161 3-2-0

- 9

- CY

162 4- 0- 0 90 163 4-0-0 - J,

- Q

60 4-0- 2 60 165 4-C- 2 - IC,

4-0-4 - (Y

30 166 30 4-0-4 IC,

0 168 - C Y S w ! e p

4-0-6

0 169 4- 0- 6 - 3, ( a = 0")

0 V 170 4-0- 6 7 7 - $ ( a = l o 0 )

160 4-0-6 Walls Dn - IC, ( a = -10")

0 1 7 1 0 eELl Use or dtrclosurt of data on l h ~ r D * J ~ 15 HEUOOrreR-r subpt to the r c r t r i c t m on the fillc paqe RUN SCHEDULE SUMMARY MODEL C100-FIB LANGLEY V/STOL TEST 69 RUN CONFIG.

~F.s. i~

NO.

(KTS) (DEG)

-

Coll. Sa 176 Waij 1-2-6 40 Q!

177 1-2-6 40 90 17 8 4J 1-2-6 40 90 (Y 1-1-6 40 93 a 1- 1.05 (Y 181 1-1-5 a 182 1-2-5 Coll. 183 1-2-5

JI 184 80 i " ? 75

1-2-5 Coll. 185 1-2-4 120 60 a 186 1-2-4 4J 187 1-2-4 a

1-1-4 I ! o

a 18 9 1-1-2 120 3 0 a 190 1-1-2 160 3 1-2-2 120 Coll. 192 a Coll. 194 Cyclic 195 6 , Sk Roll T a r 197 3 Sa 198 Roll T a r 199

v

a ! s tl

200 160 3 0 eP Roll T a r 201 0 Thrust C ecks 203 Coll. Sweep 120 a !

204 120 205 JI 120 Coll .

206 160 a !

208 JI Su 160 t eP 209 Thrust C ecsks 210 Thrust C ecks

Q sk

2P 212 120 6e a 213 160 6 , Coll.

Q 6 , 218 Coll 180 I \ 219 180 220 Wall 180 @

/

MU ttre or dircbsure of d a b en page r r HeUOOP.TBRootanam subject to the restriction on the t r t l e me.

RUN SCHEDULE SUMMARY MODEL C100-FIB LANGLEY V/STOL TEST 69 ( Continued) CONFIG i~ R U N l NO. DEG ) I NO.

- Coll. 2 2 1 Wal

- fr

2 2 2

- 6e

2 2 4

- $J

- Coll.

- OI

- $ .

2 39 2 4 1 2 4 2

- a

2 4 4 J I - a 2 4 5

- a

2 4 6 - 9

- Q

2 4 9 - 9 a 2 5 1

- *

- a

- *

- a

2 5 4 Kal

- $

- Coll. 2 5 6 Wal

- ff

- o !

- o !

- CY

2 60

- a !

- a !

2 6 2

- a !

263 r

- a !

- a

- 9 266 Wal

265 I

BELL the w dtxbrure 01 bra, on m t s ~ q c t i -owrrm subpct ID thc ratrtctlon on the (I* pap RUN SCHEDULE S U E W R Y MODEL C100-FlB LANGLEY V/STOL TEST 69 (Coi~t inued) CONFIG.

NO. RUN 1 NO.

a 2-4-6

lwalls Up - Sweep

-

a

'

-

a I

I 1

I a

I

Walls Up - 2-4- 6

Q Sweep a

Walls Dn - Sweep, Slots Open 1-2-6

a

I I

Q Q

I

-

Q Sweep, Slot: open

-

a

-

a

-

Q

-

Q

I

- a

1-2-6

-

Coll. 5-2-6

-

5-2-6 Coll.

-

Coll. 5-2-6

-

Coll. 5-2-6

-

Roll 1-2-6 Roll

-

Roll Roll

-

Roll Sweep

-

a !

Sweep 1

-

Q 1-2-6

-

(r :-2-4

-

1-2-4 9c,

-

(Y 1-2-2

I

-

1- 2- 2

-

a 1- 2- 2

-

1-2-2 #

-

a 2-0-0

-

a 2-0-0

-

CY 2-0-0

-

a I 6-9-0 Rotors O f f

-

a 3-2-6

-

a 3-2-6

-

a 3-1-5

-

a !

3-1-6

t - a

3-1-5

t

IWalls Dn - Sweep 3-1-5

RUN SCHEDULE SUMMARY MODEL C100-FlB LANGLEY V/STOL TEST 69 (Continued) RUN QDNFIG.

VF.S, i~ RPFl h/ D NO. NO.

(KTSI (DEG)

--

312 3-2-5 Sweep Walls Dn - a bo 1 75

313 3-2-5 8 0 - # 75

- a

314 3-2-6 60 90

- Q

315 3-2-6 93 316 3-2-6 60 90 - # a

317 3-2-6 - 9 120 90

- a

318 3-1-6 8G 90

- a

3-1-4 120 60

320 3-2-4 - a

321 3-2-4 60 - #

- a

322 3-2-2 ! 30

323 3-2-2 30

- $

3-1-2 - CY

324 30

- a

325 3-1-0 0 -- Q 326 3-2-0

- Q

327 3-2-0

328 3-2-0 - # 1

329 3-2-0 120

- i+

330 160 - #

- a

- b e

333 I - 6 . Swzep

334 160 - a Sweep, Upside Dn " I

i

335 60 - a Sweep, Upside Dn 6- 2-0

i

336 160 - & Sweep, Rt Aileron Dn 9-2-0

I 337 9-2-0

160 - # Sweep, Lt Aileron Dn 0

I

- a

339 80 75 Sweep 3-2-5 i

- Q 3-2-4

340 120 60 i

- Q

341 3-2-2

- a

342 30 4-4-2

- a

3$3 30 4-4-2

344 I 30 4-4-2

- J,

345 - C Y

4-4-0

346 - a

347 - #

I

343 - + 6-4-0

- a

349 0 6-4-0

I

- (Y

350 60 4-4-4

I

120 - II, 60 4-4-4

!

- a

352 4-4-6 60 90 35 3 4-4-6

60 - # 90

354 4-4-6

120 - Q 90

I - (Y

355 129 4-4-6

- Q

356 60 75 4-4-5

- a

357 80 75 4-4-5

-

358 I t 4-4-5 80 75 8 'I

!-

Walls 160 Dn - cr Sweep 4-4-5

301-099-004 A- 9