..... _:'_;''' ' 00000001

ADVANCEMENTOF- PROPROTORTECHNOLOGY

,. TASK[ - DESIGNSTUDYSUMMARY , (NASACONTRACTNAS2-5386) (N.kSA-CR-1111682) ADVkliCBBENT OF PE_ROTOE NTq-11833 TECUNOLOG¥. TASK I- DESIGN STUDY SUNM_R¥ (Bell--HelicopteE Co,) 209 p HC $12,50 CSCL 01C Unclas G3/02 23270

REPORT 300,099"003

; _'.' ',' ._ :_-)1 _I"1 i|,_ COMPANY

POST OFFICE _OX 482 f-ORT WORTH TEXAS 76101 _ _ COMPANY 00000001-TSA03

13FL.L HELICOPTEI_ (:OMI-'ANY--

CoNT._NtS ,"

I • SUMMARY I-l

II. INTRODUCTION II-I

A. The Need for VTOL B. The Failure to Meet the VTOL Need II-2

C. The NASA Proof-Of-Concept Program II-2 D. The Tilt-E_oprotor Aircraft - A Promising VTOL Concept II-4 E. The Tilt-Proprotor Proof-Of-Concept Aircraft Evaluation Program II-6

..;; III. DESIGN DESCRIPTION IIl-1 7_i A. General iII-i .....I B. Design Criteria III-9

_'_t. C. Proprotor 111-10 _" i. Blades 111-10 2. Hub III-i0 D. Drive System 111-11 E. Powerplant 111-12 I. Engine 111-12 2. Induction System 111-12 3. Oil System 111-13 4. Fuel System III-13 F. Airframe III-14 i. Wing III-14 2. Fuselage 111-15 3. Empennage 111-15 4. Landing Gear 111-16 G. Aircraft Systems 111-16 i. Conversion System 111-16 2. Hydraulic System 111-17 3. Electrical System 111-17 H. Aircraft Controls 111-18

I. Proprotor Controls 111-18 2. Flight Control Tli-19 3. Stability and Control Augmentation System 111-19 4. Proprotor Governor System III-20 5. Power Management 111-20

300-C99-O03 i O00000Ol-T.qAn4

J (_ BELL HELICOPTER C_O_IPAN_'

CONTENTS - Continued

IV. WEI_{T ANALYSIS IV-I

A. Rotor Group IV-3 B. Wing Group IV-3 C.- Tail Group IV-4 D. Body Group IV-4 E. Alighting Gear Group IV-7 F. Controls Group IV-8 G.___Eng[ne Section and Nacelle Group IV-8 H. Propulsion Group IV-9 i. General Description IV-9 2. Engine Installation IV-9 _<_, 3. Conversion System IV-9 4. A_r Induction System IV-9 _'- 5. Exhaust System IV-9

_, 76.. LFubricatiel Systemon System Engine IV-IOIV-9 8. Engine Controls IV-IO 9, Startin$ System IV-lO IO. Rotor Pltch Governor Control IV-10 ii. Drive System IV-10 I. Instrument Group IV-12 J_Hydrauiics Group IV-14 K. Electrical Group IV-14 L. Electronics Group IV-14 M. Furnishings and Equipment Group IV-15 N. Air-Conditioning Equipment Group IV-15

V. APERFORMANCE. Summary V-iV-I B. Airframe Aerodynamics V-i I. Lift Analysis V-3 2. Drag Analysis V-3

I C. Proprotor Performance V-4 i. Description V-4 2. Method and Correlation V-6

I 3. Isolated Proprotor Performance V-7 D. Powerplant Performance V-7 !

I 300-099-003 it 00000001-TSA05

I i_ B FILH__.E _LI COPT£R C:O_rAN v ! CONTENTS ,ontlnued

V-8 E. IIelicopter PerformaDce i, Hovering PerformaDce V-8

i 2. HelicopterRequired Level Flight Power V-8 3. Rate of Cli_b V-9

F. Airplane Performance V-9 i. Thrust Horsepower Required V-9 2. Flisht Envelope and Maximum Speed V-9 3. Maxlmum Rate of Climb V-10 4. Specific Range V-IO 5. Single Englne Performance V-10 i 6. Mission Profiles V-lO ,_ 7. STOL Performance V_ll VI. DYNAMIC STUDIES VI-I

':''Lk=-_._" A. Proprotor Dynamics VI-I ,_r'_; I. Natural Frequencies Vl-i _"_-- 2.- Loads VI-2 _':" 3. Blade Flapping VI-4

B. Airframe Dynamics Vl-4 i. Natural Frequencies VI-4 2. Vibration Levels VI-5 C. Stability Vl-6 1. Proprotor VI-6 2. Flight Mode Stability Vl-7 VII. NOISE Vll-i

VIII. RECOMMENDED PROOF-OF-CONCEPT PROGRAM VIII-I A. Objective and Purpose VIII-I B. Program Plan VIII-I i. Phase I - Design and Fabrication VIII-2 2. Phase II - Flightworthiness Tests VIII-2 3. Phase III- Proof-of-Concept Flight Research VIII-5 C. Program Schedule VIII-6

IX. REFERENCES IX-I

X. DESIGN LAYO_ffS x-l

300-099-003 iii 00000001-TSA06

BIEI .I 14El .ICOPTER Cn_I',_,N',"

LIST OF FIGURES

N__oo. T_tl______e Page

I-1 Bell Model 300 Proof-Of-Concept Tilt- Proprotor Aircraft 1-3

II-i XV=/._Convertiplane 11-8

Version 266 I 11-2 Operational of Model Composite Aircraft 11-9

] II-3 NASA-Ames_40-by--_0-Fo25-Foot Proprotor Performanceo_t Full-ScaleTest Tinunnelthe If-10

II-_ 25-Foot Proprotor Dynamic Test on Simulated Wing in the NASA 40-by-S0-Foot Full-Scale Tunnel ll-ll

!_._ ,_ 11-5 Folding Proprotor VTOL Aircraft 11-12 4 ]_"" II-6 Mockup of Representative Civil Tilt- Proprotor Aircraft 11-13

11-7 Civil VTOL Aircraft Mission 11-14 11-8 Military VTOL Aircraft Mission II-15

IIl-i 25-Foot Proprotor Parameters III-22

111-2 Proprotor Stiffness and Mass Distribution 111-23

111-3 Wing Stiffness arid Panel Point Weight III-24

111-4 Fuselage Stiffness and Panel Point Weight 111-25

III-5 Vertical Tail Stiffness and Weight Distribution III-26

III-6 Horizontal Tail StlffnesE and Weight Distribution III-27

III-7 Collective Pitch Versus Conversion Angle 111-28

III-8 Differential Collect_ve Pitch Versus Conversion Angle 111-29

I 111-9 Fore and Aft Cyclic Pitch Versus Conversion Angle 111-30

I III-i0 Differential Cyclic Pitch Versus Conversion Angle 111-31

I !II-ii Differential Cyclic P_tch Versus Airspeed 111-32 I 300-099-003 {v 00000001.T_An7

BEL_! HELICOPTER _;r_._t'AN_"

LIST OF FIGURES - Continued

IV-I Wing Weight IV-16

IV-2 Fuselage Weight gstimation Parameters IV-I7

IV-3 Fuel System Weight IV-18

IV-4 Gear Stage Weight IV-19

IV-5 Main Transmission Schematic and Weight Data IV-20

V-I Mode[ S00 Aerodynamic Model in 7-by-lO-Foot Tunnel V-12

V-2 Airframe Lift Coefficient Versus Fuselage ,: Angle of Attac_ V-13

V-3 Airframe Drag Coefficient Versus Fuselage _?_[ Angle of Attack V-14

_ _i. ' V-4 Airframe Drag Coefficient Versus Lift _i. :. Coefficient V-15 _ V-5 Comparison of Full-Scale Aircraft Drag, Based on Tunnel Results and Calculated Data V-16

: V-6 Proprotor Blade Station Data, Radial Station 0.075 to 0.45 V-17

V=7 Proprotor Blade Section Data Blade

• Station 0.45 to 0.70 V-18 V-8 Proprotor Blade Station Data, Radial

I Station 0.70 to 0.90 - V-19 V-9 Proprotor Blade Section Data, Radial Station 0.90 to 1.00 q-20

I V-10 Proprotor Performance Correlation V-21

I V-If Boeing-VertBlade Parameolters[3-F(Bladeoot Diameter"E") Proprotor V-22

Versun Thrust V-23 i V-12 Proprotor Hovering Power Required V-13 Proprotor Efficie_icy Versus Shaft I Horsepower Sea Level V-24 V-if+ Proprotor Efficiency Versus Shaft

I Horsepower i0,000 Feet V-25

I 300-099-003 v 00000001-TSA08

O 131::I I HEI-ICOPTElY (:C'IMI_ANY

LIST OF FIGURES - Continued ¢.

V-15 Proprotor Efficiency Versus Snaft V-26 Horsepower 20,000 Feet

V-16 Twin Engine Proprotor Shaft Horsepower Available, Helicepter Mode V-27

V-17 Fuel Flow, Helicopter Mode V-28

V-18 Proprotor Shaft Horsepower Available Takeoff and 30-Minute Power V-29

V-19 Proprotor Shaft Horsepower Available, Normal Rated Power V-30

V-20 Fuel Flow, Airplane Mode V-31

: ,- V-21 Hover Ceilings_ V-32 j# _/ V-22 Hovering Figure of Merit V-33

:_ V-23 Proprotor Shaft Horsepower Required Versus .... ' True Airspeed, Helicopter Mode V-34

V-2q Lift Distribution Between Proprotor and Airframe in Helicopter Level Flight V-35

V-25 Rate of Climb, Helicopter Mode V-36

V-26 Thrust Horsepower Required Versus True Airspeed, Sea Level V-37

V-27 Thrust Horsepower Required Versus True Airspeed, i0,000 Feet V-38

V-28 Thrust Horsepower Required Versus True Airspeed, 20,000 Feet V-39

V-29 Proprotor Shaft Horsepower Required Versus True Airspeed, Airplane Mode, Sea Level V-40

V-30 Propro£or Shaft Horsepower Required Versus True A_rspeed, Airplane Mode, i0,000 Feet V-41

V-31 Proprotor Shaft Horsepower Required Versus True Airspeed, Airplane Mode, 20,000 Feet V-42

i V-32 AirplaneFlight EnvelopeMode and Maximum Speed, V-43

V-33 Maximum Rate of Climb, Airplane Mode V-44 I

300-099-003 vi m, E 00000001-TSA09

II_ I]I-:I .I H-FI .ICOPTER _:c,.,_,^'_v

LIST OF FIGURES - Continued r,

V-34 Nautical Miles Per Pound of Fuel Versus True Airspeed, Airplane Mode, Sea Level V-45

V-35 Nautlcal' Miles Per Pound of Fuel Versus True Airspeed, Airplane Mode, 10,000 Feet V-46

V-36 Nautical Miles Per Pound of Fuel Versus - True Airspeed, Airplane Mode, 20,000 Feet V-47

V-37 Single Engine Performance V-48

_: V-38 - Research Mission V-49

V-39 Civil Mission_ V-50

% V-40 Military Mission V-50

_'_!: ._I V-41 Ferry Mission V-51 ,_ " " V-42 STOL Takeoff Performance V-52 VI-l Proprotor Collective Mode Natural Frequencies VI-10

VI-2 Proprotor Cyclic Mode Natural Frequencies VI-ll

Vl-3 Blade Oscillatory Stress at VH, Helicopter Mode VI-12

VI-4 Blade Limit Stress at VH, Airplane Mode VI-13

VI-5 Flapping Envelope in Airplane Mode VI-14

Vl-6 Airframe Symmetric Mode Natural Frequencies_ VI-15

VI-7 Airframe Asymmetric Mode Natural

) Frequeneies VI-16 VI-8 Pylon and Crew Station Vibration VI-17

VI-9 HelicopterPylon and CreModew Station Frequency Response, VI-18

I VI-lO PylonAirplaneand ModeCrew Station Frequency Response, Vl-19

i VI-II Mode]. 300 Proprotor Stability Boundary VI-20 Vl-12 Correlation of Calculated and Measured Proprotor Stability - NASA-Bell Test of

I Model 266 Aeroelastic Model VI-21

I 300-099-003 vii O0000001-TSAIO

(1_ F;IF-I I HF_LICOPTER {:()MP4_'Y

LIST OF FIGURES - Continued "

VI-13 Model 300 Aeroelastic Model Proprotor Stability (Full-Scale Frequency and Airspeed) VI-22

VI-14 Model 300 Aeroelastic Model in 7-by-lO- Foot Tunnel VI-23

VI-15 Model 300 Full-Span Model in NASA- Langley 16-Foot Transonic Dynamic Tunnel VI-2q

VI-16 Comparison of Stability Analysis Method VI-25

VI-17 Root Locus of Symmetric Free-Free Modes VI-26

VI-18 Root Locus of Asymmetric Free-Free Modes VI-27

_ _.., VI-19 Proprotor Stability Sensitivity to Wing Stiffness and RPM VI-28

1-,J_ VI-20 Root Locus of Longitudinal Short-Period

_ Mode VI-29 Vl-21 Root Locus of Longitudinal Short-Perlod Mode of Basic Airframe VI-30

VI-22 Root Locus of Dutch-Roll Mode VI-31

VI-23 Root Locus of Dutch-Roll Mode for Basic Airframe VI-32

VII-I Calculated Noise Level of Model 300 VII-3

I Ames 40.-by-80-Foot Tunnel VIII-8

VIII-I Flightworthiness Test Article in the i VIII-2 Proof-of-ConceptSchedule Program - Suggested VIII-9

. VIII-3 Proof-of-Concept Program - Alternate I Schedule VII!-I0

I LIST OF DRAWINGS 300-960-001 Three View X-2

I 300-960-002 Proprotor and Controls X-3 300-010-001 Blade Assembly X-4

i 300-010-100 Proprotor Assembly X-5

i 300-099-003 viii 00000001-TSA11

LIST OF DRAWINGS - Continued

300-960-003 Inboard Profile Nacelle X-6

300-960-004 Main Transmission X-7

300-960-005 Fixed Controls - Wing X-8

300-960-006 Fixed Controls - Fuselage X-9 300-960-007 Wing Assembly X-10

300-960-008 Fuselage and Empennage X-If

300-960-009 Crew Station and General Arrangement X-12

' LIST OF TABLES

... IIl-i Basic Data III-2

- ' III-2 Dimensional Data I_I-3

__h_ III-3 Control Travels III-7 _ _'_" IIl-q Basic Design Criteria III-9

IV-I Group Weight Statement IV-I

IV:2 Model 300 Mission Weights IV-3

IV-3 Ta_l Surface Unit Weights IV-q

IV-4 Fuselage Weight Penalties IV-5

IV-5 Fueslaga Parameters and Symbols IV-6

IV-6 Rolling Gear Component Data IV-7

IV-7 Alighting Gear Group Weight Comparison IV-8

i IV-8 Drive System Data IV-If IV-9 Transmission Special Provisions IV-12

IV-10 Instrument Group Weights IV-13 IV-II Electrical Group Weights IV-14

I IV-12 Furnishlngs and Equipment Group Weights IV-15

i V-I Performance Summary V-I

i 300-099-003 ix 00000001-ISAI2

I

Rr=l I HF_I.ICC'JPTF_R _;_ml'aNY

],IST OF TABLES - Continued r-

V-2 Parasite Dr_g Correction Parameters V-5

V-3 Drag Breakdown - Full Scale A_rcraft V-6

V-4 Engine Losses V-8

VI-I Model 300 Pitch Link Loads VI-3

VI-2 Empennage Natural Frequen_Cies VI-5

VI=/-- Phugoid Mode Characteristics_ VI-8

Vl-4 Dutch Roll Mode Shape and Phase VI-9

VI-5 Time to Half For Rolling and spiral Modes VI-9

"! ,o

! _L "

I J i

! I 7 I I

-i I 300-099-003 X 00000001-TSA13

I

I, S_MARY

tilt-proprotor proof-of-concept study Aconducted under NASA Contract NAS2-5386.aircraft Thedes_.gnresults are has orebeen- sented in this report. The objective of the contract is to

I andadvancefull-scalethe statewind-tof unnelproprotortests.technTheologyspecificthroughobjectivedesign stofudies Task I, which is now complete, is to conduct preliminary design

J stthatudiescan toperfodefinerm proof-of-concepta minimum-size tilt-proprotorflight research.reseaTherch sta_rcrudies,aft which wexe based on prior work done under Bell [_licopter Company Independent Re_earch and Development include aircraft-layouts,

J weight estimations, performance calculations and dynamic analyses. The aircraft that results from these studies (Figare I-l), ! ' designated the Bell Model 300, is a twin-engine, high-wing air- J craft with 25-foot, three-bladed tilt proproters mounted on pylons at the wingtips. Each pylon houses a Pratt and Whitney - PT6C-40 engine with a takeoff rating of 1150 horsepower. E_pty ',... weight is estimated at 6876 pounds. The normal gross weight is l:_r_'._ 9500 pounds, and the _maximum gross weight is 12.400 pounds.

_. The maximum level-flight speed of the Model 300 is 312 knots at "i"[_ 15,000 feet. The a_rcraft can hover out of ground effect at 6,400 feet and 95°F at the normal gross weight. The 4000 foot, 95°F hovering ceiling occurs at a weight of 10.300 pounds. ItB suitability for flight research and simulated civil and military missions is analyzed and found to be more than adequate to demonstrate proof of the concept.

Performance, weight and .dynamie analyses of the Model 300 are based on the layout drawings in Section X. The results are substantiated statistieally and by model-test data obtained from previous Bell IR&D programs. Drag estimates are based on one-fifth-scale model wind-tunnel tests. The method of esti- mating proprotor performance is correlated with NASA tests of a 13-foot-diameter proprotor. The proprotor-pylon stability analyses shows good correlation with a one-seventh-scale aero- elastic model tested in the NASA-Langley 16-foot Transonic Tunnel, and with a one-fifth-scale semlspan model of the Model 300 tested in a low-speed wind tunnel. High torsional stiffness of the wing provides a speed margin for proprotor-pylon stability of over 70 percent of the limit dive speed--far in excess of the 20-percent flutter-free margin required. The methods used to predict the important characteristics of the Model 300 appear to be valid, but they will be further substantiated under Task II of the contract, which will include the collection of full-scale test data in the NASA 40-by-80-foot tunnel, using a 25-foot proprotor of the same design as that for the Model 300 proof-of- concept aircraft.

As a part of the contract, a long-range proof-of-concept program has been developed. The proposed program involves three phases:

300-099-003 I-i ...... --...... - __ __ L . • , . --" .,. ] 00000001-TSAI4

1

BELl. HELIc]OPTf_R _:o_l-^n_

design and fabrication of the alrcraft, flightworthln_ss tesl:_, o.. and proof-of-concept flight research. It is directed toward establishing proof of concept in three areas: technical, economic, and environmental (noise, downwash, etc.). The schedule calls for the first infiight conversion in the third quarter of 1972, in order to permit proof-of-concept flight research to atart early in OY 1973. Alternatives to the basic program, including the--folding-proprotor concept, are discussed.

It appears that the urge,,t need for civil and military VTOL transportation can be met by tilt-proprotor aircraft. Imple- mentation of a proof-of-concept flight-research program i8 the next logical step toward meeting that need.

300-099-003 I-2 00000001-TSB01

fIf_l tif-I I(;_Jf'|| I_ _ cJM_'AN_ ...... -- 1:1 00000001 -T Rn9' '

RF! J. HELICOpTeR _X_F_N_

I].. INTROI)IICTION

Military and civil planners are becoming aware of the need for aircraft with VTOL capability in a variety of military missions and for civil transportation. The following par_- graphs discuss this need and appraise the reasons for the failure of past efforts to meet the need. A NASA proof-of- concept program which could provide the first step for the logical and expedient attainment of operational VTOL aircraft is outlined. The qualifications of the tilt-proprotor VTOL aircraft for early operational application are discussed and the role of the proof-of-concept flight research aircraft is defined.

A. The Need for VTOL

The need for practicable VTOL aircraft in both military and civil transportation grows continually more urgent (References , i through 13). On the military side, VTOL is essential to the requirements for fast reaction, operational flexibility, and economy of effort in such missions as airborne assault, local-area defense, antisubmarine warfare, and tactical logistics.

I_ systemThe "Sixcan-Daybe War"if itof is196dependent7 showed onhow largevulnerableairfields.an air-defenseViet Nam has made one point clear: the ability of our armed forces to operate efficiently in the underdeveloped regions of the world must be retained and enhanced. Initially at least, airfields will be few or nonexistent; VTOL aircraft may be the only means of access to vital zones of combat. Even in highly developed nations, VTOL would be essential to counter an enemy who had in- capacitated existing airfields and ground lines of transport.

In Viet Nam the helicopter (a slow, short-range VTOL) has proved its value in airborne assault, ground support, surveillance, medical evacuation, and rescue. Its remarkable usefulness is a good indicator of the greater 9_OL potential that could and should be developed. Larger, faster, and longer-range VTOL aircraft can realize the full capability attainable for these and other mis- sions.

On the civil side, commercial aviation can be expected to con- tinue its rapid expansion. Airline operators are preparing for

j transportthe introduwillction cutof flyingfaster timeand forlargertranscontinentalaircraft. The andsupersonicinter- continental flight in half. Giant transport aircraft with

capabilitiesin 1970. Yet ofthenearlya_r traveler500 passengersmay receivewill bepoorerenteringserviceservicein terms of travel time to his ultimate destination. Air traffic has already overloaded our airports, not only in their ability

btoaggage,handle butthe alsoairplanesin movingand passengersload and unloadout topastheisengersr finaland desti- nations. Decreasing flight times will no longer slgnif_cantly

I shorten total trip time. VTOL aircraft can help to avoid this I

I 300-099-003 II-i 00000001-TSB03

(_ BELL HEUCOPTER (_O_ANY

stagnation in civil a(r transport, and open new opportunities .. to benefit from the rnpid pr_a_es_of seronautics.

A VTOL transportation system can relieve congestion at major airports by carrying psssengers between cities up to 500 miles apart, using convenient VTOL ports near their points of origin and destination. VTOL transportation would also be availabte to carry passengers from the major regional airports into nearby cities.

The development of such a system will require the cooperation of local and federal planners in integrating the necessary facilities ] and services, and formulating regulations for the operation of the system. Most important, VTOL aircraft that can provide the desired services at reasonable cost, with low noise, ground dis- turbance, and environmental pollution, and with high standards of safety must be developed.

B. The Failure to Meet the VTOL Need ! i_r'_ Despite the urgent need for VTOL aircraft, their development _,.i" has been painfully slow. In 1968, the Director of Defense Research and Engineering told Congress that the US had built ._i.: 17 V/STOL aircraft during the preceding i0 years, and had spent

morebest designthan halfapproacha billionfor VTOLdollarsis withoutnot yet obvious.zeal results.A greatThe variety of VTOL concepts have been proposed, and many have been fllght-tested. Some have demonstrated their technical feasi- bility, but none has been selected for production and operation in the United States. The various concepts have suffered tech-. nlcal problems, undesirable operational characteristics, or economic shortcomings, or combinations of these faults.

The "requirements/technology dilemma" has also impeded the development of military VTOL capabilities. The Department of Defense has been slow to establish requirements for VTOL air- craft because of a lack of demonstrated VTOL mission capa- bilities. Conversely, industry has been reluctant to develop the needed technology without the military requirements. The dilemma even more strongly affects the development of civil VTOL aircraft.

Although VTOL technology is still not adequate, concerted national goals can and should be formulated to guide its develop- merit .

C. The NASA Proof-Of_Cgncgpt Program It is in the national interest to proceed with VTOL development

withof theallmostpossibleessentialexpediency.steps in TheaccomplishmentNASA can readilyof thisperformobjective:two (i) it can develop and evaluate VTOL technology. (2) It can I determine and demonstrate the capabilities and limitations of 300-099-003 II-2 I O0000001-TSB04

(1_ BELL HELIC..._OPTER C:()MF'AN¥

the promising VTOI, types, thereby permlttln_ operational specifi- cations and requirements to be realistically prepared In short, the requirements/technoloKy dilemma is resolved. Once these steps have been taken, industry, military, and civil planners can move swiftly to develop the needed operational aircraft for introduction into VTOL transportation systems.

NASA efforts can be most effectively focuaed in a proof-of- concept program. A proof-of-concept program should be conducted for each of the most promising VTOL concepts to provide and verify technology and to determine the suitability of the concept to fill the role of the civil and military VTOL aircraft. A proof-of-concept flight research aircraft should be tested to obtain the necessary data to determine the suitability of the concept for future development and service.

The program should include testing to establish proof of concept : in three specific areas:

_ , .,, Technical Proof of Concept

The aircraft's performance, flight characteristics and

determine if technology is in hand for the successful • development of en operational aircraft with the desired problemtechnical areascharacteristics.would be investigated and evaluated to

Economic Proof of Concept

The ability of the aircraft to perform a variety of possible missions would be investigated. Payloadlift capability, range, endurance and fuel consumption would be measured. These data along with operation analysis inputs would be used to determine economic feasibility of the concept. The cost effectiveness of the concept and its competitive position with alternate means for accomplishing specific missions could then be established.

Environmental Proof of Concept

The desirability of using the concept for particular civil and military missions would be evaluated to determine if the aircraft is a "good neighbor" and to determine its suitability for operation in batt].efield environments. This would provide data on flight safety and safety to ground personnel, internal and external i noise levels, and the effects of downwash and recircu- lation on engine injestion, visibility and dust signa- ture. Approach and landing procedures would be evaluated

I uforndevopeleorationped areasfrom toairportsprovide anddataheliportsfor the developmentas well as of navigational systems and to determine reel estate

requirements for future VTO]. ports and to permit 300-099-003 TI-3 I

i i i i i I I I I II I _ • .... ] ii li BELl- HELICOPTER COMpANy

definition of low-speed maneuver requirements of mi].itnry aircraft.

The proof-of-concept approach offers a way for this nation to make up for lost time and acquire the VTOL transportation sys- tems that are so urgently needed.

D. The Tilt-Proprotor Aircraft - A Promisin_ VTOL Concept

Of all of the VTOL concepts that have been investigated, one appears to offer great promise for an effective transportation system. This concept, the tilt-proprotor, has the charaoteris- tics necessary for economic feasibility and social accepta- bility in both civil and military roles. Economic and operational effectiveness require a high-payload vertical-lift capability, as well as high-speed cruise efficiency. Attempts to achieve these objectives with configurations that tel F on high-disc- loading devices for lift have been generally unsuccessful.

Recently,largely becinterestause of continin low-disc-loadingued technologicaVTOLl progresshas been andrenewed,the ._ demonstrated operational usefulness of the helicopter.

_ _ Studies have shown that rotor-lifted (iow-dlsc-loading) VTOL 9* J aircraft can have good cost effectiveness. Of the severa[ con- cepts for low-disc-loading VTOL, the tilt-proprotor has been

missionsshown to holdrequiringthe mosextendedt promisehovering.for transportComparativemissionsoperationsand analyses conducted by Bell Helicopter Company, Reference 14,

haveproprotorshown aircraft.the significInantotheradv-staDtudies,ages inherenBell, tWestland,in the tilLockheed,t- Sikorsky, Boeing, and the Marine CorDs have reached essentially . the same conclusion (References 2, i0 and 15 through 20).

The proprotor is an efficient lifting device that makes it possible to control hovering position precisely and to maneuver at low speed in confined areas. In hlgh-speed cruise it func- tions efficiently as an airplane propeller. It has the added advantage over some other concepts of requiring only one pro- . pulsion system for both flight modes. Since there is no weight t penalty for duplicate powerplants or propulsion devices, the ratio of payload to gross weight can be high. Consequently, the proprotor aircraft can realize high levels of mission effectiveness over distances much longer than helicopters or compound helicopters can fly. Arranging the proprotors side- by-slde makes the overall span of the lift system large. In forward flight in the helicopter mode, this feature keeps power ' requirements low, and provides exceptionally good STOL char- acteristics. This latter capability could be used to advantage in many missions where the initial takeoff can be made from an airfield.

_: The low disc loading of the tilt proprotors will minimize noise

, 300-099-003 II-4 i and dust; the low-downwash velocities will contribute to the O0000001-T£RNR

{_ BELL HELICOPTER COMPANY

safety of ground personnel. The low noise level should maI

Flight safety is enhanced bF ";he simple conversion process M_ich may be stopped or reversed at any point. The aircraft may be flown continuously with the proprotors at any conversion angle from vertical to horizontal. Its control characteristics make it safe to maneuver, climb, or deseend during conversion or reconversion. Because its rotor-lifted speed range overlaps its wing-lifted speed range, the conversion corridor is wide, and such parameters as airspeed or power need not be scheduled with conversion angle. Steep descents are possible_ there are no restrictions on the approach angle. In the event of complete loss of power, a conventional helicopter-like autorotational flare and landing is possible. The power-off reconversion capability, which was demonstrated with the XV-3 Convertiplane, - makes it possible to enter autorotation from any flight mode. _- At airspeeds of 150 knots and above, power-off reconversions can ,-_.K_ be performed without loss of altitude.

proprotor aircraft's capability to meet the requirements for civil and military transportation can be demonstrated.

_ WhileThe technologyproprotor istechnolosysufficientlyhas wellrecentlydevelopedbeen advanringthat the tilt-rapidly, I the concept is not new; zts technological development has been under way for more than twenty years. Efforts to develop the technology date back to the 1940s and this early work led to I the initiation of the Joint Air Force-Army XV-3 Convertiplane | program in 1951 (Figure II-I). The flight evaluation of the XV-3 by Army, Air Force, and NASA pilots demonstrated the sound-

I hessproprotorand safetycould beof usedthe coequallynversion wellprinforcipleliftandandshowedpropulsiothatn. a These tests also defined dynamic stabilit F problems that required

i further analysis and'correction. The evaluation tests were completed in 1961 and reported _n References 21 to 23. The program included more than 375 hours of

I wiinnd-tunnel125 hours andof flightground-runtime.time,The antestd moreaircraftthan 250was testflownflightsby t_n Government test pilots and two Bell pilots, who made a total of

I morepilotsthanmadeii0power-offfull conversioreconversions. nsFivefromof thecruiseGovernmeto helint coptertest autorotation after simulated engine failure.

proprotor government I Much of the recent work has been under sponsorship as part of the Army's Composite Aircraft Program. These efforts have been concentrated on resolving the technical

I technologicalproblems uncoveredbase forduringthe thedesignXV-3 andtests,developmentand on providingof modern a high-performance tilt-proprotor aircraft. The Composite A{r-

I craft Exploratory Definition phase was completed i_ September

i 300-099-003 II-5 00000001-TSB07

BEll- HELICOPTEr C()MPANY

1967, with the design of the 28,000-pound, 360-knot, Bell Model 266 (Figure 11-2), This work has been reported {n detail in Reference 24 and in summary in Reference 25. The program to con- struct and test the demonstrator aircraft was not implemented, primarily because R&.D funds were not available.

During the Composite Aircraft Program, design solutions for all the known proprotor problem areas were found, and the technical risk in a full-scale development appeared to have been mini- mized. Some uncertainty remains, however, since the design was substantiated only by analysis and the results of small-scale model tests. Large-scale model testing._and/or a demonstrator aircraft program can f_nally prove the adequacy of proprotor technology for application to the design and development of such VTOL aircraft. Bell Helicopter Company has designed and is fabricating, as part of its IR&D program, a flightworthy 25- foot-diameter proprotor suitable for full-scale wind-tunnel testing.

... The NASA-Ames Research Center has contracted with Bell for a _--_4_q'_ wind-tunnel test program in which the 25-foot proprotor will . _ be used to obtain data on performance and blade loads and to / investigate dynamic stability of the rotor-pylon-wing system in airplane flight. These tests, which are scheduled for 1970,

thisare illcontractedustrated workin Figincludesures 11-3a anddesign11-4.studyTheof firsta proof-of-phase of concept aircraft. This report presents the results of that study.

Under separate NASA Contract, Bell is conducting a design study of a more advanced tilt-proprotor concept, the folding proprotor (Figure II-5), which promises to extend operating speeds into the 400-to-500-knot range. The Air Force has also initiated a program (with several study contracts) to develop the technology for this advanced concept. The Bell-NASA folding-proprotor contract also calls for the design and fabrication of a 25-foot- diameter folding proprotor for full-scale wind-tunnel testing.

E. The Tilt-Proprotor Proof-Of-Concept Aircraft Evaluation

The proof-of-concept aircraft must be capable of undergoing a flight research program which will provide the verification as to whether or not the tilt-proprotor VTOL can meet the demands for future civil and military transportation. This means that the test aircraft and its test program must provide the data necessary to evaluate the economic and social acceptance of the tilt-proprotor aircraft as well as its technical characteristics. These proof-of-concept requirements are met in the tilt-

i thisproproaircraft,tor aircrafthte dBellesignModelpresen300,ted canherepositivelyin. It is demonstratebelieved thatthe technical, operational, economic, and environmental suitsbility I 300-099-003 II-6 I 00000001-TSB08

BEI-L HELICOPTER C_PA_'

of the tilt-proprotor VTOL to fill the role of c_vil and m_litnry _ transports. The mockup shown in Figure 11-6 is representative of a civil VTOL of the size and design of the Model 300 proof-of- concept aircraft.

The relatively small size of the aircraft permits it to be tested in the NASA-Ames Full-Scale Wind Tunnel and is a result of the desire to make the proof-of-concept program as economical as possible without scarificing any program objectives. In this respect the aircraft is a minimum size; however, test results and conclusions will be applicable to the largest VTOL transports envisioned at this time. The aircraft can operate in level flight to speeds of 300 knots and dfve to higher speed to investigate aircraft dynamics. The Model 300 aircraft has been designed efficiently and eoafigured so that it can show economic proof of concept by demonstrating the capability to perform a variety o£ civil and military missions.

___ Theaccommodationsthree view ofwhichthe coaircrafuld bet providedin Sectionfor X sshowsuch missions.the pas_se_ngerTypical missions which could be simulated by this aircraft are depicted _ in Figures II-7 and II-8.

i L The aircraft can investigate the effects of noise and downwash during takeoff and landings at a disc loading from 7 to 12-1/2 '-- _ pounds per square foot. The low noise signature of tilt- proprotor aircraft will be typified by the Model 300. Section VII of this report presents the predicted noise level.

A noise pressure level of 90 PNdb would be experienced 300 feet from the hovering aircraft. This compares with a noise level range of 80-100 PNdb for automobile and truck traffic noise at : 50 feet from a busy downtown street.

_i Takeoff, approach and landing evaluations can be conducted in

of flisht procedure on the noise and dust generated and to determlne the terminal navigation and traffic control system :i ureqndeuvirementseloped areforas,VTOLs.heliportsBecauasend atirporhe Modelts to300evalhasuategoodtheSTOLeffect _ performance capability, operations in this mode c_n also be investigated. A proof-of-concept flight research program with the Model 300 tilt-proprotor aircraft will define requirements and establish technology, thereby paving the way toward realizing the VTOL transportation systems so urget.tly needed.

300-099-003 II-7 00000001-TSBO9

I

_, ,. I

| ', I i I

00000001-TSB11 O0000001-TSB12

RF_I .I HF",I IC()PTIL_R C:[)AtI_ANY

Figure II-4. 25-Foot Proprotor Dynamic Test on Simulated Wing in the NASA 40-by-80-Foot Full-Scale Tunnel., 300-099-003 lI-il

O0000001-TSB14 ' " ..... , I, i, I I i I=-- - I I-'I i I lii lli i I 00000001-TSC01

(_ BELL HELICOPTER COMPANY

t-_ 0 ra 'H

Lr4

Z q_ ,_o _ _ o

I _ _ .,4 1 _ "_°

| _' I4 , H _ I-4

°H I

_ _ o

| =aoM O | . _ _

I 300-099-003 II-IA_ m -_ ___ _ mm

BEI-I- HELICOPTER _,n/_PANY O0000001-TSC03

(I_ BFLL HELICOPTER ,c:oMI_'ANY

Ill. I)EgIGN DESCI_IPTION _.

A. General

The Model 300 tilt-proprotor proof-of-concept aircraft has twin proprotors ae the tips of a forward swept high wing. The prop- rotors are mounted with gearboxes and turboshaft engines in self- contained propulsion system pods. The el=craft uses two 1150- horsepower Pratt and Whitney PT6C-40 dlrect-drive engines. : General arrangement of the aircraft is shown in the three view in Section X.

] The three-bladed 25-foot-diameter proprotors are gimbal mounted with hub springs to increase longitudinal control power in helicopten mode. The proprotors are identical in design to the

scaleproprotortunnelto ubendertestTasked eaIIrlyof inthis1970program.in the NASA-DiscAmesloadingfull-is 9.7 pounds per square foot at the normal gross weight of 9500 I pounds and 12.6 at the maximum gross weight of 12,400 pounds.

_ _ _ At normal wefght the wing loading is 54 pounds per square foot. --'. The cockpit is arranged for a crew of two and can be flown from _: ] either seat. Conversion and power management procedures are .._,._g, l simple and straightforward and permit the aircraft to be flown _ by a single pilot. Power is controlled in the helicopter mode _ by a collective stick and twist grip throttles and in airplane a ] mode by throttle levers and a proprotor governor. Conversion is controlled by fore and aft movement of switches on the pilot

i and copilot cyclic control sticks. The canopy and forward fuselage are designed for installation of Douglas Escapac I-D ejection seats for the research flight

I tinestcommercials. The cabinseating_s largeor twelveenoughtroopsto accinommodaa high-densityte eight passseaengetingrs arrangement.

I The aircraft is designed for a 2.0 g load factor in helicopter and conversion mode and 3.5 g's in airplane mode. Design limit dive speed is 350 knots. Basic design criteria are summarized in

I dataSectionin TableB. BasIII-2ic dataand arecontrsummarizedol travels in inTableTableIllIIl.-3.-l, dimensional

i Thereporfollot. wThroing ughoutdesign thelayoteutsxt aofre thisincludedsection,in Sewcheretion referenceX of this is made to these drawings, drawing number will be shown in parentheses.

I 300-960-001 Three View 300-960-002 Proprotor and Controls 300-010-001 Blade Assembly

I 300-960-003300-010-i00 InboardProprotor ProfileAssembly- Nacelle 300-960-004 Hain Transmission %

I 300-960-005 Fixed Controls - Wing

300-099-003 I[l-i 00000001-TSC04

I

IBEI J. HF__LICOPTEI_ ,:O,'_PaNV

300-960-006 Fixed Controls - Fuselage "" 300-960-007 Wing Assembly 300-960-008 Fuselage and Empennage 300-960-009 Crew Station and Cmneral Arrangement

i TABLE III-i BASIC DATA

I Aircraft Weight

| Normal Gross Weight 9500 ib 1 Maximum_Gross Weight 12400 ib Empty Weight 6876 ib

_ _ | EngineDesigo(Two)LandiogWeight 950O

._,, ] Manufacturer Pratt and Whitney

I_ Model30-Minute Rating (2 x 1150) PT6C-402300 hp MaPoxwimerumLoaContinding uousat NormRatingal Gross(2 xWeigh995)t 4.11990 ibhp/hp Power Loading at Maximum Gross Weight 5.4 Ib/hp

Proprotor (Two) Diameter 25 ft

I NumbSolidityer of Blades Per Rotor 0.0893 Disc Loading at Normal Gross Weight 9.67 ib/sq ft | Disc Loading at Maximum Gross Weight 12.63 Ib/sq ft Wing

Span 3_.2 ft Area 176 sq ft Aspect Ratio 6.6

Wing Loading at Maximum Gross Weight 70.5 ib/sq ft 1

_ 300-099-003 III-2 00000001-TSC05

J i_ BE[_I H[_I_I(.=OI:.TER c;o_PAN Y

J TABLE III-2 f"

J DIMENSIONAL DATA Aircraft Dimensions

I Overall Length (41.8 feet) 501.0 in Overall Width (Proprotor Turning)

I Overall(57.2 feeWidtht) (Proprotors Removed) 686.0 in (36.4 feet) 436.0 in Overall Height Pylons Vertical (Top Spinner at I NGWof (15.32 feeFromt) Static GL) 185.0 in Overall Height (Top of Fin - From

I SpanStaticBetweenGL) atProprotorNOW (15.7Centerlinesfeet) 188.0 in at Conversion Pivot Points ,_ I (32.2 feet) 386.0 in ', _ Static Ground Line Reference at WL ii.0 --uW_. Height of Conversion Pivot Point -_...._ Above Static GL at NGW (7.42 feet) 89.0 in -" Conversion Axis Location, Percent | Wing MAC " 39.0

I Distance from Conversion Pivot Point To Horizontal Tail i/4-Chord of

To Vertical Tail i/4-Chord of i MAC (20,5 feet) 247.1 in MAC (19.3 feet) 231.7 in

I Distance from Wing iP_-Chord of MAC To Conversion Axis 8.7 in

MAC (21.3 feet) 255.8 in I To HorizontalVertical TallTaili/4-i/C4-Chordhord of of

I GroundMAC Clearance(20.03 feet)at NGW (GL at 240.4 in WL ll.O) 12.0 in

I Main Gear Tread Width ii0.0 in Distance from Nose-Wheel Axle to Main-Gear Axles 214.0 in I Engine S0-Minute Rating I Horsepower 1150 shp RPM (Output Shaft) 30000 in-lb , Torque 2420 in-lb

i

I , , ,,

_ 300-099-003 III-3 00000001 -T.q O.NR

! I_ RFI .!. H_I _ICOPTER (_(,MPANY

I TABLE III-2 - Continued

Maximum Continuous Rating

RPM (Output) 30000 i Horsepower 995 shp Torque 2090 in-lb

I Dry Weight 325 ib

| Drlve.System Gear Ratios | Englne to Proprotor 53.1:i Englne to Interconnect Shaft 4.63:1

_ Proprotor

| DiameNumberterof Blades per Proprotor 25.03 ft "i | Disc Area per Proprotor 491 sq ft , Blade Chord 14 in. basic blade _"7 | 17 in. cuff root at _,,i/ | 0.0875R {¥:" Tapering to 14 in.

at 0.25R I SBladeolidityArea (3 blades) 43.0.08975 sq ft Blade Airfoil Section

I TipRoot (GL Mast) NACA 6464--923508 aa== 0.30.3 Blade Twist (See Figure IIl-i for

i HubDistPreconeribution)Angle -45.0+2.5 degdeg 83 -15.0 deg Onderslinging 0 deg

I MastFlappingMomentDesignSpringClearan(perce Rotor) ±127002.0 indegib/deg Blade Flapp_ng Inertia (per Blade) I05 SLug ft 2

I BladeDirectionLock ofNumberRotation, Inboard 3.83 Tip Motion Aft/Up

I Pylon and Conversion Actuator Point of Intersection of Mast and

I ConverFSsion Axes 3OO.O WL 100.0

i ConverBLsion Axis Wing Chord Location 39.0193.0 percent MAc Conversion Axis Forward Sweep 5.5 deg Conversion Axis Dihedral (Up) 3 deg

I --. I

_ 300-099-003 III-4 OONNNNNI' .T_C'O7 ' '

I_ BELt- HELICOPTER comPANy

TABLE III-2 Contiuued ,.

J Angle of Mast Axis to Conversion Axis 95.5 deg

I AngleHelicopterof OutboardMode Tilt of Mast Axis 2.5 deg Airplane Mode 0 deg

I DistanceConversionRotorAxisFlapping Axis to 56.0 in Conversion Range (Pylon Vertical = 0) -5.0 to + 90 deg

I ActuatorExtended Length 39.39 in Retracted i0.00 in

DistTravelance Engine CL from Mast CL 29.3917.0 inin

![ Wing Span (34.2 feet) 410.0 in [ Span Between Conversion Axis _" I Pivot Points 386.0 in ! Area (Total) 176.0 sq ft _' Root Chord (BL 28.0) 62.0 in I Tip Chord (BL 205.0) 62.0 in I Mean AerodFnamio Chard Chord (BL 102.75) 62.0 in

1/4 Chord at FS 291.3 Leading Edge at FS 275.8 Airfoil Section (Constant) NACA 64A223 Modified Aspect Ratio 6.63

I DihedralForward Sweep 6.52.167degdeg Angle of Incidence 3.0 dog

I Wing Twist 0 deg Aileron

sq I ArSpanea/Side[Along(AfHinget of Line)Hinge Line)(8.04 feet) 96.510.4 in ft Chord/Wing Chord 0.25

I Flap

I ArSpanea/Side(Along(AftHingeof HingeLine)(4.25Line) feet) 5.5SI insq ft Chord/Wing Chord 0.25

I Wing Loading Normal Gross Weight 54 ib/sq ft

I Maximum Gross Weight 70.5 ib/sq ft

I 300-099-003 111-5 ..... 00000001-TSC08-

I l_rl_ B_LI_ HELICOPTER co.pAN Y

I TABLE 111-2 - Continued

I Fuselage

I LengMaximthum (38.1Breadthfeet) 457.566.0 inin Maximum Depth 74 in

I CabinComparLengthtment) (Cockpit(15.25 fePluset) Cargo 183 In Cargo Compartment Length (9.16 feet) 110 in Cargo Compartment Width

I FloorMaximumLine 48.060.0 in Cargo Compartment Height

I AheUnderad Wingof Wing 60.054.0 in Cargo Floor Space (9.16 feet x

_ "_, i Cargo4 feet)Compartment Volume (9.16 feet 36.6 sq ft _,_. 4 feet x 4.75 feet) 174 cuft

"_. 7 I Vertical Tail -. _F: Span (10.33 feet) 124.0 in Total Area 57.8 sq ft -_ I Rudder Area (Aft of Hinge) 7.6 _q ft Rudder Chord/Total Chord 0.15 : m Aspect Ratio 1.84 I Sweep of i/4 Chord 36.83 deg Root Chord at WL 75.0 97.0 in Airfoil Section NACA 64A015

I TipAirCfoilhord atSectionWL 199.0 NA37C.34A 64A015in MAC Chord (WL 127.8) 71.6 in

i MACMAC Leadingi/4-Chord Edgeat atFS FS 513.8531.7 Horizontal Tall

I Total Area 62 5 sg ft Span (16.0 feet) 192.0 in

I AAspectngle ofRatioIncidence 04.1deg Elevator Area (Aft of Hinge) 17.6 sq ft Elevator Chord/Total Chord 0.30 I Root Chord (BL O) 56.0 in Airfoil Section NACA 64A015 Tip Chord (BL 96.0) 37.75 in

I MACAirfoilChordSection(BL 44.88) NA47C.46A 64A015 MAC Leading Edge at FS 535.2

I MSACweepi/4of-Chordi/4-ChatordllnFS e 22547.7.21 deg

300-099-003 III-6 nnnn-nnh1_-F -r o

I

tABLEBELL" 111-HELIC2 OPTER- Continued(:OMpANY

Main Gear

Number Wheels per I Tire Size,0f Type and PlySideRating 8.50i x i0, Type III, lO-ply

i InflationNominal OutsidePressureDiameter 25.70 2psiin Load Rating (Helicopter) 9200 ib Flat-Tire Radius 7.0 I OleoMaxlmnStrutm GroundStrokeSpeed(Total) 80i0.0kt fn

I Nose Gear Number of Wheels 2 Wheel Spacing (Dual) 9.5 in I Tire Size, Type and Ply Rating 5.00 x 5, Type III, __:=i ."! '.. 6-ply • • _ Inflat,on Pressure '9 psi I Nominal Outside Diameter 13.9 in . Flat Tire Radius 3.B in Load Rating (Helicopter) 2100 lb

I MOleoaximumStrutGroundStrokeSpeed(Total) 89.00 ktin

l TABLE 111-3 CONTROL TRAVELS

i Cockpit Controls ,

I CyclicCycl_c SSticktick LFatoreeraland Aft _6.0 inin Collective Stick 12.0 _n

I RudderFedal AdjustmentPedals _2.2.05 in

i ProprotorCollectiveContPitrolcsh at 0,75R Helicopter -2, +18 deg

I AirplaneConversion +18,See Figure+50 deg111-7 Differential Collective Fitch

I (Lateral Cyclic Stick) Helicopter ±3.0 deg % Conversion See Figure III-8 I Air_lane ±0.3 deg

I 300-099-003 111-7 . -...

I BELL HELICOPTER _:O_r'ANY I TABLE III-3 - Continued

I L, Collective Pitch Trim

I CHelicopteonversionr _0.5_0.5 ddee E Airplane ±0.5 de E

I Cycllc Pitch Tots1 ±4.0 de E

Fore and Aft Cyclic Pitch I Hel_copter ±i0.0 deg i- Conversion See Figure 111-9 Airplane 0 | Differential Cyclic Pitch : (Rudder Pedals)

@'_ ''I_ I Helicopter (0-60 kt gAS) ±4.0 deg

(i00 kt EAS +) ±I.0 deg 4_/, (60-100 kt EAS) See Figure III-ii _' I Conversion . See Figure III-i0 _, Airplane - - 0 deg

I Fixed Surfaces Aileron

i Flaps Upat i00 degdeg +15.0+21.6 -15.0-15.2 deg Flaps down 30 deg +18.0 -12.5 deg Flaps down 60 deg ,-14.2 - 5.0 deg

Elevator Trim Tab ±f20.020.0 deg Rudder ±20.0 deg I I I I

Z00-099-003 IIl-8 , _ ..... Lh 00000001-TSC11

I 01_ BEIi. HEI_ICOPTER c;omt_aNv

I B. Design Criteria

I designedCriteria haveflightbeenresearchestablishedaircraft.to provideBasic criteriaa safe andcomplyefficientlywith the Federal Aviation Regulations. Design limits, load factors,

I andrequirementsconditions ofhavethe beenFAA, established"Tentative Airworthinessin accordance Standardswith the for Verticraft/Powered Lift Transport Category Aircraft - Part XX," dated July 1968. In the areas where this document fails to pro-

I FevidederaladequateAviationdefinition,Regulationsthe forapplicablerotorcraftrequirementsand airplaneofs the_re used as a guide. In the areas not covered by any of the

I gregulationsranted, Bellanddesign/or wherepracticeexceptionsfor helicoptershave been hascustomarilybeen used. The basic design criteria and design parameters for the Model

i 300 are given in Table 111-4.TABLE 111-4 BASIC DESIGN CRITERIA I "_i"' Normal Gross Weight 9500 ib m Maximum Gross Weight 12400 lb I Empty Weight 6876 ib

_ ' " DesignHelicopterOperating Speed, EAS 140 kt I Conversion 140-170 kt Airplane 260 kt , Design Limit Speed, EAS I Helicopter 156 kt Conversion 189 kt Airplane 350 kt I Proprotor Maximum Operating Speed and RPM Tip Speed

I Helicopter (fps)740 (rpm)565 Conversion 700 534

I LimitAirplaneLoad Factors at 9500 Pounds 600 458 Helicopter 2.0

I AirplaneConversion 3.52.0 Transmission Design Power

I Helicopter 1060 hp Airplane 860 hp Conversion lO00 hp

I Single Engine 1150 hp I

300-099-003 III-9 00000001-TSC12

I (_ RFI.I. H[]IJCOPTER (:OMPANV I C. Proprotor ,.

I The 25-foot-diameter proprotor which is currently being fabri- cated for full-scale wild-tunnel testing, is designed to f]_ht- worthy standards and is appropriately sized for use on the Mode]

I for300 aircraft.efficient crAerodynamicuise in the design200- toparameters300-knot speedhave beenrange.selectedThe same requirements for reliability, service life &nd maintenance

i asthisan proprotooperationalr. Basedhelicopteron wind-tunnelwere met testin theresults,detail designnecessaryof design changes would be made prior to fabricating the prop- rotors for the research aircraft. The proprotor blades and

I hub are described below. i. Blades

The blades (300-010-001) use type 17-7PH stainless steel as the basic blade material as a result of a design study in which the

-- _. '_ stainless steel were considered. Results indicate a substantial relative merits of aluminum, titanium and several types of :"_W'_'_ weight savings for both steel and titanium compared with --'_. .f aluminum blade designs. The 17-7PH steel blade provided the •_ I desired natural frequencies and strength for minimum weight. q Thickness, taper,, twist and camber distributions were selected

l tforo meethelicopterthe varyingand airplanestructuralflight.and aerodyNASAnamic64-seriesrequiremeairfoilsnts are used with a 64-208 at the tip and a 64-935 at the theoretical ' root (blade Station 0). The thick blade root section is required high I to provide adequate blade strength when the blade is at pitch in airplane flight where torque and inplane gust loading cause high bending moments about the airfoil chord line.

The basic chord of the blade is Ig inches. Chord, twist, lift coefficient and thickness distributions are shown in Figure III-i. Blade stiffness and mass distribution are shown in Figure 111-2.

I 2. Hub

I spindlesThe hub (300-010-i00)and a universalconsistsjoint ofassemblya titaniumthat yokeis spllnedwith threeto the mast. A nonrotating, elastometric hub-moment spring is attached

I sptoringthe yokeis attachedthrough toa bearing.the transmissioThe nlowercaseendby ofstutheds. hub-moment

i Thebearingsunivermountedsal jointin assemblyaluminum consistspillow blocofks a steelon two crossopposingwith spindles and a steel fo_k with bearings on the other two spindles. These four roller bearings are not provided with

I steelinner crarossces, member.but roll A oncommonthe ease-hardenedoil reservoir _ournalsis createdof bythe oil passages drilled within the cross member. Oil level slght gages

' I are in_;talled on the pillow block housings. The bearing housings

I 300-099-003 III-I0 00000001-TSC13

I conta_.n thrust bearings to enrry the proprotor ll-forces nqd seals to retain the oil. "

l The iaboard an_ outboard pitch change roller hear_ngs assemble in the bladels integral root fitting. The inner race oi" these bearings assemble on the spindles of the yoke. A stainless liner is the I steel bonded to spindle to prevent fretting bet_,en the inner race and the titanium spindle. The pitch-change bearings are oil lubricated from a reservoir located in the I pitch horn. The three wire-wound blade retention straps have an integral

I steelthe yoke.fittingThe whichoutboardseatsfitting,at the inboardof the retentionend of eachstrap,spindleis of attached to the blade by a steel bolt through the blade root fitting, spar and doublers.

J D. Drive System

_ The drive system consists of a main transmission assembly _ (300-960-004) at each wingtip, a system of drive shafting ... through the wings connecting the two main transmissions, and

a center gearbox mounted inside the fuselage (300-960-007). case. Each transmission is attached to a steel spindle which . Theis supportedPT60-40 engineby the attachestwo outboard'directlywing toribs.the transmissionHydraulically- pylon powered and mechanically-interconnected Acme screw actuators ,i, support, power, and control the conversion of the pylon assembly about the transmission-spindle axis. In the airplane mode the actuators drive the pylon into a down stop supported by the tip rib.

In normal operation, each transmission delivers power to its I proprotor from its own engine. The interconnecting shafts in the wings operate unloaded, except during mat_euvers, single- engine operation, or asymmetrical loading conditions, where the interconnect driveshaft distributes power as required.. Design power for the transmission is shown in Table 111-4 and is

I basedengineson operathe tsameing. designTo permitorquet theof useeachof modemaximumof flightpower withfrom boththe remaining engine in the event of an engine failure, the engine output shafting and herringbone gear stage are designed for the |"m maximum engine output power of 1150 horsepower. Several factors are multiplied by the design power and torque to arrive at limit and ultimate torques for the various stages. A distribution

I whichfactor allowsof I.i0forisanapplieduneven todistributionobtain the maximumof power steadybetweenpowerthe proprotors. A transient torque factor of 1.67 is then applied

i totorqobtainue is limit1.5 timetorqs ulimite duringtorquasymmetrice. maneuvers. Ultimate The main transmission assembly supports all pylon components. I The structural parts of the assembly consists of a spindle,

I 300-099-003 III-II 00000001-T$C14

I (_ HELl H_LICOPTER _x_ml-an_

p_.on ease, ihtermediate ease, and a top (mnst) e_se. Tile oll_ine _. _,Id pylon eow].ings are also s,pported by the trallsmlss_on.

Power is transmitted from the engine by au adapter shaft _ch picks up the female spline of the PT6C-4O power turbine shaft, then through a combination power and torquemeter shaft which _s sp[ined to a herringbone pinion. The herringbone stage of reduction gears transmits the power through a one-way clutch to the two planetary reduction units. Power is supplied to the rotor masts by the planet carrier of the upper planetary stage.

The interconnect power train, linked to the main proprotor drive side of the one-way clutch, consists of a spur gear set, an intermediate sha£t (with torquemeter shaft), and a spiral bevel gear set.

The accessory gears provided for:

Hydraulic pump

Transmission oil pump

I! - AC generator i - NII governor

The center gearbox, with splash lubricated bevel gears, is mounted on the rear spar of the wing at the centerline of the fuselage. This gearbox accommodates the change in interconnect _" shaft angle due to wing sweep. A magnetic sensor is installed to provide a signal to the rotor tach indicator and proprotor governor.

! E. Powerplant

The Model 300 is powered by the Pratt and Whitney PT6C-qO free turbine turboshaft engine. This engine is an advanced version of the PT6 turboprop engine widely used in executive, third- level airline and utility turboprop aircraft. The PT6C-40 is a direct drive engine with an output speed of _0,000 rpm. : The turboprop gearbox is removed and the lubrication system ' modified for vertical operation. The engine has takeoff and 30-minute ratings of 1150 horsepower and a maximum continuous power of 995 horsepower. The engine is rigidly mounted to the ! proprotor transmission case. Engine torque is read by a Simmonds magnetic torquemeter located on the main transmission.

I 2. Induction System The engine induction system (300-960-003) is designed to give

, maximum total pressure at the engine inlet screen, to provide

i 300-099-003 III-12

- . : - := .... _- .... r_ L = =_. 00000001 -T.qF_n I

I /_11_ BELl HEIJcOPTER _x_mPanY

I anti-iclng, and to protect the engine from dust and sand ingestion. Air enters through the nacelle i_llet and diffuser °"

I effectiveduct. A 90-inertialdegree particleturn into separathe engintor eto plenumremove providesdust and ansand. A screen is provided which ices over during icing conditions

I to increase the separator efficiency. Moisture and debris are removed from the primary engine air in the by-pass duct and ejector. By-pass air spills out of the I by-pass duct to a plenum, then moves across the engine and - transmission oil coolers to the engine accessory section. Air passes over the accessories from the coolers, through a bell- -: • mouth and blower to provide a power source for the ejector. " I The ejector then emits the by-pass air, induction air contam- ination, and heat from oil coolers and engine accessory section. I " I 3 OilSystem i: The engine is supplied with oil from a 2.3-gallon tank that is _ I an integral part of the compressor inlet case on the engne.4 "-_ • Oil flows from the tank to the accessory reduction gears, engine • bearings and filter. Scavenge oll is directed through the oll =_"- / I cooler located behind (airplane mode) or below (helicopter mode) thermostatically regulated bypass to prevent high surge pressures _ theduringaccessorystarts gearunder case.cold weTheatheroil conditions.cooler is equippedAir for withthe acoolers is driven I provided by a mechanically blower. A shaft from the accessory drive pad on the engine provides power for the blower. The oil leaves the coolers and is returned to the tank forming

I aindi"coldcationtankof system".system operationThe tank isis providedvented overboard.by oil temperatureContinuous and pressure instruments in the cockpit. Warning lights are

I alsoture. provided to indicate low oil pressure and high oil tempera-

4. Fuel S_stem

I Fuel is supplied by two separate systems, one for each engine. Each system is composed of two cells interconnected to form a

I singlepounds. tankThe incellseach arewing,constrwithucteda totalof a fuelflexiblecapacityrip-resistanof 1600t material. Continuous support for each cell is provided by the

I accomplishedstructural honeycombthrough panelsfiller ofcapstheinwing.each ofGravithetyinboardrefuel_ngcells._s One dc fuel-booster pump is provided at each inboard cell.

I checkEngine valve,fuel passesfuel filter,from thefirewallbooscer-pumpshutoffdischargeand conversion-swivelthrough a fitting before entering the fuel control. A pressure ga_e {n

I pump.the cockpitAn interconnectindicates thebetweendischargethe twopressdischargeure of thel_nesboosterperm{ts ene pump to supply both engines if a pump fails. Opening the "' • tank interconnect valve will allow inter-tank gravity transfer I of fuel to the operative pump.

i 300-099-003 III-13 O000000J-TSDOP

i I I

I F. AirI!rame

The following specific objectives were established for the wing

I design (300-960-007). Place the elastic axis far forward to minimize the torsional deflections resulting from coupled proprotor/

I pylon/wing motions. - Provide high torsional stiffness without undue weight

I penalty. - Provide the maxit0um possible flap and aileron area,

to minimize the projected wing area and hence the l _ and design them to deflect to a large angle, in order • aircraft download, during hover.

Provide an unobstructed passageway to route controls ! and the transmission interconnect.

I Provide fuel space

Thesestructurobjectivesal box andaresweepingaccomplishedthe wingby forwaa forward rd 6.5-degreeslocation ofto the relationship wing center I obtain the desired between the of pressure and the conversion axis. Fuel cells are located inside the structural box; the controls and the transmission

I interconnect shaft are located aft of the rear spar. A lightweight, torsionally efficient structure Ks obtained by

I Ausingll of sandwichthe skin constis reffeuctionctive foinr theboth skinsbendingof andthe sttorsion,ructural wherebox.as with a plate-stringer combination the skins may be in a buckled state under load, and the stringers provide no torsional stiff-

1 hess. ==J A significant factor in obtaining high torsional rigidity is the

,., highcross-sewing-ctionalthickness area. ratioSincewhichthe prorigidityvides a varielarges wwithing_boxthe sqt,are of the area, high torsional stiffness results. In addition to

concontribtributingtes toto thethe wingtorsionalbending rigiditstiffness.y, the highThis thicknein turnss resraultstio in the low structural weight required to carry the high bending moments resulting from the lift being concentrated at the wing tips during vertical fl_ght. The aluminum alloy front and rear spars are designed partially

I ments.by stiffnessThe frrequirementsont spar is anda shear-resistantpartially by stweb.ructuralThe rearrequire-spar Ks honeycomb sandwich. The outboard two ribs of the wing support the conversion spindle and the tip rib supports the conversion- actuator pylon Bto_. l spindle and down The intermediate ribs _orm

I 300-099-003 III.-14 NNNNNNNt _T_nn_

I BFL=I HELICOPTER I:()MI_ANY

I t:he bulkheads for the extremities of the fuel eel.l.s _nd provide for. redistribution of loads from the ailerons and flap hinge r I ribs. The leading-edge structure is honeycomb sandwich, hinged for

I access to the conversion Interconnect shaft. El's and GJ distribution and panel-point weights for the wing

are shown in F_gure III-3.

2. Fuselage

: I The.fseml-monocoqueuselage (300-960-008)structure of _s2024a conventionaland 7075 aluminumnonpressurized,alloy. The four main longerons run continuously above and below the

_; I cbreakutoutsup reqtheuiredskin forpanelsdoorstoandthe therequliredandingsi=e.gear. MajorStringebulk-rs heads are provided for the ejection seat rails at both sides of m the entrance door,_ front and rear wing spars, at both ends of

I Nosethe landinglanding gegearar bay,supportand beams,for the atverticalBL 7.75 stabilizerextend betweenspars. Stations 131 and 237.

The cabin extends between the canted bulkheads at Stations 219.8 and 3_7. The inside cross-sectional dimensions are 60-

I inchesthe wing).wide, The127-inchesentrance longdoor andopening,60-incheslocatedhigh at(54theinchesforwardunder left side, is 28-inches wide and 52-inches high. Emergency exits are provided on each side of the cockpit and in the cabin

I aluminumon the righthoneycombside betweensandwich Stationswith a rigidized319 and 347upper. Thesurface.floor is an

I EI'sfuselageand areGJ distributionshown in Figureand panel111-4. point weights for the

i 3. Empennage Three spars of the vertical stabilizer (300-960-008) attach to canted bulkheads of the fuselage. The horizontal stabilizer is

I stabililocatedzerat andapproaximattaehtelyas theo thelowerfrontthirdand spancenterof spar.the verticalRibs and chordwise stiffeners to break up the skid panel are located

I between the three spars. The structural box of the horizontal stabilizer (300-960-008) is a single-call configuration consisting of two spars antl

I hesuarfaceds arecovperoringsvided ofat honeycoelevatmorb hingesandwichpointsconstranduction.at the inBulk-ter- section with the fin.

I El's, GJ, and mass dist_ibutlon for the vertical and horizontal tail are shown in Figures 111-5 and 1116. !

I 300-099-003 III- 15 000000nl-T.qmnA

I I

q. Landin_ Gear

I wingA fuselageconfiguration.mounted mainThe geargear wasretractschosen intobecausethe sidesof the ofhigh-the fuselage. Flush doors are provided between the bulkheads at

i Stationthe gear 34to7 andclear410.the Thelowergearlongeronsgeometry whenwas itdevelopedretracts. to Aperml:dual- wheel nose gear retracts into the compartment between Stations 131 and 169. Shock-absorption system is a conventional air-oil

I oleo. G. Aircraft Systems

I Thei. conversionConversion Systemsystem provides controlled rotation of the pro- pulsion pod from the vertical to the horizontal position and

I• return.any intermediaIt cante safelyposition.lock theThe pylonsystem inalseithero serveextreme,s as a or in reference for the control system by providing a phasing control motion as a function of flight regime. The conversion actuator

J'-_.... I holdsmode. the pylon against the down stop on the tip rib in airplane ._ Conversion is controlled by a switch on the control stick grips. __t I Forward movement of the switch rotates the pylons forward from • _ helicopter to airplane flight position, and rearward switch movement returns them to the helicopter position. The conver-

I Theslon normalof the conversionpylons may timebe stoppedis approximatelyor reversed nineat seconds.any position.

I orShouldelectricalone conversionfailures, actuatorthat unitfailis todrivenfunctionby thedue actuatorto hydraulic motor on the opposite wingtip through the mechanical inter- connect shaft. In the event of a complete dc power failure, a

I mechanicalreconversion backT-handleup system,locatedoperatedin theby cockpipullingt,_positionsthe emergencythe hydraulic valves to cause the actuators to move the pylons to

I the helicopter position. The major components (300-960-007) of the conversion system 'I include the double screw conversion actuators with hydraulic I motors and electrically powered servo-valves, the interconnect shafting and a control phasing gearbox located on the forward side of the front spar of the wlng. The hydraulic motors that

I activated,power the conversionthree-positionactuatorsservo-valvesare control].edwhich receiveby dual,feedbackpilot- information through a small gearbox on the interconnect shaft

near each wingtip. The control phasing gearbox, provides a linear output, propor- tional to the pylon an_le, that phases the various fixed controls during conversion. Thls unlt also provides the signal for ,be • pylon-conversion-angle indicator in the cockpit. A convers_,_n brake assembly is incorporated in the phasing gearbox which

300-099-003 111-16 ONNNNNNI_T I- n

I

locks the pylon when the aircraft is on the ground with hyd _u].ic ..

i power off. 2. Hydraulic System

I systemThe hydraulicutilizingsystemMIL-H-5606is a MIL-|I-5440fluid at anTypeoperatingII (-65°Fpressureto +275'F)of 1500 psi. It has two independent transmission-driven hydraulic

I pumps.operate Sincewheneverthe thepumpsproprotorsare drivenare byturthening,transmission,and they arethey independent of engine power.

connected, to side of the dual I The primary hydraulic system, one flight-control actuators, is powered by the hydraulic pump in the left pylon. The utility system is powered by the pump in

I the urightilityt pylon.system Afteris separretaraction,ted from thethe landiugflight-congeartrolportporiontionof by an isolation valve. The system powered by the right trans- '_ _ mission pump, isolated from any utility function, then becomes I a second primary system for one side of the dual flight-control actuators. Dual or single power is provided to the hydraull- "_' rally operated components as shown below:

L_I_I I Primary Utility (Left Pump) (Right Pump) Function

I x x Cyclic --_ x x Collective ii I X x Ailerons × x Proprotor Governor

I X ConverLeft sion Actuator x Conversion Actuator

I x x SeASRight

i x LProprotoranding geaTrimr

I 3. Electrical System

I Thestarterelectricalgenerators,systemandcontwosists13-ampere-hourof two 200-ampere,batteries,28-voltprovidingdc primary de power. Two 250-va, l15/200-volt, slngle-phase 400-

I Hertzbusses acareinverte_sconnected providein parallelthe ac thropower.ugh aTwobus-tieessentialrelay.de Each 200-ampere starter generator supplies power to one of the

I essential de busses. Each essential dc bus has a 250-va inverter

I 300-099-003 III-17 ...... _!- . ___ 2 • _ ,...... 00000001-TSIDO6

J , ii

I i(i_ BELL HELICOPTER c:oMl_Ny

I with its essential ac bus. '['here are two essential ac and de busses, and one nonessential dc bus. _

I The electrical system is designed to provide complete dual ac and dc power sources. These sources and their essential and

I nonessentialsources and theirbussesbusseare8 designedin the eventfor completeof any failisolationure. of the

i II. Aircraft Contro!s i. Proprotor Controls

I collecProprotiveor cohentrolsad assembly(300-960-002)above theconsistproprotorof aandrise-and-falla monocyclic (fore and aft) swashplate below the proprotor. i I The collective head is attached to the proprotor mast. A non- rotating tube, extending inside the mast to the collective boost cylinder, gives vertical motion to the rotating collective head.

i Athecollectivecollective leverhead.is Aattachedcontrol totubeeachextenof dsthe frthreeom onetrunnionsend of of " each collective lever to a pitch horn. At the other end of each

I collective lever a tube goes to the rotating swashplate. The rotating swashplate (outer) is driven by the lower ring of

i theto theproprotortop casespinner.of the transmissionThe nonrotatingand swashplateis free to tiltis attachedabout only one axis. The cyclic cylinder is attached to the non- rotating swashplate (300-960-003).

I Collective control inputs, which increase or decrease the pitch of all blades at the same time, are introduced by means of a

I casetandembelowhydraulicthe mastcylinder(300-960-003).which is attachedThe servo-valveto the trlinkageansmissionof the collective cylinder receives its input from the pilot through a swivel joint, on the conversion axis, which connects to the fixed controls I in the wlng (300-960-007). The input motion is introduced along the conversion axis so that the collective system functions in the same way in both airplane and helicopter

I modespitch. of operation, though with different ranges of collective

I Theone-per-revcyclic controlvariationscylinderin bladetiltspitch.the swashplate,The servo whichvalve causesof the cyclic cylinder is actuated by the pilot through a linkage (300-960-003) which is automatically phased out as the pylon I converts from vertical to horizontal. This phase out is accomplished by having the fixed controls in the wing (300-960- 007) impart vertical motion to the end of the cyclic input tube

I ducedthat iswhenlocatedthe inponutthetubeconversionis verticalaxis.(helicopterAxial motionmode).in intro-No axial motion occurs when the tube is horizontal (airplane mode) ! 300-099-003 III-18 I I BELL.EL,COPTeR The design of the control linkage boost cylinder permits the pilot to control the proprotors manually in the event of a double hydraulic system failure.

2. Flight Control

The flight-control system combin_s the basic elements of con- ventional helicopter and airplane control systems. The cockpit controls for the proprotors and control surfaces are arranged so that a single pilot can maintain full control in all flight regimes, including conversion. Each of the two crew stations (300-960-009) has complete controls for pitch, roll, yaw, and thrust in all modes of flight. They consist of control sticks, rudder pedals, and collective levers, for both the pilot and copilot; a single set of power-management controls and a flap control on the center pedestal, and a rotor trim control on each cyclic stick. Dual-twist grips on the collective levers are interconnected with the power-management controls on the center : i i pedestal. -_ i_ In helicopter mode, the controls apply blade-pitch changes to " _ produce powerful control moments and forces. Fore-and-aft cyclic .._i pitch provides longitudinal control, while differential-cyclic "'_ pitch produces directional control. Collective pitch is used for vertical flight and differential-collective pitch controls roll.

In airplane mode, the controls actuate conventional control surfaces which provide the control response characteristics of a conventional airplane. These control surfaces are also actuated in the helicopter flight mode, but they have minimal effectiveness because of the low dynamic pressures and high control moment capability of the proprotor8.

Conversion or reconversion can be made within a wide range of variables such as airspeed, conversion angle, and fuselage attitude. Mechanical phasing of the proprotor control authority minimizes the need for control inputs during conversion. To provide the proper control authority during conversion (or reconversion), some controls are phased out, others are phased

changesin, and inthe controlsauthority as oftheotherspylon isis alteconverred.ted Thefromautomhelicopteratic mode (-5 degrees to +15 degrees conversion angle) to airplane

modethrough(+90lll-ll.degrees conv_rsion angle) is shown on Figures III-7

3. Stability and_Control Augmentation Syste m

A stabilization system is used to enhance the flying qualities in helicopter and conversion modes. The three-axis stability and control augmentation system (SCAS), uses rate gyros to sense pitch, roll and yaw. Electrical signals from pilots control motions as well as signals from the rate gyros are input into appropriate shaping networks of the SCAS. The result is an air- craft that is stable and well damped for external turbulences,

l 300-099-003 III-19 • , .w. ONNNNNNI_T nn

yet is highly responsive to pilot control inputs. As a normal .. proeedure, the pilot will engage the SCAS prior to takeoff and fly the aircraft in the helicopter and conversion modes where SCAS is phased out as the helicopter controls are phased out.

i The SCAS actuates the rotor controls through servo actuators located on the collective and cyclic hydraulic cylinders. Redundancy is provided by dual actuators which are operated I individually by the two hydraulic systems. 4. Proprotor Governor System

i The proprotor governor system is used to simplify power manage- ment and rpm control and to prevent engine-power adjustments and external disturbances from changing proprotor rpm. The system is a closed-loop control system that maintains a pilot- selected proprotor rpm by controlling collective blade-pitch in the airplane mode. i '-II_ _ The proDrotor governor system detects any error between the command rpm and the actual proprotor rpm. This error signal is am pll_ted and used to drive a hydraulic actuator in the collec- / ' tire control system. With a constant proprotor rpm setting, • _ '_ increasing power with the power management levers will increase _ the collective blade-pitch to hold a constant rpm. This will result in increased aircraft velocity without changing proprotor rpm. Decreasing power will decrease the collective blade pitch and reduce aircraft velocity.

The proprotor governor is a fail-operate type system. Sufficient redundancy and monitoring circuitry is included so that a single failure will not result in loss of the proprotor governor. If a failure occurs, a warning light will be illuminated. If a second failure occurs, the proprotor governor system will auto- matically shut off and the pilot will control proprotor rpm manually with the collective lever.

5. Power Management

Power management is simple and is designed for straightforward cockpit procedures. Power control is provided by two control systems. For helicopter flight, the engine power-turbine governors maintain selected proprotor rpm by increasing or decreasing power as manual changes are made in collective pitch. In airplane flight, the proprotor-pitch governor maintains selected rpm by increasing or decreasing collective pitch as manual power changes are made. Thus, the Model 300 may be flown in helicopter mode in the same manner as a conventional he]i-

copter,turboprop andairplane.in airplane mode in the same manner as a conventional

, Throttle and proprotor governor rpm select levers are mounted on to the pedestal (300-960-009), convenient both the pilot and

I 300-099-003 III-20 00000001-TSD09

I I_ NLLL ._LJCoP'rr_R ,:,_MrANY I copilot, The pilot and copilot collective st_eks have dual twist grip throttles and turbine governor rpm beeper switches. _'

Conventional helicopter rpm droop compensation is provided for each engine. The collective stick is connected through the droop-compensator linkage to a droop cam on each engine, The droop cams position the load-signal shafts on the engine fuel control, which in turn schedule limited rpm changes to compensate for the engine droop characteristics.

Engine output power for both helicopter and airplane flight is regulated by the fuel control on each engine. Movement of the power-control shaft on_each engine controls fuel flo_. This shaft is positioned by the th=ottle levers. The throttle levers are controlled remotely by the twist grips during helicopter flight.

,iI, I l 300-099-003 IJI-21 m O00000N 1-T.q N i n

BELl_ HF::I_ICOI::'TER c;,r:,MI'ANY

I 1..0 _ ___J -] .- o. ,,,,,_-- DESIGN C L

I °

, ,_; _' AERO TWIST

-,.i,: I 4

m_ m : ...7

o _,

I 0 0.2 0.4 r/R 0.6 0.8 I .0 I I Figure III-l. 25-Foot Proprotor Parameters. '!

I 300-099-003 111-22

. __ = 000000()1-TSD11

I_ BELl HEI-ICC)PTER [".()/_4PANY

3.0 r_

i00

I O0 0.2 0.4 0.6 0.8 1.0 SPANWISE LOCATION - r/R

Figure III-2. Proprotor Stiffness and Mass O£stributfon.

300-099-003 IIT-23 {_/%/%t-tt%t%r% 4 -rt-,r% A ,_

;00

i0 150 .-,

O

8 200

,, 4 100 I

' ® 2 50 I

I O0 gO 80 120 160 200 0 WING STATION - INCHES ! Figure III-3. Wing Stiffness and Panel I Point Weight. !

I 300-099-003 III-24 00000001-TSD13

I

I _ F::IF_=EL H f_: I .|C_oF_,l.r_R c:e JMF,AN, f I ®

I 35 ¢) ..... - PANEl, POINT WEIGIIT

I 30 .Ely

I / _-EI, z 1 I,' \ ,"/ N

,._ N Q

i 15

t_ t_ '1/ 0 \ / /i \x\ i s

0 i

I I00 200 300 gO0 500 6q0 FUSELAGE STATION - INCHES FUSELAGE PANEL POINT WEIGHT - i G

I Figure lll-g. Fuselage Stiffness and Panel Polnt Weight. I

I 300-099-003 III-25 00000001-TSD14

BFI .L ulgl .ICOP'T'HR (:()Mr'ANY

W - SIANWI) SE"7 WEIGIFI' DISTRIBUTION - LI}/IN 0.5 0.9 1.3 1.7 2.1 2.';

140 _ TIP

w/ I ,

)i :% I

60

40 GJ ' 1 \\ ; I _o , o0 ,i20 \40 560d 80 i00

I Ely GJ - LB-IN 2 X 10 8

i Elx m LB-IN 2 X i07 Figure III-5. Vertical Tail Stiffness and Weight Distribution. I I 30n-099-003 I II-26 --_ ,, i¸ _ _=" - • O000(l(lI{_T£_ntl

i

I i_ BELl. HI_.LICC)PTER (:c)m_,aNY I

I 24 .4C 'I [i "

_ J _,

I _ _ _ _ _.o_ I "_'_ I TIP I 4j _o4 I

I °o 2o 4oBL - INCIIE_oS 8o _oo° I

I Figure III-6. WeightllorizontalDistribution.'rail Stiffness and I

I 300-099-003 III-27 0

-lO I -lO 0 i0 20 30 40 50 60 70 80 90 PYLON CONVERSION ANGLE - DEGREES I

I Figure III-7. ColLective Pitch Versus Conversion Angle. I

I 300-099-003 [II-28 NNNNNNNI.T_n'_

,i. . i

J i_ BELL, H E Lie:() i: ,TER, :c)l,_ll,AN,,, I ...... l..... I I ! 2 " I

o 3 _'_'_' | _

o

gt ,., 2

0

< H Z

_ -IO 0 i0 20 30 40 50 60 70 80 90 PYLON GONVERSION ANGLE -- DEGREES ! I Figure III-8. Differential Collective Pitch

i Versus Conversion Angle.

I 300-099-003 [I 1-29 00000001-TgF04

I

I I I I I _--F-'-- I...... ,it _- ! t...... t AIRPLANE MODE _- ! ! | lo 9 ! 8

_ 7

> ? : m , 6

"+l "-" 5

_ g

_ 3

0

I -i0 0 l0 20 30 40 50 60 7(1 80 90 PYLON CONVERSION ANGLE DEGREES | Figure IIi-9. Fore and Aft Cyclic Pitch

i Versus Conversion Angle. !

I 300-099-003 III-30 ONNNNNNt_T_n_

I I ! . It' ! ! i

M _g

I

_a

O N I °_2

Z

b.4 ! 0

i -I0 0 I0 20 30 gO 50 60 70 80 9() PYLON CONVERSION ANGLE - DEGREES I Figure lll-lO, l)II:ferent_alCyclic P][t¢:h

I Versus Conversion Angle. I ,

I 300°099-003 [I 1-31 i. .... O0000001-T£FnR

i i

I i(_ BFI i HFI.IC:()PTF]R,:o._m.A._ _ I • I

I 6 '

i _ \ +__2 5 IN PEDAL TRAVEl.

M

o_, 0 oJ _ /_--CONVERSION ANGLE S--% _ ii i'--'__'-_ -- + 5"7° " f _ so°_ - __ I _ / I " /

i _ 600

3 45 ° _ _ _ iii_ 75 ° I _ _ ,, I 90 ° \ i 00 20 40 60 80 I00 120 140 EQUIVALENT AIRSPEED - KNOTS I Figure III-ll. Differential Cyclic Pitch

I Versus A_'rspeed. [

I 300-099-003 I_ 1-32 00000001-TSE07

I _ BI_LI. HF_I .ICOI:_TER ,'_t'ANY

IV. WI,:I (;I IT ANAI,YSIS .,.

I Weights of the Model 300 proprotor were determined from detz _| I fabrlca:l'|'O n drawings. Weights for other major structural components such as wing, fuselage, empennage and landing gear were determined from layout drawings and statistical curves

of of other aircraft. 'rr_n s - I based on the weights components mission weights were estimated from layouts by use of empirical formula and by comparison with similar eomponents in present use.

I wereWeightsdeterminedof items fromsuch manufacturersas engine, instrspeclf.,catlons.uments, and electronics,

I Layoutsin SectionusedX. forTheseestimatingdrawings andarecalculatingreferred toweightsin the areapplicableincluded group weight analysis discussion that follows. These discussions also detail the methods used to determine the weights.

I A group weight statement, Table IV-l, shows the weight empty is 6876 pounds. Based on this weight, four mission weights have

I been established and are shown in Table IV-2. The research _-_':_ mission is shown for the normal gross weight of 9500 pounds and __ _ makes provisions for instrumentation, ejection seats, pilot and • , I copilot, and 1600 pounds of fuel. For the civil mission, the _;'_ _ interior is furnished to accommodate eight passengers and 240

givepoundsa grossof baggage.weight Theof 10,300fuel loadpounds,is adjustedthereby allowito 1028ng pohoverunds to I out of ground effect at 4000 feet, 95°F. The military mission gross weight is ii,200 pounds. This weight provides for a crew of three with three additional men being hoisted aboard, at the

I mid4000 poifeet,nt of95°F.the mission,The ferrywhilemissionhoveringis shownout ofat groundthe maximumeffect at gross weight of 12,400 pounds. A discussion of the performance

i for each mission is containedTABLEinIV-ISection V, Performance.

I GROUP WEIGHT STATEMENT Rotor Group 910

I Blade Assembly 580 : Hub Assembly 270 Spinner 60

IN, Wing Oroup 700 _- Tail Group 2q5 Horizontal 'Fail 136 I Vertical Tail 109 Body Group 1055 I Alighting Gear 350

I ! 300-099-(303 IV-I ! ONNNNNNI_T n

I I_ BEI I HEI.I(;O[_TER _:c_I,ANY

I TABLE IV-I Continued

I Flight Controls Group 532

I CockpitRotor, NonrotBtingControls 23653 Rotor, Rotating l_l Fixed Wing 72 I Engine Section 167 Engine Mount 16

I FirewallCowl 5992 Propulsion Group 2086

I Engine Installation 72g Conversion System 126 : Air Induction System ]8 i' (_ I Exhaust System 16 --._r_,_. _ Lubrication System 32 :.._._ Fuel System 91 -'__ . ' • Engine Controls 59 Starting Syste_l 59

DriveRotor GSystemovernor 21 -, _ Gear Boxes 761 Transmission Drive 74 Rotor Drive 45

I Instrument Group ii0 Ilydraulic and Pneumatic Group 78

I Electr_cal Group 283 Electronics Group 84

I FurnishingsPersonnel Accommodationsand Equipment 74 192 Niseellaneous Equipment and

I EmergencyFurn{sh[ngsEquipment 645g Air Conditioning Equipment 58

I Undrainable Oil 12 Unusable Fuel 14

I WEIGIIT EMPTY, POI_DS 6876

| ! I 300-099-003 IV-2 O0000001-TSE09

q

i ABL,_ IV-2

i blODEL 300 MISSION WEIGIITS Research Civil Military Ferry

i Load Condition Mission Mission M_ssion H_ss{on Crew 400 340 600 400

I FPassengersuel (8 at 170) 1600 10281360 1600 1600 Auxiliary Fuel 1059 3133 I Engine Oil 35 35 35 35 Mission Equipment

m ArmorAvionics 171 300361 Oxygen Installation 110 _, I Rescue Equipment 250 - I terior : - '; • Eight-Place Commerieai 250 | Ejection Seat Increase i14 _\: Third Man Crew Seat 37 7"4m_

I AuxiliaryInstrumentationFuel Tank 475 82 246 Baggage 240

I Total Useful Load 2624 3424 4324 5524

I Weight Empty 6876 6876 6876 6876 Gross Weight 9500 i0300 11200 12400

I A. Rotor Group The Model 300 lift and thrust system consists of two three-bladed, semi-rlgid proprotors, gimbal mounted with a hub spring which I provides flapping restraint. The Rotor Group weight was derived from detail computations of

I isfabricationshown on drawingsBell drawingand 300-010-100is 910 pounds.in SectionThe proprotorX. assembly B. wing Group

I The wing of the Model 300 is basically a two-spar, single-cell structure utilizing bonded aluminum honeycomb sandwich con- struction for Lhe upper and lower panel of the main structural I box as shown on Drawing 300-960-007. i I 300 -099 -003 !V-3 00000001-TSE10

I (_ BFI L H[_I.I_()f"|I_R .... mpaNY I W_ng primary structure _ight was based on pane[ and spar Ihiek~ _. hess requirements est:ablished by stress analysis• _ight o1 I additional structure was estimated from layout drawings. Figure IV-]. presents a statistical wing weight estimation curve

I Thisdevelopedcurve fromincludesavailablean adjustmentdata on utilityto compensateand cargo-typefor the aircraft.torsional and dynamic effects of supporting the pylon by a spindle assembly

i estimatesat the wingfortipthes. BellThis Modeladjustment266 tilewas proprotor,based on dethetailedcomponentsweight of which were sized by a detailed stress analysis. Model 300 wing weight taken from this adjusted curve is 598 pounds. An

I additionalsection with102nontaperedpounds wasskins•includedThe becautotalse estimatedof the constantwing weightwing : is 700 pounds.

• ! C. Tail Group

I The Model 300 tail assemblies consist of conventional stabilizer-

I_ elevatorassemblies andwerefin-rudderestimated configfromurations.layouts. WeightsThe unit forweightsthe tail _" _ obtained were consistent with those for similar designs operating

"'_ I in comparable flight regimes as shown in 'Fable IV-3.

TAIL SURFACE UNIT WEICHTS

I Model Horizontal(ib/@q ft)Tail Vertical(ib/sq ft).Tail XV-5A 1.81 2.13

I X0-142 2•31 2.21 AO-I 2.23 1•86

I 262 2.44 2.44 266 Bell 2.68 2.28

I 300 Bell 2.18 1.88 l:

I D• Body Oroup

• 1

i Thestructurefuselageshownof theon DrawingModel 300300-960-008.is a nonpressuri_edThe basic seml-mofuselagenoeoqte weight was estimated from Bell-developed equations with penalties added for the flooring, doors, windows, and windshields.

I The windshield weight was based on 0.50-inch thick stretched acrylic, and the same material with s thickness of 0.188 inch

I is used for the side and top panels•

| 3q0-099-003 IV-4 O0000001-TSE11

!

I_ Ir"ll_l I tt|_-I I(;.(_)P'II_.I_ ' (:_)ml.an¥ I The estimatJ, u_ method of l_efeFence 26 Wa_ used to verl]:y 1:|1(' ,,.-

i o[use].ageperating weight.in a higherInasmuchspeed asregimethis ruethodthan the _s Modelbased 300,ol] n_rer_1 ;lllSs t t, l.t: that the results will be eonservative.

I Thethe estimatinsum of theg methodbasic weightconsidersrequiredthe totalto providefuselageminimumweightskins,t(, bu stringers or longerons, and circumferential stiffeners to resist

i basicconcentratedflight loads andph.ns redistributeweight panaltiesaroundincurredcutouts toandsupportthrough joints. Basic weight, FB, is defined by the expression

I F B = i.!23 S + f (Nz, Q, L, h)

I The 1.123 constant is based on a minimum skin of 0.040-inch 7075 aluminum plus 0.038 equivalent gage to account for stiffeners

I square(i.e., 0.078 inch x 0.i0 pounds/cubic inch x 144 square inches/ foot equal 1.123 pounds/square foot). Because the Mode] .... 300 loadings and design permit use of rainimum skin thickness of 0.020 aluminum, this factor was reduced to (0.020 + 0.020) x _--_' 0.i0 x 144 = 0.58. Therefore,

I FB = 0.58 S + f (N Z, _, L, h) (with the function "f" taken from Figure IV-2) = 306 pounds

I Penalties were then determined as shown in Table IV-t_.

I FUSELAGE TABLEWEIGHTIV-4PENALTIES

I Nose Gear Penalty 33 Bulkhead = 0.00025 W NL 5

I Body Cutout = 0./4 ib/in x 38 in 15 Door = 2.0 ib/sq-ft x 4.1 sq-ft 8

I Door Mechanism 5 Main Gear Penalty 19t

I Bulkhead = 0.O01 W NL 21 Body Cutout : 0.8 ib/in x 58 x 2 93

i Door Mechanism-- 2.0 lb/sq-ft X 12.5 sq-ft x 2 3250 Canopy and Windshield Penalty (from 166 I Figure IV-2)

i 300 -099 -(](]3 _V- 5 00000001-TSE12

i

I (_ll_ BELL H EL.I_](}PT_N COMPANt

I TABLE IV-q Continue(]

! Cockpit Penalty , 11.6 Bulkhead = 2 lb/sq-ft x 25 aq-ft 50 I Body Cutout : 0.5 Ib/In x 74 in 37 Floorin E = 1.0 Ib/sq-ft x 29 sq-ft 29 FB X I Production Joint Penalty = 0.025 x J 16 Tail Support Structural Penalty = 0.[5 x WT 4l

I Wing= 0.0005AttachmeNnZt W Structural Pe_lalty 24

I EquipmentPenalty =Support0.5 ib/cu-ft x 20 eu-ft i0 Cargo Floor Penalty = 1.0 ib/sq-ft

x 47 sq-ft 47 -_. ,.; Door Penalty 31 1 Body Cutout = 0.4 ib/in x 28 in ii 1 z--_ , Doer = 2 ib/sq-ft x I0 sq-ft 2_

I Miscellaneous Penalty = 0.i x Total of Above Penalties 68

! Total Fuselage Penalty Weight = Fp, Pounds 748

Total Fuselage Weight = F B + Fp, Pounds 1054 Model 300 Fuselage Weight, Pounds 1055

| - , Parameters and symbols used in the above equations are shown

I in Table IV-5. TABLE IV- 5

-- l • FUSELAGE PARAMETERS AND SYMBOLS Parameter Value for

i or Symbol Description Model 300 FB Fuselage Basic Weight 306 ib

I S Fu_lelage Wetted Area 520 sq-ft NZ Ultimate Flight Load Factor 5.0 _ 1 Q Weight of Fuselage and Controls 3483 ib |

I 300-099-003 IV-6 00000001-TSE13

I {_ I]I'I I IIF:I It-7(Z)I_TITR _:_MI'4NY

II TAI;1,F: IV-5 _ Contl.nued, .

I 1, Fu,_cl.nge Leugth 38.1 ft h Fuselage Depth 6.2 ft

I W Design Gross Weight 9500 !b Fp Fuselage Penalty %_eight 748 lb | N L Ultimate I,a_,d_.ngLoad Factor 2.25 WA Windshield Area 56 sq-ft J Production Joints 2

I WT Tail Group Weight 2q5 lb

: ,, _ L , ...... f ,.

I E. Alighting Gear Group

"; I Theretractable,laDding getricyclear of theconfiguratiModel 300on iswitha hydraa udual-wheellically operated,nose gear _ and a single-wheel main gear.

f I Gear structure weight was taken from a gear design layout and

' stresswere estimatedanalysis. fromHydrauliclayout drawings.system and Rollingcontrol gearsystemcomponentsweights

I caandtalogtheirdata,associatedare shownweights,in TablewhichIV-6.were taken from vendor

I ROLLING GEARTABLE COMPONENTIV-6 DATA

I Nose Gear Main Gear Item 'Number Size Weight Number size We igh_- (ib) (tb)

I Tire and Tube 2 5.00 x 5 12 2 8.50 x 10 5_

I Wheel 2 5.00 x 5 7 2 8.50 x i0 26 Brake - lq

Total 19 91

I A comparison of the Model 300 gear weight a_ a percentage of design weight with similar data for curreot generation V/STOL . f I alrcra t is presented in Table IV-7. I

I 300-(}99 -003 IV-7 00000001-TSE14

l

i I A)}I,I. IV-7

ALIGIITING GI_R GROUP WEIGIIT COMPARISON Design Gross Gear Group

I blodel Weight (ib) Weight (lb) Percent X0142 37474 1211 3.23

I X22A 14500 432 2.94 XV-5A 9200 420 4.56 XV-gA 7200 _l 4.04 I 300 9500 350°° 3.68 I F. Controls Group I The Model 300 flight control system consists of conventional ,_'_l_, helicopter cyclic and collective controls and airframe aileron, flap, elevator and rudder controls, along with phasing controls

J I which provide the proper combination of the two systems during _ _:, • transition from helicopter to airplane mode.

_' All rotating control components and all other critical compo- i nents were sized by a preliminary stress analysis and weights were estimated from layout drawings of these components. Weights

and existing hardware weights. . for other control system components were based on routing lengths

G. Engine Section and Nacelle Group

I This group includes the engine mount, firewalls and wingtip- mounted pylon cowling.

I One PT6C-40 engine is mounted at each wingtip in the pylon assembly shown on 300-960-003. The engine face is bolted directly to the transmission case. Weights for the engine support l were estimated from layouts. Firewall weight includes the forward, aft, upper and lower induc- tion firewalls. Aluminum extrusions which support the " firewalls, and%_ighactts wereas supportsbased onfor0.020the cgwling,tltanlum arefor websincludedand sintiffenersthis welght.

The cowls consist of al.uminum skins and support structure. Weights were based on gages determined by a preliminary stress analysis and were estimated from layout drawings.

I 399-099-003 IV-8 00000001-TSF01

I I

I II. Propulsion Group -

1. C_neral Description Power is supplied by two Pratt & Whitney PT6C-40 engines. 'rhis

andpowerplanetaryis transmittedgear stagesto thein proprotorsthe main transmission.by means of herringboneSpiral bevel gears are utilized in the interconnect shaft system which

providesof a singlepowerengineto bothfailure.proprotorsA centerfrom gearone enginebox, within thea seteventof bevel gears, is provided in the interconnect driveshaft system to account for the wing sweep angle.

2. Engine Installation

i I andfacturer'sinstallationspecification.hardware wereEstimatedadded weightsto completeof residualthe 724-poundfluids I total for two engines. i I[_ Dry3. weightsConversionfor Systemthe PT6C-40 engines were taken from the manu- u li,_ 1 Thismotor systemat eachconsistspylon. ofAn aninterconnectedAcme screw actdriveuator systemwith a ishydraulicincor- |£ porated to provide for operation of both actuators simultaneously

l andin theto provideevent of powera hydratoulicboth syactuatorsstem failufromre. oneA smhydraulicall control-motor phasing gearbox is provided at the cen_ .r of the interconnect shaft. Weights for this system were estimated from layouts and I taken from vendor catalog data. 4. Air Induction System

I Engine inlet air is inducted from an opening adjacent to the splnner immediately aft of the proprotor rotation plane. Weight

I forthe thisinductionfiberglassby-passductejectorstructwasure, estimatedseals, screeos,from powerplantblower and layouts.

I 5. Exhaust System An 0.030 stainless steel exhaust pipe with an 0.030 stainless

I e_ectorsteel turningbaffle,vanewhichis isclampedconsideredto eachpartengine.of the Theexhaustexhaustsystem, is 0.020 stainless steel. Weight of the exhaust pipes, attach-

i ment r_n_, ,urning vane and exhaust ejector baffle Is 16 pounds. 6. LubricaZion System Engine

I Thisware withsystemthecontainsoil tankan asoilan cooler,integrllfilter,part ofplumbingthe engine.and hard-The blower is carried _n the air induction system. _ights.were

I based on equivalent-slzed components used in existing alrcr_ft.

I 300-099-o03 IV-9

,i= O000000i -T.£ Ff_9

! I_ BE_I I HEI .IC()PI"E_ c:_)_rAN_ ! 7. Fuel. System _.

I TheBladderModelceils300 hasare loctwoatedseparatein eachfuelwingsystems,as shownone onfor Drawineach Kengine. 300-960-007. The fuel system weight was estimated from a

I ofcontiueiractor-developedcapacity as showncurve onforFigurefuel systemIV-3. weightFuel systemas a weight,function based on 246-gallon capacity at 0.30 pound per gallon, is 74

i aponunds.additionalDue to17-poFAAundrequpenaltyirements wasforaddedtwin-enginefor a totalinstaestimllation,ated fuel system weight of 91 pounds.

I 8. Engine Controls The engine control system primarily consists of droop compen-

I sationthese weandre estimatepower-leverd usingcontrolssimilarfor _{-ieach componenengine.t Weightsweights forand allowing for the increased cockpit-to-engine routing distance. ; I 9. Startin_ System

_ The 200-ampere starter-generator provided at each engine is

I on equivalent-sized components used in existing aircraft.

: _ • 10.poweredRotorby Pitchtwo 13-ampere-hourGovernor Controlbatteries. Weights were based _W | The rotor pitch governor is an eiectro-hydromechanical system which maintains selected rotor rpm. Electronic equipment,

I pactarisonsuator, ofandsimilarassociateditems linkageused inwereexistingestimasystedtems.based on com-

I ii. Drive System Engine power is transmitted to each proprotor directly from the

i engineetary gethroughar stagesthe asmainshowntransmissionon Drawing by300-960-004.herringbone andDesignplan-data for the entire drive train, including the interconnect shafting and center gearbox, is given in Table IV-8.

I All main transmission cases are magnesium except for the ring gear (nose) case which is cast aluminum. A large pylon case,

I todriveswhichforthethefirewallsinterconnectand cowlsshaft aresystem;mounted,the housessteel spindlethe bevel attaches to this case. ! ! ! 3OO-1199-OO3 IV-lO ! 00000001-TSF03

I I TABLE IV-8 ."

I DRIVE SYSTEM DATA

Speed Design Max Cont I Component (rpm) (hp) Torque (in-lb) Engine Output 30000 2150 2412

I Main Transmission 458 946 130127 Output

I Interconnect Drive 6485 633 6144 Shaft il ,I Gear Ratios

' I Interconnect Driveshaft 0.2162 I_ EngineRotor ShaftOutput 0.0188Basic

_i_,_ Hydraulic Pump Drive 0.2010

Engine Nil Governor 0.1400 I Center Gearbox 0.2162

I Gear Stage SpDeateda Max Cont Sta_e (rpm) Torque (in-lb)

I *Engine Output 30000 2412 *Main Transmission Input 8487 8532

I First Planetary Output 1776 33557 Second Planetary Output 458 130127

I *MOutpainutTransmission Spur 8487 I)704 *Main Transmission Bevel 6485 6144

I Output

*These components designed by single engine operation

I power. w_

i Weightspreliminaryfor thelayoutsModelby 300use driveof empiricalsystem werformulase estimatedand, wherefrom possible, by comparison with similar components in present use. I

i 3o0-099-003 IV-ll O0000001-TSF04

!

BE|-[- HELICOPTER Co_ANY I 1'he main transmisslon weights were verified by the estimntln_ method contained in Reference 27. Gear stage weights were rend

I welghts_rom Figurewere IV-q,read whichat the wasuppertakeendn fromof theReferenceweight range27. Alland then reduced by ten percent.to account for improvements in materials and methodology occurring since the . written. The i total of these reduced weights was multlpliedreport wos by the factor for magnesium c@ses, given in the report, to derive a basic housing weight. Welghts were then added for special mounting and shape I provisions and accessory drives. . These %_ights, taken from design estimates, are shown in Table IV-9. I TABLE IV-9

TRANSMISSION SPECIAL PROVISIONS Case extension to Support engine flrewalls and cowl and attach conversion spindle 56

J Freewheeling unit 7

- _ Lube system (with oil) 36

Torquemeter 17 i Accessory drives i0 Support installation (spindle and bearings) 39

TOTAL, Pounds 165

! Figure IV-5 shows the main transmission gear stage schematic and gear stage weights derived from Figure IV°4, along with a tabulation of the data used and the summation of total estimated

weight.weight by the veriflcation method as compared to Model 300

I I. Instrument Group The znstrument group consists of engine, flight and navigation

i IV-IO.instruments,Weightstranforsmitters,the instarndumentsinstallationsand transmittersas shown werein Tablebased on those currently in use on present day helicopters. Instal- lation weights are assumed to be the same except for wiring,

I whichbetweenhasthebeencockpitincreasedand propulsionto compensategroup.for greater distance ! !

I I 300-099-O03 IV-12 O0000001-T£FN5

I (_ i_l_:|.L HEL|CO|_TEIq f:OMI_AN Y

I TABLE IV-10 INSTRUME,N_r GROUP WEIGfrTS q'

Indi- Trans- Instal- eators mitters lation "l'oi:,_I

I AltimeterInstrument Noi. (ib)1.5 (ib) (lb) ([b)I. 5 Airspeed 2 i. 7 0,9 0.9 3.5

Cl OCI( D I I 2 i 0 i 0 Standby Compass i 0.8 0.8

I AngleVerticalof AStpeedtack 2 2.52.4 22.5.4 Turn and Slip 2 3.8 3.8 I I Attitude i 2.8 2.8 , m Vertical Gyro 1 6.0 9.0 4.2 19.2

_- _ _ • Oyro Compass I 6.5 7.3 q.8 18.6 o_ • Outside Air 1 0.2 0.2

_ Temperature " I Fuel Flow 2 1.8 1.8 [] Transmission oil 2 1.8 2.0 3.8 Pressure

I Engine Oil 2 1.8 0.4 2.2 Tempe rat ure Engine Oil Pressure 2 1.8 2.0 3.8 I Fuel Pressure i 0.6 1.2 1.8 Transmission Oil 2 1.8 0.4 2.2

I GasTemperatureProducer 2 1.8 1.6 3.4 Tachometer

I Fuel Quantity i 0.5 3.0 3.5 Dual Torquemeter 2 1.8 2.0 1.2 5.0 Triple Tachometer I 4.4 2.4 1.2 8.0 I Hydraulic Pressure 2 1.8 2.0 2.0 5.8 Turbine Inlet 2 1.8 1.8

I RPMTemperatureWarning i 0.3 2.2 2.5 Position, Flap 2 2.2 0.i 0.5 2.8 Gear I Position, Main I 0.3 0.9 2.0 3,2 Conversion 2 1.1 0. i 0.5 [ •7

I TOTAL INSTRUMENT GROUP WEIGIIT 109.6 I I 30O-099 -003 IV-13 00000001-TSF06

I J. !I_,draulic._ Gro.p

[lydraul].c power is utilized to e×tend and retract the lnnHnp[

i _ear, to. power the aileron py]on conversion actuators, m_d( " Io, provide boost capability in the proprotor c_clic and co].]eetive control systems. The Model 300 has a completely dual 1.500 ps¢ system. i hydraulic Only the _ghts of the pumps, reservoirs, accumulators, valves, and inter"connecting plumbinE are _ncluded in the main system weight. Weights of components and plumbing

I ofprovidingthat syspowertem. to a specific system are carried in the weight

I K. Electrical Group The ac-de electrical system on the Model 300 is powered by starter-generators attached to the engines. Two 13-ampere-hour I batteries are provided to furnish power for the starter- generators. Weights of these and other major components are based on vendor data. Wir£ng and hardware weights were estl-

"-_"._[Lt I matedthe componentsfrom wiringare diagramslisted inandTableroutingIV-II.layouts. The weights of TABLE IV-If

,. Weight (ib)

I DC S),stem Batteries 48 Battery Installation 2

I TransformerVoltage Regulator 64 Switches, Rheostats _nd Panels 4

I Relayswiring and Miscellaneous 8719 Equipment Supports 16

I AC System Inverter 26

i AmmetersSwitches, andRheostats,Voltmetersand Panels 262 Circuit Breakers and Fuses 10 Junction and Distribution Boxes 3

I WiringRelays 8nd Miscellaneous 141 Lights 15

I TOTAL ELECTRICAL GROL_ WEI_{T 283

Electronics I L. Group Electronic equipment consists of AN/ARC-If4 V]IF-F_| and AN/ARC-I] 5 • s i VHF radio , Collins 613L-2 transponder syste_n and a four-station

I 300-099 -003 IV-14 00000001-TSF07

J I_ BELL HFI ]COPTER C:[I_,I'ANY

l ICS, C-6533. Weights for these systems _ere taken from ,- existing installations used in current aircraft.

l M. Furnishings and Equipment Group

I Thefurnishingsfurnishing miscellaneoand equipmentus equgroipment,up inch:desand emergencycrew aeeoeqmmodation,_,ulpm_nt The weights shown in Table IV-12 were based on similar equipment

i presently in use on Bell helicopters.TABLE IV-12

i FURNISHINGS AN_O EQUIPMENT GROUP WEIGHTS

i Accommodations Weight (Ib)

: _" I Crew SeatsSafety Belts 626 _" '. Crew Shoulder Harness and Inertia Reels 6 ._ ]J Miscellaneous Equipment |

_, Windshield W_"per 14 I InstrumentConsoles Panel 1520

i Furnishings Soundproofing (cockpit) 15

Fire Detection System 5 Portable Fire Extinguisher 7 Engine Fire ._xtlngulsher 42 i II Emergency Equipment EQUIPMENT I TOTALGROUP FURNISHINGSWEIGHT AND 192

I N. Air-Conditioning Equipment Group The air-conditioning system of the Model 300 is used for forw_ird window and and of the ment. Thedefoggingenvironment heatcontrol_ng unitcoolingused is the cockpitsame as eomparthat nowt- i_stalled on the Bell Mode]. AH-IG. ! !

I 300 -0'99-003 IV -I5

O0000001-TSF09

I_ BELl.. HELICOPTER (x_MPANV

FUSELAGE BASIC WEIGHT PENALTY

2O

o H

1 1.0 _ 2.5_ . 3.0 3.5 ___ NZ QL 1![ _7_o_

_ WINDSHIELD OR CANOPY WEIGHT 3O0

i

H 200

I | _ I _ _oo

| _ O, 30 qO 50 60 ! TRUE AR_ - WINDSHIELD OR C_OPY - FT 2

I Figure IV-2. Fuselage Weight Estimation Parameters. ! 300-099-003 IV-17 -I 00000001-TSF10

I _ BFI .I. FII_I_ICC)pTIqF_ ¢':c)A41"ANY I 8.- I I I s,o I IIIIII ] I I IIIIII ! I I i_ ' ' lllllm SELF SEALING ' ' = ;. I ,,lllll= PA,T,A_'SELF_E,' ,L,NG

UH-ID • ," _- -_C119C BSOD-- _] _o '.0.J"= SELFI S_ALI I I I I'I,NGD --I C m ! _ ! , =L-?3ow f I 111,.I II Ic'47 I - i _ ....,.,.U,_,H-1B---.--Jr_.J-PARTIAL SELF SEALING_--_...--. -C-

I _ _...9.. _1__'},"TIll%_ I I I .!-"m_._'"_ c__!__I -i L _"-- I i_oo _....:-_'._____. " -,30,- I , ' , c(p11_ :" I_I I o.1 , I L 2o0 ,ooo L 'lll 1oooo,

FUEL CAPACITY - GALLONS

I Figure IV-3. Fuel System Weight. I I I 300-099-003 IV-I8

O0000001-TSF12

1 I

1_ BFI-I. HELIC_O|_TER C=OMPANY

PROPROIOR

! a t /I ENGINE I(_. ,_ INTERCODRIVESHAFTNNECT | - Gear Stage "Design Output/Side Wo/S_e Torque '-_,--: _ I Number Type (hp) (rpm) (in-lb) (ib)

:"_' :_ • Engine i150 30000 2412

Idler *6 _, I 1 Spur & 1150 8487 8532 20 26 2 Planet 946 1776 33557 29 m 3 Planet 946 458 130127 74

i 45 BevelSpur 633633 64858487 46144704 2136

I Weight DataSide LMain Transmission Item (lb 9

I Gears, bearings, shafts, etc. : W o 151 (derived from Figure IV-4 with

I a 10% reduction) Housing : 0.372 WG 53

i TSopectalialweightprovisionsper side 316569 Model 300 weight per side (estimated) 366

| *Left transmission has two.....idlers; rig ht transmission, one. Weight shown is average for both sides. !

I Figure IV-5. Main Transmission Schematic and Weight Data.

i 300-099-003 IV-20 00000001-TSF13

I (1_ B_LL HELICOPTER _:r_

I V. PERFORMANCE _ •

I A. Summary The performance characteristics of the Model 300 are summarized in Table V-i for four mission weights. These missiorls are: m flight resea2ch, simulated eLvil mission, simulated _ilitary mission and ferry mission.

I The aircraft has three-bladed 25-foot-diameter proprotors with 14-inch chord blades. Disc loading is 9.7 pounds per square foot at the normal gross weight of 9500 pounds. Power is

I tsuppliedurbine enginesby two wPrattith takeoand lWhif tandney 30-minPT6C-40ute direcratint-drivegs of i150free- horsepower and a normal power rating of 995 horsepower. The

I broadpermitsrangefull ofpowerrpm forto beefficientextracted operationat maximumof therpm freeduringturbinetakeoff in helicopter mode and near optimum specific fuel consumption to be obtained in airplane cruise at reduced rpm. Proprotor tip per I speed is 740 feet second in helicopter mode and 600 feet "' per second in airplane mode.

[_ _ At 9500 pounds gross weight the aircraft hovers out of ground _:i S effect at ii 600 feet on a standard day and 6400 feet on a 95°F _ day. At this weight the service ceiling is 26,000 feet. Max- I feetimum speedand serviceis 312 ceilingknots atks 15,00020,000 feefeett. atHoverthe maceilingximum gross_s 2600 weight of 12,400 pounds. A STOL takeoff can be made in 900 feet at 4000 feet on a 95°F day at this weight. Maximum single engine speed is 215 knots at i0,000 feet for this weight. Maximum range for the research mission _ncreases from 397

I feet.nautical Ferrymilesrangeat seisa 18level70 nauticalto 630 naumilestlcal atmiles20,000at feet.20,000

Thcivile perfandormamnceilitaryof VTOLthe Moadpeplil c30ations0 is suffiA cienttypicalto civildemonstratemission could be demonstrated carrying the equivalent weight of eight passengers end 240 pounds of baggage at a cruise speed of 260

kno(350ts statu(300temilesmiles).per hoAnur)aircrewfor a recrangeoveryof mission306 nauticalis selecmilested to demonstrate military application. With a crew of three and 550

I po300undsknotsof aatrmori0,000and feet,rescue pickequipment,up threethedownedaircraftairmencan atdasha at distance of 375 nautical miles, make a hovering out-of-ground

i effect takeoff at 4000 feet 95°F and return. B. Airframe Aerodynamics

I MarchA one-f_f1969th inscathele Ling-Temco-Voughtmodel of the Model 7-by-10-foot300 airframe low-speedwas testedwindin tunnel (Figure V-l). This test will be reported in Reference 28. The model was tested with a smooth finish and also with artifi- I cial roughness. The roughness was introduced to insure that

300-099-003 V-l O0000001-TSF14 00000001-TSG01

1

I (_ BELl, HELICOPTER comrAN_

! boundary layer transition was consistent with that anticipated _. on the full-scale vehicle.

I All airframe components, with the exception of the extended landing gear were tested in the wind tunnel. Full-scale llft

I andReynoldsdrag characteristicsnumber correctionswereto determinedwind-tunnel by daapplicationta. Since ofthe fixed elevator settings in the wind-tunnel test did not represent the full-scale Model 300 trim setting, values for trim, lift and were I drag obtained by computation. i. Lift Analysis

I Both model and full-scale lift curves are given in Figure V-2 for flap settings of 0 and plus 30 degrees. Comparison of the

I modelthe Reynoldsand full-nusmbercale oncurvesthe liftshowsdatathatwerethe amajornegativeeffectsshiftof in angle of zero lift and an increase in the lift curve slope. The minus 1.4 degree shift itl the angle of zero lift was due to the I Reynolds number effect on the 23-percent thick wing section. This unusual behavior was determined by extrapolating the data from Reference 29. i

in Figure V-2 was due to an increase in the wing lift curve I The 8.5-percent increase i_ the overall lift curv_ slope shown I slopein Figuafterre V-2correcis antionincreasefor Reynoldsin wing numbermaximumeffects.lift coefficienNot apparentt. A 0.15 increase over wind-tunnel data was estimated when Reynolds number corrections were applied, but this effect was offset by

I mentsthe elevatorfor lifttrimwereforcemade reqtouirements.the fuselage,No pod,Reynoldsand empennagenumber adjust- lift curves.

I 2. Drag Analysi_

Full-scale airframe drag was computed from wind-tunnel data by each individual for I correcting the value for component Reynolds number differences between the model and the full-scale ship. The resulting total airframe drag coefficient is shown in

I data.Figures TheV-3 methodand V-4usedand tocomparedcompute withthe ftheull-scslemeasured dragmodelis tesout-t lined below.

I a. Lift and drag values for each component were obtained by com- paring different test "runs" with and without the component being evaluated.

b. The profile, or parasite drag coefficients, were obtained by subtracting the induced drag from the measured total for each i component. The induced drag coefficient was cslculated from CDI = CL2/_eAR w

I e = 0.90 300-099-003 V-3 I 00000001-TSG02

I BELL HELICOPTER COMPANY

I c. Reynolds numDer corrections were applied to the model profile r drag coefficients to obtain the full-scale coefficients. Corree-

I tion factors are shown.to Table V-2. d. The corrected profile drag was then added back to the induced

i drag at a given lift coefficient to obtain the final component value. . . e. The profile drag coefflclent for the elevator was obtained I from Reference 30, using interpolation for the desired airfoil section. In order to account for roughness, the drag values from Reference 30 were.increased first by "fairing out" the drag

I "buckets"25 percent.and Thetheninducedby Increasingdrag for thethe resultantelevator minimumwas obtainedCDo by from calculations based on the required llft and for flight eon-

'_ I ditlons. f. The total full-scale aircraft drag values were then found by summing the component values plus a five percent increase in I total profile drag to account for leakage protuberances, irreg- r_4_r, ular surfaces, etc. A full-scale drag breakdown in terms of

._/. equivalent flat-plate area is shown in Table V-3 for s lift In Figure V-4 it ts seen that the minimum drag for the full- scale flaps up condition does not occur at zero lift but at CL = 0.20. This comes about because of the fuselage and pod effects on total llft. The wing alone minumum drag occurs at CL = 0.07, but w(th the addition of the fuselage and pod, the

I minimumcarried onshiftstheseovercomponento CtLs = for0.20 anglesbecauseof ofatZackthe negativeup to 4 degrees.lift

i Figumodelre testV-5 comparesdata with thethe aircdragraftpredicteddrag as bydetconventionalermined from dragthe estimating methods. Predicted parasite drag area is 6.4 square feet with a resulting CDo = 0.0364.

I For the helicopter mode of operation an additional flat-plate drag area of 4 square feet and 32.9 square feet were added to

I atively.ccount for the extended landing gear and vertfcal pods, respec-

C. Proprotor Performanc_

I i. Descriptipn

I chord,Each of thickness,the 25-foottwistdiameterand amountproprotorsof camberhas threechangeblades.as func-The tion of the radius. These d_steibutiona are shown in Figure

I lll-1the bladeand areinto simulatedfour sections.in the Theseperformancedlvislonsanalysisare madeby dividingat the nondimensional blade Stations of 0.075, 0.45, 0.70, 0.90, and % ! 1.00. The blade is not represented between Stations 0.0 and

I 300-099-003 V-4 00000001-TSG03

300-099-003 V-5 00000001-TSG04

I

BELl- HELICOPTER COMPANY

i 0.075 as this segment is contained within the spinner. For cai- culatlon purposes, each of these sections _s further subdivided

I such that the entire blade _.s represented by 22 blade "strips". Airfozl sectlon characteristics were estimated fo_" each of the

i fotesurts bladeof Referenceseetlons 31.based.onF1guresV-6the BellthroHelicopterugh V-9 showairfollsamplessectofion the airfoll sectlon data for Mach numbers of 0.3, 0.5, 0.6 and 0,7.

I 2. Method and Correlat{oc All proprotor performance _s obtained w_th the a_d of the Bell

I Hellcgptera hellcopter compperformanceuter programcompFu3ta5(J).tlon Thisprogramprogandram is_sdescribedprimarily in detail in Reference 32. The validity of this program has been

i pdatarovenforin botheth pofast-theby fllghthe texc_glmesellent ctorrhatelaationnormalw_thelicopterh fl_ght-test rotor experiences: hover and helicopter forward flight. TABLE V-3 "" ,_ , ! DRAG BREAKDOWN - FULL SCALE AIRCRAFT Component Flat Plate Drag Area 1 I Fuselage 1.60 i Wing i. 57 Vertical Tall 0,55 i Horizontal Ta_l 0,59 Pods - Horizontal 1.37 Miscellaneous (5% Parasite) 0.28 I Airplane Parasite 5.96 Airplane Induced (CDi) 1.6___22

I VerticalTotal AirplPod aneIncrement 32.90 7,58 Flaps Down In_rement 8.63

i HelicopterUndercarriageParasite° 4.00 51.49^ Hel_copter Induced 2.82 2

I i Total Hei_copter -- 5,.31 i i. These values reflect flight conditions that exist for

i a250grossknoZsweightin theof alI0,300rplane pounds,mode, orandseaiO.000level,feet50 altftude,_nots In the helicopter mode (airframe CL = 0.38).

I 2. Induced drag cCL2alc/u_eAR.lated onWherindividuale e = .components based on the relation 0.9

1 3. Estimated undercarrlage not tested...

i" I 300-099-003 V-6

• • r, 00000001-TSG05

I B LL. ucoPTcoMPanY .

I Correlation of F35(J) with proprotor performance _n the airplane mode is shown in Figure V-lO. The test results shown are for a

i This13-foottestdiawabmeterconductedBoe_ng-Verintothel NASA-Amesproprotor mo40-by-80-footdel with "E" windblades. tunnel. Blade parameters for the proprotor are shown _n

i obtainedFigure V-ll.by usingThe calcbothulateda limitedvaluesamouusednt offorBoeing-Vercorrelationto[ weresection data and a set of Bell Helicopter drooped airfoil section data. The variation in the two sets of calculated results show the

i thatimportancethe compofuterusingprogramthe proF35(J)per sectioncan accuratelydata. It predictis conclprop-uded rotor performance for any flight regime when the proper airfoil

i section data are used. 3. Isolated Proprot0r Performance

4, I Figures V-12 through V-15 show proprotor performance in hovering II and alrpiane flight. The hovering performance is shown in Figure V-12 in the form of power required versus thrust. Per-

i formance in airplane flight is shown in Figures V-13 through _ V-15,proprotorin theshafformt horsepower,of propulsivefor seaefficiencylevel, asi0,000a functionfeet, andof

i these figures for the steady level flight condition at a gross • weight of 10,300 pounds. 20,000 feet standard day operation. A dashed line is shown on i D. Powerplant Performance Figures V-16 and V-17 show the power available and fuel flow,

i performance.respectively, Figforuresthe V-18hoveringand V-19and helicare opcarpeter t forwaplotsrd forfl_ghthet airplane forward flight power available as a function of air-

i speeratedd andpower.altitAlsoude onshowan standardon those dayfiguforres takeoffis the and1720 normalhorsepower design torque limit. A typical fuel flow carpet plot is pre- sented in Figure V-20 for standard day operation at [0,000 feet.

i Tlosses.he power available numbers reflect all engine and transmission

i TtheableinstV-4alledshowsenginethe lossperesformanforce.each Thenginee powerusedavailablein estimatingand fuel flow were obtained by inputting these losses into the computer

i program supplied by the engine manufacturer. The ram efficiency factor (_RAM) was determined from Section IX, Figure 18 of Reference 33 and is expressed by the relation

i _RAM 1.0 - 14.4 \_,Vl

I where V is the velocity in knots, W a is the weight flow of air in pounds per second, and _' is the air density ratio. !

I 300-099-003 V-7 O000000]-TSG06

L i

i (1_ BELL H E LICOpT1ER COMPANY

I TABLE V-4 _'

I ENGINE LOSSES

Inlet Pressure Loss 5.8 in H20 SLS

I Exhaust Pressure Loss 3.0 in H20 SLS

I Inlet Temperature Rise 2.7°F Ram Recovery Loss (l - TRAM ) q in H20

I SLS Extracted Horsepower Loss 1]..5 hp

I Twin Engine Transmission 0.98 Efficiency

I SingleEfflhiencyEngine Transmission 0.97

_--_! I E. Helicopter Performance

I andAll ohuet-of-ground-effeclicopter performancet condiis tions.calculated for 30,000 engine rpm

l. Hovering Performance

! ° Hover ceilings are shown in Figure V-21 for standard and 95°F day conditions. Included in these data is a seven-percent I download that is experienced by the vehicle in the hovering mode. The effect of this download is made evident in Figure V-22 which shows both the isolated proprotor and the overall Flgure of Merit

I Figas uarefuncoftionMeritof isthea gross0.98 transmissiweight. onAlsoeffiinclcienudcey,d _n the overall

All data shown in Figures V-21 and V-22 were derived from the in V-12 and the I isolated proprotor data shown Figure power available data shown in Figure V-16.

I 2. Helicopter Level Flight Power Required Figure V-23 shows the helicopter level flight power required

I forrangesea-levelof 8500 standard-dayto 13,500 pounds.operationsIncludedand forin thesea grossdataweightis the sharing of the total vehicle lift between the airframe a_d proprotors as a function of airspeed as shown in Figure V-24. power I accountA 15-percentfor thedecreaseslde-bM-sidein rotoreffectinducedof the proprotors.was also usedAll to data were determined for a 15-degree mast angle and a 30-degree

I flaps down condition. ! 300-099-003 V-8 00000001-TSG07

i ,, i

I (_ BELL HELICOPTER comPANY

I 3. Rate of Climb

I Standardfunction dayof almatitudeximum isratesho_..1of climbin Figureat normalV-25 ratedfor a powerrange asof a weights. These values were determined by using the relation:

I R/C (helicopter mode) = excess power x 33000 x 0.85 ft/min gross weight

I whereavailableexcessand pominimumwer is powerthe differencerequired. betweenThe 0.85normalfactorratised a power climb efficiency number.

I F. Airplane Performance

i. Thrust Hqrsepower Required

I Data for thrust power required versus airspeed for sea level, i0,000 feet and 20,000 feet on standard day conditions are • shown in Figures V-26 through V-28 for several gross wights. I These data were determined using the airframe lift and drag -_";_ coefficients shown in Figures V-2 through V-4 and the

I . • expression: i I DV 55O Thrust horsepower required = I where D = CD qS

l _ I Figures V-29 through V-31 contain the above mentioned thrust I horsepower required data expression as proprotor shaft horse- power. The relation used for this converslon was:

_-_ Proprotor shaft horsepower required = thrust horsepowe T rgquired

I Wpro p

where _-rop is the proprotor propulsive efficiency, obtained I from Figures V-13 through V 15. 2. Fli_ht Envelope and Maximum Speed

I Figure V-32 contains the standard day airplane flight envelope for normal power rating and the maximum speed for takeoff power rating, both for a gross weight of 10,300 pounds. The low-speed the is stall the I end of flight envelope governed by wing and high-altitude end by the absolute ceiling. The maximum speed was determined by the intersection of the power available and

powerpounds.required curves for the given gross weight of 10,300 ! ! 300 -099-003 V-9 00000001-TSG08

I

(_ BELl- HELICOPTER cOMrANV

I 3. Maximu____Rat____ee___mm of Climb

I reRatelat_on:of climb in the airplane mode is calculated using the

I R/C (airplane mode) = excess thrustgrosshorsepowerweight x 3,3000

e where excess thrust horsepower is the difference between the

I thrustclimb forhorsepowernormal raav_11ableted power andis ploreqtutedired. versusThe maaltitudeximum raforte of various gr?ss weights and is shown in Figure V-33 for standard

day operatlons.

m 4. Specific Range . • Figures V-34 through V-36 contain nautical miles per pound of fuel versus true alrspeed for various gross weights at sea level, 10,000 feet and 20,000 feet on a standard day. These curves

%41_._..'_ I were prepared from the fuel flow and power required data. _-_r;_, 5. Single Engine Performance _i " 'r "._,: _ Figure V-37 shows the single engine performance for both the helicopter and airplane modes superimposed on a twin engine J _'-" hovering celllng. For this data a 150 and 200 foot-per minute rate of climb was maintained for the helicopter and airplane modes respectively.

6. Mission Profiles Mission capability of the aircraft is illustrated in Figures

i V-38 through V-dl. Miasion profiles are shown for the following: Research m_ssion

I civil mission Military mission

I Ferry mission

The civil and military missions are shown to illustrate the to I capability of the Model 300 demonstrate economic feasibility on simulated missions. The civil mission is an interclty commuter flight of 8 passengers and 240 pounds of baggage. Range

I 260is 306knotsnautical(300 mph).miles The(350 milistatutaryte missionmiles) atis aancruisealrcrewspeedrecovery.of A dash speed of 300 knots is used to traverse 375 nautical miles,

I Tpickakeoffup threeweight suinclrvivors,udes aandcrewretofurnthreeat long-rangeand 550 poucruisends ofspeed.armor and rescue equipment. At the _id-point pickup the aircraft can ! hover 4000 feet on a 95°F day or 9000 feet on a standard day. 300-099-003 V-lO ! 00000001-TSG09

I _ BFLI _ NEE|COpTER COMpAN _

I The research and ferry missions are performed at long-range cruise speed; whereas, the civil and military missions are

I mission.flown at higherPayload speedsrange morecurvesappropriateare shown forin Figuresthe particV-3ular3 an_ V-41 for the takeoff weights of the research and ferry missions.

I 7. STOL Performance STOL takeoff distance to clear a 50-foot obstacle is shown in

I Figfollowingure V-42takeofffor a techniqSTOL takeoffue is assat u4000med: feepylonst on aare95°Frotday.ated Theto a 20 degree conversion angle, maximum power and full forward

I cyclicacceleratarees appliedto 70 knoasts,the aftbrakescyclicare isreleased,applied theto rotateaircraftthe aircraft and lift off, climb out is made at 70 knots. The pylon tilt and forward cyclic tilt the tip path plane 30 degrees. The I 70-knot lift-off speed was selected to provide 1.25g maneuver capability at the maximum gross weight. A lift off at lower speed would decrease the takeoff distance. Takeoff distance is

• weight where the aircraft can hover out of ground effect. I 900 feet at 12,400 pounds and 700 feet at the 10,300 pound

I I • I I

I i

i

300-099-003 V-II I1_ I]FI I- HT:.I _l(_:()l "Tf:R _:r ;mr,anY

I °" O0000001-TSG11

i i

I i_ BELL HEt._ICOPTI__R rX_MI.AN Y

l 1.8 •

t.6 ,- "_

1.4 i i s+$fI ," I - / ; "," i _.c V/) l 1,4 £ I l i

7:_'<:I __>°° /,,_' 'I1I !/IX/// I ' ' o /!_ I I Y A" / I FORCEDTURBULENCE-

' /i ______WIND TUNNEL-- ! ., / / FULL SCALE _ /I ./J I I / I _ DATA ' ._ /kt'l1III!,I-

i . FUSELAGE ANGLE OF ATTACK Flgure V-2. Airframe Lift Coefficient Versus

_ I Fuselage Angle of Attack. " I

I 300-099-003 V-13 00000001'TSG12

Figure V-3. Airframe Drag Coefficient VersUs Fuselage Angle of Attack.

300-099-003 V-14 00000001-TSG 13

HELICOPTER ,--"?M I "A_ V

•16 .... /

_LAP S / / I .14 ___ DOWN 30" WIND TUNNEL \ / , FORCED !FULL SCALE ._ ,\/ // I ._ , ___ H_o "_ ,_

• I o .... _-- /7 /

CFULL SCALE _ WIND TUNNEL .06 " / / "_--DATA (SMOOTX)

I I 0 I -.2 0 .2 ._ .6 .8 1.0 1.2

i LIFT COEFFICIENT

I Figure V-4. Airframe Drag Coefficient Versus Lift Coefficient. I

i 300-099-003 V-L5 00000001-TSG14

[ (1_ BEI i HELICOPTER ¢:¢)MI_ANY

.18

.i6

i .... / I "_" /

| _ .10 FgRCEFULL _SaLETUrBULeNCEDATA \ /

' _ --

CDo = .0364 / AR = 6.63 / /_ .o6 • = .9 / /

I .04 _ f

.02 ! 0 i. i -.2 0 .2 .4 .6 .8 1.0 1.2

i LIFT COEFFICIENT Figure V-5. Comparison of Full-Scale Aircraft Drag,

I Data.Based on Tunnel Results and Calculated I I 300-099-003 V-16 00000002

00000002-TSA04

at _ F_FI_L.. HF|J(;OPTER C_)MI_ANV

I .18 i .... I!_/I i "_° ""A°_"_"_":'_

I ._ 1

' " "" I o .i0 I o t

, / ' 1/

._ .._

.02 I _ . I 0 0 .2 .4 .6 .8 1.0 1.2 1.4 LIFT COEFFICIENT

i FiKure V-8. ProprotorRadial StatiBladeon 0.Station70 to 0.90.Data, !

i 300-099-003 V-19 00000002-TSA05

(li_ BE LI_H E L|COPTI=R c;oMpAN,,,' I I .1E I I I .12 I

I o .i0

t.,_:I .08

.06 I

I .Og ! .02 !

I 00 .2 ,g .6 .8 1.0 1.2 l.g

I LIFT COEFFICIENT

i Figure V-9. ProprotorRadial StationBlade 0.90Sectionto 1.Data,00. I

I 300-099-003 V-20

ONNNNNNg_T_n7

I (_ BELL HELICOPTER (_'OMPANY I .4

I _. _ --DESI ON CL_

I I o I I 3O

" i _ "AERO TWIST

_.-- o_20 | _M f-_

I ,_LI 0 <..,#,_ I _ _o_ _ _ -

i := _< 0 " ' ,, m -I0 I 0 .2 .g .6 .8 i .O r/R I

i Figure V-ll. Boe£n_-Vertol 13-Foot Diameter Proprotor Blade Parameters % (Blade "E") . I

I 300-099-003 V-22 00000002-TSA08

I _ A_LI. HELICOpTER_MrANv

I 3600 I 3200

II 528oo_ / _.oo /

/ I _oo / o / I _ 800 / o_o_oRO_D_F_o_

I 400 I 0 I 6000 8000 i0000 12000 14000 16000 18000 20000

I THRUST/DENSITY _TIO

I Figure V-12. Proprotor Hovering Power Required Versus Thrust. l

! 300-099-003 V-23 00000002-TSA09

I I_ BELL HELICOPTER co_p^r_,," I 1 i _ _ _. I , ,. , /I I -'LEVEL FLIGHT CONDITION -GROSS WEIGHT = 10300 POUNDS ! 90 _6oI I_ _,_,._. .. 200 240 KNO_8 I \ _____._..j _'8_- ___ %_ _ooo ,..,I ,o L_" > _ /

7,,7-,,, ,I _ 60 +o,_

I ,. / =o / -Z,2,o

! _.o / STANDARD DAY

I 600 FT/SEC TIP SPEED 30 I

20 I 200 400 600 800 i000 1200 I_00 1600

I PROPROTOR SHAFT HORSEPOWER PER PROPROTOR

I Figure V-13. Proprotor Efficiency Versus Shaft Horsepower Sea Level. I

I 300-099-003 V-24 -- k 00000002-TSA10

I (_ B_ [.I_H.___F.L|COPTER (]nMPANY I g-, I i00 I | 90

I 80

I

70

. , ] Z

• _ 60 o

I 0 50

I I I

l 200 400 600 800 i000 1200 1400 1600 PROPROTOR SHAFT HORSEPOWER PER PROPROTOR I Figure V-I_. Proprotor Efficiency Versus Shaft i Horsepower i0,000 Feet. l I 300-099-003 V-25 00000002-TSA11

i

I ¢ BELL HL_I-ICC)PTER cQ_P,,',N',"

I I 1 1 ..

I AIRPLANE MODE i00 .I . I _ ,

%6C 2( 0 ---,-- _ .-_o___

, .o ,.i2S0 I //>_ >_ "_",_o1 >_Coo_o,_-

,"/_,,I _, ooIx/ / \ I _ / i l'/i' _o \ I a_ so-- -- fCONDITIONSL_VEL_ICHT______A GROSS WEIGHT = 10300 POUNDS

m JTA NDARD DAY

I 600 FT/SEC TIP SPEEE 3C I

I 200 400 600 800 I000 1200 1400 0 PROPROTOR SHAFT HORSEPOWER PER PROPROTOR I

I Figure V-15. HProprotororsepower Efficiency20,000 FeetVersus. Shaft I

I 300-099-003 V-26 00000002-TSA12

1

I (_ BELl- HELICOP'_ER ,:_',MI_NY I

I 30000 ENGINE R_

i HELICOPTER MODE I I i_ 24000

" " 2, 1 _oooo\\,\\\ I _16000

, \\ k 1 TAKEOFF POUR

I __2000 _. _o__ _o_ I __g8ooo \ k\ I 4ooo x x

, o ,,,\\,_ \_ I .00 800 1200 1600 2000 2400 2800 I TOT_ PROP_TOR HORSEPOWER AVAILABLE

I Figure V-16. Twin Engine Proprotor Shaft Horsepower Available, Helicopter Mode. I

300-099-003 V-27 0000OOOP-T._A ig

I (1_ BEI .I IIF_I .ICOPTI_R cnml-,aNY ! I I

I ii0 J 0

900

_i_ o_ 800

] _ 70o

J 600 ! 50G I

I 400

I Figure V-17. Fuel Flow, Helicopter Mode. I I

300-099-003 V-28 00000002-TSA14

I _ RF:_LL HF'I .IC(3P'I-ER ,:_Ml'an_"

|, , 1 3600 ...... Jr 24300 ENCINE P.Plg__ I STANDARD DAY

3201 _-

._ 280 ...... _,_0YS %60

I

I I 800 -_,_ I .oo I I I

I F_gure V-18. Proprot_r Shaft Horsepower Available, Takeoff and 30-Minute Power. I

300-099-003 V-29

O0000002-TSB02

Figure V-20. Fuel Flow, A_rplane Node. I

I 300-099-003 V-31

II I .... i ii i|. I II l r i i- i ii " " .... - -_ _1 ]..... I O000000P-TRRN."4

I @ BELL HEL|C_fOPTER ,X',MI'ANY I , t I --- OUT OF GROUND EFFECT_

J 7?o DOWNLOAD INCLUDE[ ; 30000 ENGINE RI_ --

mu - I

t oooo,\, \ .... HOT DAY

16000 \ _ X \ ' \ \ 12ooo _ \ \ \ H k

SINGLE \ _/ I < ---8000 =_" \ \// K" ' I k 4000 • \

\ \ \ o \\ _ \ | 2ooo _ooo 6000 8000 Ioooot2ooo_ooo GROSS WEIGHT - POUNDS I

I Figure V-21. Hover Ceilings. I

I 300 -099 -003 V-32 00000002-TSR04

I (_ BELl- I-IELICOPTEI_ c:oMr, aN,t

I 1.)1 I

i 1.0

I .90

' p---PROPROTOR FIGURE OF MERIT /GROSS WEIGHT = THRUST I .80 TRANSMISSION EFFICIENCY = 1.00- i i j. i_._< l .,o .._- --_-_

• 50 i, ! / OVE L FIGURE OF MERIT GROSS WEIGHT = THRUST/I.07 .40 TRANSMISSION EFFICIENCY = 0.98- I .30 I

I "206000 8000 i0000 12000 14000 16000 18000 20000

i GROSS WEIGHT - POUNDS I Figure V-22. Hovering Figure of Merit.

' |

I 300-099-003 V-33 O000000PIT.qRn_

I @ B_LL H_LICOPTER COmPANy

HELICOPTER MODE S_A LEVEL 1800 l I II/I

, H POUNDS Ill

-: ,- , . I _IZO0 _=_oo I//////I1

_,,'_I _ --_oo _ _.oooi1500 /2/// I _ _o_oo\\_/// i ssoo I Ow_ ooo ',._

! 4OO ! I I 200 0 _0 80 120 160 200 240 TRUE AIRSPEED - KNOTS I Figure V-23. Proprotor Shaft Horsepower

I HelicopteRequired r VersusMode. True Airspeed, % !

I 300 -099-003 V-34 I (_ BELL H E LICOPTER cx_,w-Ar_'r I _0 i\ .. I \

----PROPROTOR LIPT

1 oo \

i / l _ / ' 30 /

I _o., / /

I0 MAST ANGLE -15 ° _'-

1 °o 60 8o ioo 12o 14o 16o 18o TRUE AIRSPEED - KNOTS I

I Figure V-24. Liftand Airfr&_eDistributionin HelicopterBetween PrLeveloprotor Flight. I

I 300-099-003 V-35 O0000002-TSB07

: 300-099-003 V-36 O0000002-TSB08

I

I _ BE_-LL HELICOPTER t.'O.'_IpANV

1600 ---"

'! I !AIRPLANE!lMODE /

II _ 12oo

.... " I I , +o, I _ _oo i15oo__ ----_:.z'V//' E_ 1050q_ _ / I --9_00. /_,/ 8500 ._ _ / I 400 STANDARD DAY I

I 200 1 0

i 80 120 160 200 240 280 320 360 TRUE AIRSPEED - KNOTS

Figure V-26. Thrust Horsepower Required Versus True Airspeed, Sea Level.

300-099-003 V-37 O0000002-TSB09

I ¢ BFI.. i . H FI .ICQPTFR (;OMI..A,._ Y I ,_oo -/ ..

Y 1 o _o_o_ //I

1 _ _Too\ ./

_-,",,, _ 800 /. y/ 1 ',_._..L.,-- ___ I 400 STANDARD DAY S 20O I

I 080 120 160 200 240 280 320 360 TRUE AIRSPEED - KNOTS I

I Figure V-27. ThrustTrue Airspeed,Horsepoweri0,000RequiredFeet. Versus l

I 300-099-003 V-38 00000002-T£Rl13

iL ,

I _I_ 13HI_I_ HELII3OP]'ER _:c_,_l'aN_" I

, /,'_// ,, Ii ORO_IOSS .T-_ _ ./// _2oo } POUNDS j / t i _ ] Too,.4__./O_

9_oo.1 11500'_ /.i/./.,

600 -- 8500\ I-/ ____. "-'- _""

i 400

i STANDARD DAY 200

I --I .... I I I , I 080 120 160 200 2-'_0 280 320 360

I TRUE AIRSPEED - KNOTS

I Figure V-28. VersusThrust HTr_leor3epowerAirspeed,Required20,000 Feet. I

I 300 -099 -003 V-39 00000002-TSR11

I

I 2200

i 2000 __

1800 °_o STANDARD DAY il

Ul

1400 __.,_..-, ,_

_'_" GROSS WEIGHT- =.S.'_:, _ 1200- POUNDS i i I 80o I05009_oo_ 1 ,, / 600 _ 8500 .._.._I I

I 40080 120 16") 200 240 280 320 360

i TRUE AIRSPEED - KNOTS

i Figure V-29. VersusProprotorTrueShaftAirspeed,HorsepowerA_rplaneRequiredMode, Sea Level. I

! 300-099-003 V-40 00000002-TSB12

I (_ BEI .L HEI..ICOPTHR ,:_Mf'^NV

! 2200 -- _' , 2000 -- AIRPLANE MODE 1800 -- STANDARD DAY /

H

_ 1600

_ ig00 GROSS WEIGHT - / POUNDS ! _oo _oo / [ o _ooo ... if500 \ _ 9"9 | _ 10500_--J,, ,_ I \ /, y/ ,I 800 19.500_- _" 8s0o,_ _ /

I 6OO _ j /

I ,,,

I 40080 120 160 200 240 280 320 360 TRUE _IRSPEED - KNOTS I

I Figure V-30. ProproVersustorTrueShaftAirspeed,HorsepowerAirplaneRequiredMode, I0,000 Feet. I

I 300-099-003 V-41 O0000002-TSB13-"

I (1_ IglEl.I I-IEI..Ic()PTER _x_mt,aNf

I 2200 air"

i i I 4 ..... 2000 ..... AIRPLANE MODE ......

I 1800 ----- STANDARD DAY , GROSS WEIGHT- n 1400 POUNDS " ,/// _, 13500

i000 / : _ _/7 I I I 800 9500,. 8500, --/ I | _--- 600 !

I g0080 120 160 200 240 280 320 360 TRUE AIRSPEED - KNOTS !

I Figure V-31. ProprotorVersus TrueShaftAirspeed,HorsepowerAirplaneRequiredMode, 20,000 Feet. !

I 3OO-099-003 V-42 . __ . . • O0000002-TSB14

I 1 I ... I I I

| 24000

z | = _12000

I 8000

t 4000 i 0 80 120 160 200 24C 280 320

TRUE AIRSPEED - KNOTS

Figure V-32. Flight Envelope and Maximum Speed, Airplane Mode.

300-099-003 V-43 00000002-TSC01

I |

(_ BELly HITI .I(:C, PTI_R _=(_I',%N',

i STANDARD DAY AIRPLANE MODE ENGINE RPM

i 30000 R_TED

| I

: " _ 20000 ,_ _o_ _-",l. ,'_ \ ",, , %1

i,,

I m lO000 i \\'\\\\I_\ \ '_' i i o0 \.\_2000 / /4000 t 6000 RATE OF CLIMB - FEET/MIN I Figure V-33. Maximum Rate of Climb, Airplane Mode. ! , I

I 300-099 -003 V -44 - i l _ _ , :+i_i - . . 0NNNNNNg_T_r"no

I _ BI_LL II[ILI(_:OI'T[£1_ _:+_MI'ANY

I .48 •

II "+"11I _+_°°_-I

2 I - ,, '

5i iI| i "+ T_ RDDAY +++ !+';;' ' l

_ + ?" l + "32

I m._+ ++oo<..,5"-_'_ -_+Y I 9500/ /2' / / /

II ._--t 12_°°500

.20 I ..... I 120 160 200 240 280 320

i TRUE AIRSPEED - KNOTS

I Figure V-34. Nautical M_les Per Pound of Fuel Versus True Airspeed, Airplane Mode, Sea Level.

I _0-099-003 V-45 0000000P-T£_n_

1 1

I (I_ F]ELL H F_LICOPTER { x:,_l ,AN_'

o _8

I _ "

i __ STANDARD DAY __

GROSS WEIGHT - ./ I _"_L i i. o ,oo o,,

I _oo.

13500 I

160 200 240 280 360 i .20120 ] I TRUE AIRSPEED - KNOTS 1

Figure V-35. NauticalVersus TrueMilesAirspeed,Per PoundAirplaneof Fuel Mode, i0,000 Feet.

I 300-099-003 V-46 NNNNNAN9 TcrnA

,

(_ BELL HELICOPTER ¢:oMPANV

.52 '_ i _- _ANEMODE-- _-

GROSS WEIGHT - \

I _ POUNDS 8500

o 950C ,

.36 l _ _o / H 1 _ ._ -._T_ o_ / /_ 1 _RQU__M_T_L

i .28 ! I .24 I 120 160 200 240 280 320 TRUE AIRSPEED - KNOTS I

I Figure V-36. VersusNauticalTrueMilesAirspeed,Per PoundAirpolanef Fuel Mode, 20,000 Feet. I

300-099-003 V-47 00000002-TSC05

| r I

I (_ BELL HELICOPTER ComPANY I

STANDARD DAY I I " I I i

I 20000 AIRPLANE MODE ,_/" 200 FEET PER / MINUTE RATE

16000 HELlOO _R 10M DE / OF CLIMB I _ 150 FEET PER /'%_-' f-HOVER _CEILING - I , MINUTE RAT______/ M OF CLIMB k_ / TWIN ENGINE

12000 _ /

8000

bOO0 I

0 I 2000 4000 6000 8000 i0000 12000 i_000 GROSS WEIGHT - POUNDS I

Figure V-37. Single Engine Performance. {

I 300-099-003 V-48 00000002-TSC06

I (_ BELL HELICOPTER c.()MpANY

I 2400 1 ' I I I I I I

i oo-.JI III I 10 PERCENTtl FUEL RESERVE 2o0o 1",. I I [ CREW INCLUDED AS PAYLOAD

1800 1600 I _ 14oo ,I ! _2oo. I' III I _ooo.800 _ I _ _oo I _, " 400 I 200

0 0 I llo0 200501 I 300I I 400I I I0 6OO 7OO

I RANGE IN NAUTICAL MILES ! CRUISE AT VLR C -- 240 KNOTS i', / \ i RANGE ) PAYLOAD

HOVER OGE LAND I 11600 FEET, STD DAY 10% FUEL RESERVE 6400 FEET, 95°F DAY : GROSS WEIGHT = 9500 POUNDS

I F_gure V-38. Research Mission. I :300-099-003 V-49 !

i F. . ,_ .,,2 ...... _ -OOOO0O02-TSCO7

I

(I_ BEl-I. HEI ]COPTER C:O_PANV ..... I I

I CRUISE AT 260 KNOTS 10,000 FEET, STANDARD DAY

I RANGE = 306 NAUTICAL MILES PAYLOAD = 1600 POUNDS

HOVER OGE LAND , 9000 FEET, STD DAY ,10% FUEL RESERVE 4000 FEET, 95°F DAY

_ GROSS WEIGHT = 10300 POUNDS

- Figure V-39. C_vll Mission. I I I CRUISE OUT AT 300 KNOTS

// RETURN VLR C = 250 KNOTS

I i0,000 FEET, STANDARD DAY HOVER 5 MINUTES OGE 9000 FEET, STD DAY

I RADIUSi0 PERCENT= 375FUELNAUTIRESERVECAL MILES 4000 FEET, 95°F DAY PICK UP 3 SURVIVORS HOVER OGE (600 POUNDS) I 61600200 FEET,FEET, 95ST°DF DAYDAY GROSS WEI_{T = ii200 POUNDS I I Figure V-_0. Military Mission.

I 300-099 -003 V- 50 00000002-TSC08--

O BELL HELICOPTER co_f,A_

] sooo

I _ _0 PERCENT FUEL RE3ERVE

4000 _ CREW INCLUDED AS PAYLOAD i \ °3000 \ i oI \ I _° 2000 \ < i m i000 _ V'-MINIMUM PAYLOAD "

I o0 400 800 1200 1600 2000 2400 2800

I _NGE IN NAUTICAL MILES I CRUISE AT VLR C = 275 KNOTS I 20,000 FEET, STANDARD _Y

I RANGE,(SEE ABOVE)PAYLOAD

I 2600HOVER FEET,OGE STD DAY LAND G_SS WEIGHT = 12400 PONDS i0 % FUEL RESERVE I I Figure V-_I. Ferry Mission.

I 300-099-003 V-5[ 00000002-TSC09-

I /_ REI .I- HELICOPTER (:o,V,l'anY

I _- /4000 FT, 9 s°_ PROPROTOR TILTED 30 °

I 1200 LIFT-OFF AT 70 KNOTS u II

i i000

! '-'I 800

? o u_

i ,4_oo

! °_ 400 H ! ° I _.oo !

I 09000 i0000 ii000 12000 13000 14000 GROSS WEIGHT - POUNDS I F{gure V-42. STOL Takeoff Performance. I ! ,

I 300-099-003 V-52 00000002-TSC10

(1_ HEll HEI.IC:OP'I'[_R c:(_ml_Nv

'@ VI. I)YNAMIC STUDIES _.

Dynamic characteristics have been carefully considered {n the design of the Model 300. Particular attention has been paid to the proprotor and flight mode stability in airplane mode. The structural dynamics analysis has also included investigation of airframe vibration and proprotor dynamic loads. Proven analyt- ical methods and computer programs were used for the study. All have been correlated with either flight-test or wind-tunnel test results to verify their accuracy.

. A. Proprotor DTnamics

1. Natural Frequencies

TheBHC proprotor0omputer Programblade naturalC02. 802frequenciesis a Myklestweread-typecalculatedanalysisusingfor a rotating, twisted beam. It includes the coupling between

beamwlseand collectiveand chordwlsepitch. deflectionsGood correlationresultinghas beenfrom achievedbuilt-in withtwist . the measured frequencies of many rotor designs, including three- bladed semi-rigid rotors, such as those of the Model 300. The

wereblade obtainedmass and fromstiffnessthe detaildistributionsdrawings shownreleasedin forFigurethe £II-2fabri- cation of the 25-foot proprotor. Particular care was taken in

I wererepresentingessential thein stiffnessdeterminingof yokethe frequencyand spindleof regionsthe fundamentalas they modes.

I Figures VI-I and VI-2 ahow the calculated natural frequencies as a function of rpm and collective pitch. The frequencies are presented in terms of collective and cyclic modes. The collec-

I revoltlve umodestion areis anthoseintegerexcitedmultipleby airloadsof the numberwhose frequencyof blades per(i.e., 3, 6, and 9 per rev). Nonmultiple harmonic airloads excite the

i cyclicnatural modesfrequencies(i.e., i,shown2, q,in andFigures5 perVI-Irev).and TheVI-2 coupledare noted by crosses (X) and diamonds (_). The crosses denote modes whose largest deflection is normal to the plane of rotation; the

I diamondsrotation. areThethosesolidwhoselineslargestdenote deflectionthe frequencyis inof thean uplanentwistedof blade at zero collective pitch (i.e., uncoupled beamwise and

I chordwiseparing the frequencies)uncoupled andandcouarepledprovidedfrequencies,for reference.the effect Byof com- built-in twist on the blade frequencies is apparent. | Past experience with the design and testlu_ oL three-bladed semi-rigid rotols has shown that frequency placement of the first inplane (cyclic) and second beamwise (collective) modes

I bepose_sufficientlythe most severemovedre reqfromuiremeints.per revTheto firstavoid inplanehigh l-per-revmode must loads but cannot be too high or 2-per-ray loads wilt be a prob-

I lea; a frequet_cy of 1.5 per rev results in the lowest oscillatory ! 300-099-003 VI-I O0000002-TSC11

I loads, llowever, frequencies as low as l..25 per rev are accept- able. Meeting this requirement c_n be difficult _.I] propro_ors _ because of tile wide rpm and collective ra_Ke required for above I efficient operation. The second beam mode must be located 3 per rev to avoid h_gh loads and airframe vibration, rn heli- copters, keeping this mode out of resonance has been a problem;

I however,thicker andit stifferis less bladeof a problemroot sectionswith proprotorsrequired becforausestaticof the strength•

I The frequency location of the major blade modes is as follows: The first inplane mode varies from 1.42 per rev to 1.27 per rev in helicopter and conversion modes (565 rpm), and from 1.44 per rev per I to 1.3 rev in airplane mode.(458 rpm). The second beam mode is 3.7 per rev at high colleetl\e- pitch in helicopter mode and 3.92 per rev in airplane mode. Close proximity to 4-per-rev _ resonance is indicated for the second cyclic mode at high pitch -" • in airplane mode (458 rpm) and 6-per-rev resonance for the third I • , , • r _ collectlve mode at hlgh pltch in hellcopter mode (56_ rpm).

:_ I anticipated. However, should these resonances pose a problem • Veryin wind-tlowunnelairload or exflightcitation testing,in thesetuningresonantweightsharmonicscan be isused to :"'" raise or lower the mode'_ frequency, as required. - i I frequency per TheThis firstmode torslonalis rigid-bodynaturalblade pitching ison locatedthe controlat 4.5_ystem rev.

I stiffness.torsional deformation,The second torsionalis much higher.frequency, which involves blade

i 2. Load_..____s Pa_t Bell studies of proprotor loads have shown that two flight conditions impose the most severe blade loads. For oscillatory

I loads,the mostthesevere;maximum thislevelis alsoflight truairspeede for conventionalin helicopter helicopters.mode is For design limit loads, the maximum results from a gust encounter

I maneuversin airplane in mode.all modesSeveralwere otherexamined.flight Inconditions,all cases theincludingblade loads were less severe than those of the two abovementioned

i flight conditions. Blade loads were calculated using Bell's "Rotoreraft Maneuver Program," Computer Program C-81. In C-81, a trim condition of

I actingthe aircraftat the tsaircrafestablishedt's cg. by Rotorbalancingcollectivethe forceands cyclicand momentpitchs and the aircraft fixed controls are adjusted to obtain the trim

I condition.harmonically Onceanalyzedtrim isand achievedused, withthe thebladebladeair freqloadsuencyare re- sponse, to ce.lculate the bending moments. They are presented in h, rmonle iormat. If desired, a maneuver may be entered and

I loads calculated at specified time intervals during the maneuver. At any time during the maneuver a discrete gust encounter may be slmulateJ• ( with the gust's shape and magnitude spec!lfied as

I required. 300-099-003 VI-2 I 00000002-TSC12

I Figure VI-.3 shows the (-sleulnl:ed hlade oscll" 1 ntnly- - st:ross for Vmax in helicopter mode, sea level, 140 knots TAN, 9500 pounds ...

i grossThe peakwe_.gbt,stress 30-degreeof 22,500 flapspsi, occandurs 0-de_ree_t 60 percentconversion radius;an_l.e, the allowabl.e design endurance stress for the 17-7PII, condOr:ion '. TIII050, steel used for the Model 300 blades is 30,000 psi. The i I principal frequency of the loads is l per rev with a small amount of 2 per rev and 3 per rev. The pea_ stress [n the titanium yoke/spindle is calculated to be 7400 psi; the endurance limlt

I is estimated to be 15,000 psi. The limit blade stress resulting from a 50 foot-per-second vertical gust encounter at VH, zn airplane mode, is shown limit stresses obtained by apply- I in Figure VI-4. The are ing the 80-percent alleviation factor specified for the air- craft gust iced factor determination. The (I- cosine) gust shape

I wasthe alsosuddenexaminedgust withand anproduced80-percentthe alleviationsame magnitudefactor.of loadsThe ascom- pressive buckling stress is also shown in Figure VI-4 as blade

I ativelybuckling assumingis more marginalthat the bladethan abucklingtensile failure.strength isBy equivalentconserv- to the initial buckling stress, it is seen from Figure VI-4 that the blade can carry an ultimate load 1.5 times the limit loads resulting from gust I also a crltical theone 260-knotfor the hub. eneo,mter.The criticalThisstressconditionoccurs _sin the titanium hub yoke, at the junction of the spindle and yoke

I Withring. an Theallowablecalculatedultimatelimit bendingstress stressfor thisof region130,000is psi,89,000the psi. margin of safety is slightly negative.

I Control loads were also calculated. Several methods were used to determine loads in the helicopter mode including sn empirical method based on measured loads for a number of helicopter rotors.

I conservativelyThe results are establishedtabulated below.as 345 ±The345designpounds.pitch-link load was

I MODEL 300:PABLEPITCHVl-iLINK LOADS ,

Conditlon/Method I Steady Oscillatory Vma x - helicopter mode | Calculated 180 ib + 130 Ib ! Empirical i_ 175 ib Model test ± 120 lb

Vm_ _ - ai_p_ane mode Calculated - power on 163 ib

power off 332 1.b - ":.,,, , --.,

300-099-003 VI-3 T

6 ilil i r- _ ...... , _£ "= ...... i In "l Ilin I. • . _l , 00000002-TSC13

i

3. Blade l.l.apllng I o

i • J l,Proprot:orlapp_ng Wasfl.appingcalculatedin airplaneuslng severalmode ismethsummarizedods. Forinsteady-stateFigure VI-5• maneuvers, the mast angle of attack and pitch rate were multi-

i pliedThe gusttimesresponsethe respectivewas calculaflappingted withderivativesComputer Pr(_ogram/@_ and0-81._/@q).

It should be pointed out that the Model 300's flapping sec1_itiv-

I thatity toof anglethe XV-3.of attaThisck inisairplanedue to themodehigheris onlytip aboutspeed 40(600pc,cent,_ps versus 356) in airplane mode• The rotor following time is also

I less2.1 fosincer the theXV-3.Model 300 blade lock number is 3.76 compared to

i Flappingthe 12 degreesin helicopterallowed andby theconversionmechanicalmodesflappingis alwayssteps.less than

B. Airframe Dynamic _

i. Natural Fre,quencies

Placementguided by oftwo theconsiderations:wing-pylon-fuselagefirst, naturalresonancefrequenciesof the coupledhas been airframe nodes with proprotor i, 3, and 6 per rev must be avoided •_ to have satisfactory vibration charac_:eristics; second, the frequencies must be adequately separated to avoid aeroelastie instability. For the Model 300 the wing beam and chord and the fuselage bending stiffness resulting from strength requirements

I Hprovideowever, forthe satisfactorywing torsionallocationstiffnessof thesefor strfundengatmentalh requirmodes.ement resulted in a win_ torsion resonance at airplane mode rpm. The wing torsional stzffness was increased by 60 percent to provide I adequate separation from l-per-rev resonance. The resulting frequency locations provide good vibration isolation and as a result of the torsionally stiff wing, a high level of proprotor- I pylon stability. The symmetric and asymmetric airframe natural frequencies are

I an$1eshown. in TheFiguresrange VI-6of freqanduencyVI-7 ofas thea functionsymmetricof modespylon withconversiongross weight is indicated by shadings; the higher frequency correspond-

maximum gross weight. The asymmetric modes do not vary signif- i ing to minimum operating weight and the lower to the 12,400-pound icantly with gross weight. The frequencies shows are for the 9500-pound design gross weight.

I Two resonance conditions are passed through during conversion. The asymmetric wing chord mode is in l-per-rev resonance at

• I pfaurtiselageal coresponsenversion toangles.hub shears.This modeAlso,hastherelasecondtively winglow beam modes (symmetric and as_nmetric) are in resonance at 3 per rev for certain fuel loadings and conversion angles• These modes |

I I 300-099--003 VI-g i II i i I 00000002-TSC14

I o °°ELL hub shears, they are a_so damped by fuel sloshing under part:_al. fuel loading conditions and by aerodynamic damping, The pylon

yaw natural frequency Is above .4 par rev..

empennage.A preliminary Thefrequencycantileveranalysisfrequencieshas a_sofor beenthe mfinade andfor tailthe plane are tabulated below: (The tall plane mass and inertia £n

i the fin calculation.) TABLE VI-2

EMPENNAGE NATURAL FREQUENCIES I i lstf!nbending I / 1st fzn torszon [ 27.0 CpS

. The airframe natural frequencies were calculated using two1 I analyses. Program A75D, a state-vector, crossed beam analysis was used to determine the couoled fuselage-wing-pylon frequen- cies. A75D includes coupling between beam and torsion deflec-

i tcorrelatedlons, and shearwith shakedeformateststion ofandtherotaryXV-3 inerconvertiplanetia. It hasand been several helicopters. The empennage frequencies were calculated

withmass Programand stiffnessDF1789,distribuwhich tionsuses theusedfinitein theelementairframemethod•vibrationThe analysis are shown in Figures III-3 through III-6.

i 2. Vibration Levels Estimated 3-per-rev vibration levels for helicoptez,, conversion,

I andFigureairplaneVI-8. modesFor theare helicopshown terfor athend conversionpylon and crewmodesstationthe v_bra-in tion level was derived by using calculated hub shears from the

i bladeThe callcouadlatedcalcu3-per-raylations andvertictheal airfrersameponsefreqatuencythe pylonresponse.and crew station and airplane modes is shown in F_gures VI-9 and VI-IO respectlvely.

I The airplane mode vibration level was estimated by using pylon 3-per-rev vibrations measured in the August wind-tunnel test of

i tdirehe cMtlyodel from300 aterheoemealastisurced model.data (theThe modelpylon ivibs Froratiudeon scaled).is scaled The crew station level was determined by multiplying the ratio of the response level at the crew station to that of the pylon, pylon I times the measured acceleration.

I 300-099-003 Vl-5 00000002-TSD01

| ! 7 -

I (_ BELl_ H_L|COPTER I_oml-anY

I C. Stabi] it_ _,

I I. Proprotor The high wing tor_Lonal stiffness, resulting from the require-

I mentof proprotorfor avoiding/pylon wingstabilitytorsionalfor resonance,the Model 300.providesOthera highfactorslevel are the moderate value of pitch-flap coupling and the hub restraint. A summary of the proprotoP stability is given in I Figure VI-II. At sea level the margin is over 170 percent of the 350-knot dive

I airspeed,requirement•V D, farFor referencein exeess theof theairspeed120 percentabove whichV D fluttheter-freeblade tip helical Mach number exceeds 0.9 is shown. In the following

i paragraphs300 proprotorthe stabilityanalytical boundarymethods andusedtheto sensitivitydetermine theto Modelrotor rpm and wing stiffness are discussed.

I Proprotor/pylon stability eharaeteristies were determined with I BHC Computer Program DYN4. DYN4 is a linear, twenty-one degree- _, of-freedom proprotor stability analysis. It is capable of deter- oo.oo path-plane representation is used for the proprotor dynamics and

I_ i stabililinear tyaerocharadynamiccterisfticsunctionsof tilt(eL,-prCoD)protareor assaircumedr.aft. A tip- Ribners' method (Reference 34) for propellers is used to correct for compressibility effects. Rotor details such as pitch-axis

I restraintpreconing, areunderslinging,included. Thepitch-flapfirst inplanecouplingbladeand modeflappingis repre- sented and control system flexibility may be simulated quasi-

I statically.sented; wing Fivebeam, coupledchord, wingtorsion,/pylon pylelasticon pitchmodesand areyaw. repre-The six fuselage rigid body degrees of freedom are included which

i allowsaircraft theflightsymmetricmodes andto beasymmetricsimulated.free-freeInput modesto DYN4andisthein terms of lumped parameters describing the dimensions, inertias, stiffnessas, and klnematios of the aircraft being simulated.

I Standardthe airframeaircraftaerodynamics.stability Outputderivativesof DYN4are isinputin termsto representof the system eigenvalues and eigenvectors. DYN4 has been extensively

I ccorrelaorrelatedtion withis good,small-sevencale formodelcompletestx aeromechanicalstability data.typeThes of instability. An example of the degree of correlation is shown in Figures VI-12 and Vl-13• The measured data in Figure Vl-12 are proprotor I from a joint NASA-BeI_ test of the Model 266 aeroelastic model, that of Figure VI-13 are from the August tests of the Model 300 aeroelastic model. The Model 300 model is shown in

I FiguTure Vl-15VI-14 showsin the thesemispsnfull-spanconfigmodelurationas ittestedwill bein moAuugustnted. in the NASA Langley [6-foot transonic dynamics wind tunnel next

i sprlng.

I 300-099-003 VI-6 00000002-TSD0P

I Two o1:her programs are used to investigate areas where DYN4 lacks capability. These are Programs DYN3 and DYNS. Program

I flutterDYN3 combinesanalysis.the proprotorThis permitsrepresentthe ainfluencetion of ofDYN4thewitproprotorh a wing on wing flutter characteristics to be determined, as well as

I tpylono investigatestability. the Belleffectstudiesof wingof aerodynthese aeffectsmics onhavethe shownproprotonlyor/ small mutual effects of flutter on proprotor/pyLon stability and vice versa. DYN5 is a nonlinear ope_-form proprotor aero- I elastic analysis. It uses strip theory and a tabular repre- sentation of the blades aerodynamic functions (eL, CD, and Cm) to account for stall and compressibility. Tables for different

I tionprofilesfrom mayrootbe tousetiN.d to Oorrelationaccount for ofvariationDYN4 andinDYN5the bladeis shownsec- in Figure VI-16, and indicates only a small difference in the stability boundary. DYN5 does indicate a slightly lower level I of damping than DYN4.

Figures VI-17 and VI-18 show the root loci of the Model 300 I symmetric and asymmetric modes as a function of airspeed. Note I that the symmetric wing chord mode becomes unstable first. In .,,_ calculating these root loci the airframe structural damping was • assumed to be zero. This is conservative since i to i-1/2 per-

lift curve slope was corrected for compressibility for airspeeds cent of critical is inherent in the wing structure. The blade i Thus,up to 350the knotsstabilityand theboundary350-knotshownvaluein Figureused atVI-IIthe higheris represent-speeds. tative of the stability margin at the 350 _not dive speed.

I The Model 300 proprotor stability is not sensltiv_ to rpm as evident in Figure VI-19. Also, a very large loss of wing struc- tural stiffness or pylon attachment stlffness can occur before

I alsoinstabilityshown inwouldFigureoccuVI-19.r insideThetheprincipalflight envelope.dynamic problemThis isthat would result from the rotor rpm exceeding the ±I0 percent rpm

i limitresonanceor fromwhichlosscoulofd strbe uavoidecturald stiffnessby changingwouldrotorbe rapm.l-per-zev Blade motion stability for the Model 300 proprotor is assured

I characteristics.by the selection ofTherotfiror stparameinplaneters frtehatquencyprovideis abovestableoperating speed which eliminates mechanical instability (ground resonance).

I Theweavingbladewillis massnot occbalancedur. Thesuchbladethateffecpitch-flaptive centfluter tofer gravit[or is at 24-percent chord with the effective aerodynamic center In hover being at 26.1 percent. The 2-i/2 degrees of pitch axis system prevent I preconing and a stiff control are used to pitch- lag instability. Positive plteh-flap eouplit,g of 0.268 (_3 = -15 °) is used to prevent flap-lag instability in airplane mode.

I 2. Flight Mode Stability Stability characteristics in airplane mode were determined over

I a speed range from 150 knots to 350 knots and from sea level to

I 300 -099-003 VI 7

* =- ..... , 00000002-TSD03

| ,. 20,(100 feet. The predicted stability characteristics are accept- _. able in accordonce with MIL-F-8785(ASG).

I The fin and talkplane of the Model 300 have been sized conserv- atively to over compensate for the destabilizing influence of the i proprotors. The flight-mode stability characteristics were calculated using Computer Program DYNd. The equations of motion for the basic I airframe (less rotors) are those suggested by Etkin in Reference 35. The proprotor influence on the flight-mode stability is accounted for directly by using DYNd. This permits the coupling

I treatedeffects simfromultaneously.the proprotor-pylon-wing and flight modes to all be

I Thefrom stabilitythe wind-tunnelderivativestest forof thethe Modelbasic 300airframei/5th werescale obtainedforce and moment model. Where necessary, estimates were made using the methods suggested in NASA TN-1098. The derivatives were corrected I for Math number effects using the Prantle-Glauret rule. a. Longitudina ! Modes

The root loci of the short period mode, as a function of airspeed and altitude, is shown in Figure VI-20. That of the basic air-

I frame, less rotors, is shown in Figure VI-21 for comparison. Analysis of the phugoid mode was made using Etk_ns method as shown in Reference 35, The frequency and damping are given by !

_n = CLo and _= t CDo I _# CL 0

The time to half, conservatively estimated by assuming the prop- rotor zero, presented I forces to be is in tabular form below: TABLE VI-3

I PHUC,OID MODE CHARACTERISTICS I GW = 9500 ib Flaps up

I CGRotorlocationrpm = 294458 (aft) MastNASA anglestandard= 90day° Speed Altitude Period Time to u/8

I (kt)150 Sea(ft)level.... (see)35.1 Half52.0(Sec) 0.Magn706 Phase-94.2(de_). 250 Sea level 58.6 59.1 0.706 -96.2 I 350 Sea level 83.5 42.3 0.704 -I02.2 250 20000 58.4 100.6 0.706 -96.2

I & ,350, 20000 81.8 79.2 0.708 -96.5

I 300-099-003 VI-8 00000002-TSD04

I

b. Lateral-Directional Modes .

I andThe altituderoot loci isof shownthe Dutchin Figureroll mode,VI-22. asThata functionof the ofbasicairspeedair- frame, less rotors, is shown in Figure VI-23 for comparison.

I Notethe basicthat atairframe.low airspeedsThis istheduedampingto rotoris higherthrust thandamping.that Atof higher speed the thrust damping is less than the negative damping contribution of the proprotor side force and the Dutch roll damp- therefore I ing is lower than that of the basic airframe. Mode shape and phase data for the Dutch roll at selected air-

I speeds and altitudes are presented below: TABLE VI-4

I DUTCH ROLL MODE SHAPE AND PHASE

._ Speed Al£itude Period Half Magn Phase Magn Phase i 9 (kt) (ft) (see) (See) (deg_ (de_) | Time to I -_ 150 Sea level 4.73 0 885 1.187 3.7 1 190 188.7

350250 Sea levellevel 1.662.58 0.6550.741 11.820.475 149.0.4 11.740.449 204.9198.2 I 250 20000 3.54 1.740 1.778 4.2 1.720 189.5 350 20000 2.32 1.680 1.190 7.2 2.070 1.94.9

I ,,, , ,,

_! i The times to half for the spiral and rolling modes are tabulated :_ I below. For these modes the proprotor provides an increase in .+_ stability.

I TABLE VI- 5 TIME TO HALF FOR ROLLING AND SPIRAL MODES

! -- Time to Half Amplitude

I Speed Altit(ft)ude Spir(sec)al Mode Rolling(see) Mode

I 250150 Sea level 11.405.96 0.7730.473 350 Sea level 147.00 0.363 I 250 20000 6.90 0.885 350 20000 10.95 0.695 | ,

I 300-099-003 VI-9 00000002-TSD05

I Figure VI-I. NProatuprarotorl FreCoqulleencitives.e Mode I I

- I 300-099-003 VI-IO nnnnnnno_

_I_ ) BI_.I_L FIFLI¢OPTER (;OMPANY i O0000002-TRI")N7

1 I

[ @ BI=LL H EUCOPTER c°MPAN_"

SEA LEVEL , 9500 LB GW i 140 KNOTS TAS -- FI,APS - 30 ° _ ; CONVERSION ANGLE - 0 O

"! . . 28

DESIGN MAXIMUM _i 24 ALLOWABLE 30 KSI

50 20 ""2 i

¢4+.(D 8 / /

4

0_ 0.2 0.4 0.6 0.8 1.0 BLADE SPANWISE LOCATION r/R

Figure VI-3. Blade Oscillatory Stress at VH, Helicopter Mode. I

I 300-099-003 VI-t2 i i r, - 00000002-TSD08

I

0 BELl-- HELICOPTER cOmPANY I SEA LEVEL 160 9500 LB 260 KNOYS TAS FLAP S 0 o CONVERSION ANOLE - )0° I i.o 1 i

I' 120 X <

R VE BUCKLING " _ _/. STRESS -:":_i:I ' ' _ 8o _ \¢D

• 60 ? 50 _/SEO SUDDENOUST J' • 7 (W/O ALLEVIATION)

-!,, _ LTMIT STRESS

40 7 '_ (80%/ ALLEVIATION FACTOR)

20

0 0.2 0.4 0.6 0.8 [.0 BLADE SPANWISE LOCATION r/R

Figure VI-_. Blade Limit Stress at VH, Airplane Mode.

} 300-099-003 Vl-13 00000002-TSD09

(I_ BELL HI_'-'LICOPTER (.;OR_IPANY

12/////////////_-"Ae'K_N-OS.'roe.s,/////. /C,/////,4

i00 150 200 250 300 350

AIRSPEED - KNOTS

Figure VI-5. Flapping Envelope in Airplane Mode.

I 300-0_9-003 VI-I_ 00000002-TSD10

I

l 30"

' 20

• _ 15 ...... IST TORSION ?- FUSELAGE VERTICAL BENDING

to l_

t 5 _IST WING B .

IST BEAM

_ 0 ° CONVERSION ANGLE 90 °

i Figure Vl-6. NAirfraturalame FrequSymmetricencies. Mode J

300-099-003 Vl-15 00000002-TSDll

I

I _ BEI,I.HELICOPTER oCtroI-aNY I ' I _.-"

,! 25

J 20

, _

m 15

IST B_

I 0 ° CONVERSION AN_ 90 °

Figure VI-7. AirframeNatural FreAsy_etriequencies. Mode "I

300-009-003 VI-I6 00000002-TSD12

J

£.o .... I

I .9.8 I ...... i PYLON VERTICAL .7-- _

II P I ->.6 7-- HEf =Ug28.OPTER2 C_,ODEPS -- _-AIf_NE= 22.8 MODECPS I I __ >°.. /i _'_ 7

.4 0 _: 0 100 200 300

I _ I I I I I I l CREW STATION VERTICAL .M._L-A:_8._9_AS_L_!MIT.G.) ..,.,.

. HELICOPTER MODE L ANE MODE I I I ! .06 i--f = 28.2 CP_ I f = 2.2.8_CPS > .04 I "_--

.o2 ,. i-

I O0 100 200 300 AIRSPEED - KNOTS ! Figure Vl-8. Pylon and Crew Station Vibration. I

-o I

I 300-099-003 VI-17 00000002-TSD13

I ...... %. ii i_I...:,:,:":I....ill! MOED f

i SYMMETRIC I WING BEAM i0.0 2 WING TORSION

i 3 FUSELAGEBENDING VERTICAL 4 2ND WING BEAM I °_ 2 O O I "-.° 1.0

" |

MODE I iO. 0 i WING BEAM

I 2 WING TORSION 3 2ND BEAM

O O I

I

0.1 I 0 5 iO 15 20 25 30 35 FREQUENCY - CPS

I Figure Vl-9. Pylon and Crew Station Frequency Response, Ilelicopter Mode. I

I 300-099-003 V1-18 O0000002-TS D14

• ..... _- I - _ II ! III _:T iu

( O0000002-TSE01

I (1_ BF_LL HI_LICOPTER (;OMPANY

I 4. I I ' I VD I _o _o., I

II /_ I PROPROTOR /, STABILITY _il _ _ , _o_ _i- _ _v_o,E I , f_R = 600 W _i0 /'// // , ::,'ii :?,z/>, I /';/-/- /,'/. /,

I 0 0 600 800 TRUE - KNOTS !

i I Figure Vl-ll. Model 300 Stability Boundary.

I 300-009-003 00000002-T.qI::N9

1 j

I ii..

I 6 Q® O WING CHORD DAMPING 5 l I _-_ _°----°-°-°

3 DATA FROM RUN i O ' NASA/BELL TEST. I _ 2' MODEL 266 SEMI-SPAN Q MODEL 815 RPM I _ ._.,_,_

--CALCULATED '':_ :" I 0 _ ioo l_o 200 250 I 6

I 5 WING BF_M DAMPING

| s @i_. 0...--" N I _ ,_o

II _ o " ,,o\ %o o _ l , \- '_ I 100 i50 "\ 2"00 250 VELOCITY - FT/SF_C ! FiEure VI-12. Correlation of Calculated and Measured

I ModelProprotor266 AeroelStabiliasttyic - Model.NASA-Bell Test of I

I 300-009-003 V1-21 ONNNNNNg_T_n_

| I

...... I • ...... I-- I I I=I T "- ]Ii it, .... , - • __ I 00000002-T£FNR

+.+ i I I @ B_LL.EL,COP+_o+A++ I 3

I ++, "'"_

--DIN 4 (LINEAR DYNAMICS/AERODYNAMICS) ' ------DYN 5 (NONLINEAR DYNAMICS/AERODYNAMICS) I f_R : 600 CANTILEVERED WING (MODEL) _ I , . _ . , , _ I00 200 300 400 500 , ! ' o ",,/ || i ',"',/ 0 i00 200 300 400 500 I TRUE AIRSPEED - KNOTS

I Figure VI-16. Comparison of Stability Analysis Method. I

I 300-009-003 VI-25 00000002-TSE07

I (_ BEI..I. HELICOPTER COMPANY

I 230 ._- I

YAW

I 9500SEA LEVELGW I _o

80 PROPROTOR

• _, 60 WING I TORSION | 4o

WING I _oo250_oo350---r_.450oo_o,650,_oKNOTS CHORD

PROPROTOR I _--'_Ct " 0 -16 -12 -8 _I 4 DAMPING- RAD/SEC I Figure VI-17. Root Locus of Symmetric

Free-Free Modes. I

300-009-O03 VI-26 O0000002-TRFO8

I (_ BELL HELICOPTER COMPANY

I 230 ,,..

YAW I 210 • QR = 458

i S9EA5OOLEVGWEL

I 190 1 PROPRoTOR I 2 ! o • _, 80

,t :'h- [ f,,WING TORSION .,i

1 202so,as.o__0 300 4oo'_o:_6.7so__os°_:oTs w_G_

2O PROPROTOR %

-16 -12 -8 -4 0 4 DAMPING - RAD/SEC

Figure Vl-18. Root Locus of Asymmetric Free-Free Modes.

300-009-003 VI-27 00000OO9-T._PnQ

BELL HELICOPTER COMPANY

FREE-FREE

80 -- CSEOMPRESSIBILITYA LEVEL INCLUDED

ASYMMETRIC

60 CD m MODES _2 ,..._ SYMMETRIC

) , 20 _ Z r

4_'_ 700

600 _"

//j /// _/'/j"/j ASYMMETRIC _ ___

RANGE '// o_ _ _ODES 4OO ///'/z// 7//////I//_ '/A ' _, ---- SYMMETRIC I

300 0 100 200 330 400 500 600 700 AIRSPEED - KNOTS

Figure VI-19, Proprotor Stability Sensitivity i to Wing Stiffness and RPM.

300-009-003 VI-28 O00OOOn_-T.qmtn

I (_ BELL HELICOPTER CUMpANY | r ! 8.0 ! I i | ,.o MIL-F-8785 I V I \ SEA "I 350 KT , _ /- LEVEL --- I

_D

I _.C _ 250 _, ":k_, I _ 200

, _oooo_ / '_'\__ ! 20000 FT \ \ \ -7.0 -6.0 -5.0 -_.0 -3.0 -2.0 -1.0 0

DAMPING - RAD/SEC

Figure Vl-20. Root Locus of Longitudinal Short-Period Mode.

V1-29

300-099-003 00000002TSEII

I

I (_B_,'"E''COPT_oM,A.y i@ I MIL-F-8785 I 8.o V/ (ASO)

6.0 _ ""_,, _ 10000. FT i 300 "__ _', -- 20000 FT

250 %.

0, 2ooX", Z i % I i_.o \_, \ I " ,_o%{X\ i 2.o _'_\ I \ 1,0 k ! ..... \ \

I -7.0 -6.0 -5.0 -_.0 -3.0 -2.0 -i.0 0 i DAMPING - RAD/SEC

Figure VI-2i. Root Locus of Longitudinal Short- I Period Mode of Basic Airframe. I

I 300-099-003 VI-30 00000002-TSE12

(_) BELl- HELICOPTElY r:(),,,,,IF,Ar,,',,,

- -- I II,

_,0 L

I 7

!_I 6.0 \ I \ /_AsG_MIL-F_8785 - "_ 5.0 ' ? I _ 300 ""_-., , \ I _ _.o . ...._.,, -_.,/-_oooo

I 2.o 2oo

1.0 _ .

o \ -)..4 -]..2 -1.0 -0o8 -0.6 -O.P, -0.2 0 DAMPING - RAD/SEC

FiEure VI-22. Root Locus of Dutch-Roll Mode.

300-099-003 Vl-3t (")NNNNNNg_T_ _ t

I _1_ 8F-I .I- HEI-ICOPTIER _:O_IPANY I g, ,

8,0 I

J 7.0 •NIL-F-8785 I \/-c_o_ --

6°0 - ! ,, ° \ _.M_I_ o_ .0 350 KT ,W /---SEA LEVEL \ _ _ . _ r _oo0o_\ "_: - _ 4.0 -- _oo\,"'kq.. _. _ 20000

3.0 2SC \ I

200 ,.._ _.o _50\_' ,.X. \ 1 "_ 1.0 I I \ I -1.4 -i,2 -i.O -0.8 -0,6 -0._ -0.2 O

i DAMPING - RAD/SEO

Figure VI-23. Root Locus of Dutch-Roll I Mode for Basic Airframe. I

I 300-099-003 VI-32 O0000002-TSE] 4

__ HRLICOPTIER c_Mr-AN_"

VII. NOISE ,.

An analysis of the merits of any particular VTOL design must include the conslderstlon of noise radiated into communities and areas adjauent to VTOL sites. Close to such sites, the community noise problem will probably be worse than that experienced at our busiest airports today. Three factors tend to make the small downtown or suburban heliport more of a community noise problem than conventional airports: the relatively high power settings necessary duning landing and takeoff, the small separation be- tween the aircraft and exposed communities, and the low ambient noise environments at many of these locations.

Although considerable amount of experimental data and prediction methods exist for analyzing noise of fixed-wing aircraft and of conventional helicopters, the complex noise eharacterlstics of future VTOL aircraft cannot be fully described for many of the less orthodox propulsive systems. However, sufficient informa-

il_ " tnoiseion excharacteristicsists to permit oaf reasonableindividual escomponentstimate to beof madeany givenbased on : propulsive system, i.e., propellers, jet engines, rotors, etc. i _.._ [ A measure of noise called the perceived noise level, which is

by acoustical instrumentation to the subjective impressions i_ epeoplexpressedexperiencein units whenof PNdb,they arerelatexposedes the mtoeassouruend.ments Thisrecordedmeasure rates the annoyance or noisiness of complex sounds and is used in this country and abroad for subjective-judgment studies of traffic, industrial and aircraft noise. Perceived noise level takes into account the distribution of sound energy over the audible frequency range and gives a direct and simplified measure of the subjective noisiness of different VTOL aircraft.

Perceived noise levels calculated for the Model 300 are compared in Figure VII-I with estimated levels of VTOL aircraft proposed for (60-passenger) short.-haul service and with measured levels of present-day helicopters. These curves show that rotor-type aircraft will be least noisy, while jet types will be the noisiest. The noise of the Model 300 will be 20 PNdb Less than that for tilt-wing VTOL and 40 PNdb less than that for jet-lift aircraft. The Mode]. 300 will also be quieter than conventional helicopters because it will not require a tail rotor, whose sound contributes £o a helicnpter's perceived noise. The maximum perceived noise level of the Model 300 in helicopter mode will be 93 PNdb at a distance of 300 feet.

Figure VII-1 also shows the range of perceived noise levels for heavy industrial areas. At _ distance of I000 feet, the noise of a hovering Model 300 will be no greater than that generated

300-099 -003 Vll-1 O0000002-TSFO]

I_ BELL HELICOPTER c;OMPAN¥

in heavy industrial areas. In addition, the Model 300 will be " quieter when in airplane mode because the proprotor_s tip speeds are substantially reduced. Cruising at an altitude of i000 feet the noise of the Model 300 will be approximately 65 PNdb, a level comparable to ambients measured in commercial areas with light traffic. The Model 300 will not be loud enough to be heard in busier areas.

J i I,

- |

300-099 -003 Vl1-2 i

EIHI I- ttEI_ICOPTER c:o,'_l "ANY 0, O000000P-T.£ Ffl."4

• i I

Vlll RECOMMENDED PROOF-OF-CONCEPT PROGRAM

A. Objective and Purpose

The purpose of the tilt-preprotor proof-of-concept a_rcraft program is to extend and evaluate technology for this prom_s_ng concept and to determine its suitability for future development and service. The ultimata objective is to enable Industry, militany and civil planners to proceed in a straightforward manner to develop operational aircraft and VTOL transportation systems.

The technical results from the program will establish that the concept is technically feasible and that technology is adequate for the successful development of operational aircraft. In addition to flight research to obtain technical data, tests will be conducted to evaluate economic feasibility, battlefield acceptability and social acceptance for a variety of civil and military missions. Measured performance on simulated missions _ / will establish economic feasibility for the concept and will ___ . provide basic data to establish the cost effectiveness of this type of aircraft in a VTOL transportation system. Measurement --' of environmental effects on the aircraft and the alrcraft's safety, noise and other characteristics will determine its acceptance to society and its ability to operate successfully in differing military roles and battlefield conditions.

B. [rogram Plan

The recommended program plan for the tilt-proprotor aircraft proof of concept is presented and discussed in the following -, paragraphs. This plan is based on the s0ccessful complet{on of the NASA tilt-proprotor.technology program and assumes that _ the resulting technology will be available for the design of the proof-of-concept aircraft. It is also assumed that the aircraft would be _ proof-of-concept the same or very similar to mm the deslgn presented in this report.

This same plan would be applicable for a folding proprotor air- craft. The design study for a folding proprotor proof-of-concept aircraft is in progress at BHC under a separate NASA contract, NAS 2-5461. The basis of that study is to utilize the Model 300 tilt-proprotor aircraft airframe with modiflcations as required to install the stop-fold provisions and the cruise propulsion system. Because these concepts are closely related, " these programs could be combined. This is discussed _n more detail under program schedule.

Thethreeprogramdistinctivefor theelementstilt-proprotorof work candan/orlogicallyphases. be Thesedividedare:into !

300-099-003 VIII-I 00000002-T,qFO4

I I BELl_ HELICOPTER COMPANY - Phase I - Design and Fabrication of Aircraft ""

I Phase II - Flightworthiness Tests

Phase III - Proof-of-Concept Fllgbt Resesrch

The tasks to be covered in each phase are outlined below.

I i. Phase I - Design and Fabrication a. Design

I The aircraft design approach will utilize the latest available technology ih terms of the concept, materials, design, systems, engines, etc. However, items not fully developed or proven _:_. _ will be avoided as it is not the purpose here to develop sup- l porting technologies. For instance, the development of engines or new material applications is not intended. The design will

I representative of the state of the art. Close weight control will be maintained in the design stages so that the resulting

i empty weight to gross weight ratio of the proof-of-concept air- operational .aircraft. The aircraft design criteria will be • selected to permit the aircraft to explore safely the extremes craft will be indicative of useful load ratios for future I maneof theuver airandcrafttdives condiperfotrionmansce shouldenvelopep.ermitIn thoaddiroughtion, flightdesigntest- ing to investigate the technical problem areas of the concept.

I b. Fabrication

Fabrication of the aircraft, test items, and spares will be done with prototype or soft type tooling except for components such as blades and transmissions where hard tooling will be required to maintain quality, Quantities of components to be manufac-

I Phasetured wIIillandbe IIIestablisheand tod provideto _atiefyadequatethe sparestest requiremto supportents ofthese programs.

i 2. Phase II - Fllghtworthlness Tests

a. Component Qualification Tests

Critical components which are not off the shelf and have been designed or repackaged for the proof-of-concept aircraft will undergo tests to qualify them for flight research. This will include functional tests, proof tests, load-cycllng tests, liTe tests, etc. for such components as conversion, flap and control phasing actuators, hydraulic boost cylinders and ejection seat installation. Drive system bench tests will include engineering development testlng.as well as green running of the components for the flightworthlness and fllght-test articles.

300-099-003 VIII-2 O0000002-TSF05

I

I I_ B_ |-L H E LICOPTER _°_r'^mY

b. .Fatigue_.Tests _ A mznlmum amount of fatigue testing will be conducted on critical blade, hub and proprotor control components to establish a safe

i specim_llfe for-ns at.lewillastbe 200testedhourofs eachof fllghtcomponent.testing.Two Typiwillcally,be testedfour prior to flight based on estimated loads and the remaining two after loads have been measured in exploratory flight test. I Fatigue failures will be used to establish S-N curves and a safe test llfe. .

I c. Propulsion System Tests _. Development and enduran?e testing will be accomplished in a ! • ground test rlg to @ual_fy the prnpulsion and related.systems i_ I for flight. This will Include the proprotors, the drive system i. with interconnect, powerplant installations, ?onversion actua- tors, and the associated h[draullc and meehanlcal control systems _ _

I donefor thwithe proaproflighttors artlcle,and.englnes.but is.Thissometlmestype of donetestingon ancan.be"iron bird". A combination of flight article and iron bird is recom- '_' • mended for th_s program. All components and systems will be _q I designed and fabricated as flightworthy systems and installed in a flightworthy wing. The complete package of wing and systems will comprise a flightworthiness test article. The test article

thewill prbeoprotorssupportedin propellein a testr modestand aswhichwell willas inpermithelicopteroperatingmode. After completing functional and operational checks of all systems,

arunning150-hourat qualificspecified lioncombinatirun owillns ofbe made.power, Thisrpm, ma;twill angleconsist andof control setting. Some of the time will be with only one engine

operative to qualify the interconnect drive system. d. Full-Scale_Wind-Tunnel Tests

I Wind-tunnel tests will be conducted in the NASA-Ames Full-Scale Wind Tunnel to identify deficient areas and establish a safe flight envelope (within the speed range of the tunnel) thereby permitting a more rapid and safe flight test program in the helicopter, conversion and low airplane speed ranges. Rather than use the flight-test article, it is ;ecommended that the flightworthiness test article be used for these tunnel tests. The same remote control and instrumentation systems used for the flightworthiness ground tests would be used in the tunnel tests. The hard points installed on the wing could be designed so that they would also serve to mount the test article in the tunnel. The flightworthlness article is shown in Figure VIII-I as it would appear mounted in the 40-by-B0-foot wind tunnel.

Tunnel tests wlll be conducted to determi_le that the steady and oscillatory proprotor and control system loads are within design limits. Aircraft forces and moments and control positions will be recorded aud analyzed to determi(Je if the aircraft's flight

399-099-o,'13 VI [i-3 O0000002-TSF06

| BELL characteristics are safe and that the conversion control phaslngr r.

I Boundariesrelationshipsof tdehesignedconversioninto thcorridore controlwillsystembe exareploredcorrecandt. established by testing at various combinations of wing angle of

i attack,stall andconversionoscillatoryangleloadsand willpower.be usedWing tobuffetestablishthe rotor-bladelimitations and boundaries for safe conversion.

I e. Aircraft Ground Tests The flight aircraft will be thoroughly checked to determine

I andpropermechanicalfunctioningsystems..and operationThe controlof itsyss telectrical,em will be hydraulicrigged and proof loaded. A ground vibration frequency survey will be

i conductedframe modes.to deterIf anymine frequenciesthe frequencyare locationfound to ofbe thesignificantlymajor air- different from predictions, stability analyses will be redone to evaluate these differences and to determine if any corrective action is necessary. 1 The aircraft will be run on tledown to further check all systems i and to check the dynamic components. Before releasing the air- | craft for flight, a minimum of fifteen hours of running time will

be accomplished. This will include running at various combina- i tions of conversion angle, control position and power settings. f. Exploratory Fli_ht Test

N andExploratorywill be guidedflight byteststhe willresulbets conductedof the wind-tunnelin a cautioustest mannepro-r grams. Conversion would not be attempted prior to conversion

N teststhe wind-in ttheunneltunnel.envelopeExpawillnsion be ofdonethe inflighta bu_id-upenvelopemannerbeyondby extrapolating load and stability data to the next speed increment prior to flight at that speed. Analysis of telemetered data will flight. I also be used to monitor each Inflight shakers on the aircraft and appropriate controls will be utilized to determine damping of the proprotor, pylon, wing and empennage modes of

I verifyvibration.properSufficientengine operation.powerplant measurements will be taken to

I Acompletipreliminaron yof flightthe fatiglouaeds tessutrveofy willthe dybenamiconductedc componenttso permitand establishment of a safe component llfe for the fllght-research program.

I A safe nominal flight envelope in terms of altitude, gross weight, speed and load factor will be established. As necessary, adjust-

U mthie nts s envelopeor modifiwcailltionbes suffiwillcientlbe madey largeto theto airdemoncraftstratesuch thethat performance and general characteristics of the aircraft. Upon • completion of the exploratory fllght-test program, the aircraft ! proprotors and transmission will be disassembled and inspected.

I 300-099-003 Vlll-g O0000002-TSF07

BEt.L HELICOPTER cOMPANY

i Parts will be replaced as requ__red and the aircraft reassembled _ and readied__for proof-of-concept flight research.

3. Phase III- Proof-of-Concept Flight Research

I The test program is directed at establishing proof of concept in three areas; technical proof of concept, economic proof of concept, and social proof of concept. Dividing the test program

I proceedinto phaseswith wotheuld emphasisnot be rebeingquired,placedbut onthe theeffortechnicalt will logicallyareas early in the program with emphasis switching later to tests

directedmental accepat testablishingability. Theecofollowingnomic feasibiliparagraphsty andbrieflyenviron- discuss the type of testing that is anticipated in the three

proof-of-concept areas ...... a. Technical Proof of Concept

I areasThe aircwillraft'bes invepersftigatedormance, andflightevaluachtedaracttoeristdeteicsrmineand ifproblem = technology is in hand for the successful development of an

I operaEngineeringtional aircrflightaft teswitsth willdesirbeableconductechnicalted to characdocumentteristhetics. aircraft's flight-handling characteristics, proprotor and air- frame dynamic stability, airframe and proprotor loads and aircraft The defined the I performance, flight envelope as by exploratory fllght-test program will be expanded to define fully the aircraft's capability and limitations.

Evaluation of proprotor-pylon wing dynamic stabilit_ and the effects of compressibility on proprotor performance will be extended to the vicinity of 400 knots by using maximum power I and diving the aircraft.

b. Ecqnomic Proof of Conce_t

I Specific data will be recorded to establish mission effective- ness and economic feasibility. This would include data on

I conpaysloumption,ad/lift rangecapability,and endusinglrance.e engineSelectedperformanmisscie,on profilehover sfuel will be flown to demonstrate the capability of the aircraft to t perform various civil and military missions. The fllght-test ! data along with operational analysis inputs will be used to determine economic feasibility of the concept.

C. I Environmental [roof of Concept The desirability of using tilt-proprotor aircraft for particular

I civilwill beandoperatedmilitary inmiandssionfsrom willvariobeus established.environments Theto deteaircraftrmine environmental effects on the aircraft and the aircrafts effects on its surroundings. Engine failure and other emergency tests provide data I situations will be simulated. These will i

I 300-099-003 VIII-5

_L _%" , _" Z_ __ ". -- • O0000002-TSF08

f

on flight safety and safety to ground personnel, internal and _ external noise levels, and the effects of downwash and recir- culation on engine ingestion, vlsabillty and dust signatures. Approach and landing procedures will be evaluated for operation from airports and heliports as well as undeveloped areas to provide data for the development of navigational systems and to determine real estate requirements for future VTOL ports and to permit definition of slow-speed maneuver requirements for I military aircraft. The proof-of-concept program results will define the capabilities

I andrealicharacteristicsstic VTOL operationof athel stilt-proprpecificationotors andaircraftrequiremenand twills toenablebe prepared. Civil, military,, and industry planners will then be

i abandle VTOLto motveransspowifrttaltyionto devesystlems.op the needed operational aircraft C. Program Sqhedule

1 A suggested schedule for the tilt-proprotor aircraft proof-of- ,L; concept program is shown on Figure Vlll-2. The current | NASA tilt-proprotor technology program is shown as the first "_' | item on the schedule. It is assumed that the next step, the

designcommenceandon fabricationcompletzcn ooff the a25-fircraftoot prandoprotortest farticles,ull-scale would tests are I w/nd-tunnel which scheduled for the first half of CY 1970 in the technology program.

J Thisat thescheduleend of shows1971 andthe theflightworthlnessflight research testaircrafarticlet rollcompleoutte in the second quarter of CY 1972. Exploratory flight test would

j cfull-ommencescale whitluennetl.he fllghtThe wflightorthinessprogratesm twouarldticbele pawacsed inby thethe progress in the tunnel. The first in-flight conversion should occur in the third quarter of 1972 which would permit the program I Cpr7o_of-of-1_9/_. concept flight research to start early in

I Thepriatesameforproagramfoldingplan proanpdrotyptore profoof-of-cscheduloencepwoult d aircrafalso tbe prapprogram.o- The schedule would be essentially the same with events in the

I fthanolding-proprotorthe schedule aircraftshown for programthe tilt-poccrurringoprotor twelveprogram.monthsThelater phasing of the two programs could be arranged to permit the tilt-proprotor fl_ghtworthiness test article to be modified for

I components.fllghtworthiness and wlnd-tunnel te_ts of the foldlng-proprotor

I AnThisalternateplan combinesprogramtheplantilt-propand schrotoredule andis shtheownfoldlng-propon FigurerotorVIII-3 programs into one proprotor proof-of-concept program. This

shseparateould ma_eprogramthe s overalland permitprogramthemoproof-of-re econocmionccalept thanflightthe two

t 300-099-003 VIII-6 00000002-TSF09

i I _)BELLHEL,COPTER,_oM_ANY research to be conducted simultaneously with the two research "

I beaircraft.capable Sinceof testingthe singleboth typesfllghtworth_neof proprotorss articlesystems muwithoutst now a long delay for modifications, the alternate schedule shovs the design and fabrication of the fllghtworthiness article as I the first step with the design and fabrication of the ships following. This phasln K should permit the first eonve=slon to occur at the end of CY 1972 and the first stop-fold to follow

I sthreeart minonththes middlelater. ofProoCYf-of-197c3on. cept fliKht research would I

' I I I ! I I ! ! I

I 300-099-003 VIII-7

00000002-TSF13

I_ BFI-I. HE.COPTER _'OM_AN_

IX. REFERENCES _

I 1 S_L-V/STOL City Center TFan_'_ort: A_rcraft 8tudv, McDonnell Aircraft Corporation, Or'tabor 1964.

2. Seherrer, Richard, Oarrard, Dr. W. C. J., Davis, Edward M., and Morrison, Wm. D., NASA-Lockheed Short-Haul Trans-

I _ort Study, NASA Report No. SP-I16, April 1966. 3. STOL and 9TOL. System Study_ San Francisco-Los Angeles

l Model, Federal Aviation Agency, December 1966. 4. Statler, W. H., Applications of VTOL and STOL Aircraft, Boston to Washington, AIAA Paper" No. 66-964, December 1966.

l 5. Veno, L. B., A Perspective Look at VTOL, General Electric z Company, SAE 9aper No. 670686, December 1966.

_, AIAA Paper No. 66-943, December 1966.

7. Nelson, Robert A., Northeast Corridor Transportation Pro- ject - A Challenge for Systems Analysis, AIAA Paper No. 66-916, December 1966.

8. Statler, William H. and Blay, Roy A., Role of the Rotary poowoeaAoWing in Futur@ Short-Haul Transportation, Lockheed Horizons, I Lockheed-California aompany, BurbanK, California, July 1967. 9. Hayward, Donald W., Stud_ of Aircraft in Short-Haul Trans- portation Systems for 1985, The Boeing Company, AIAA Paper October No. 67-771, 1967. 10. Armstrong, LTC Marshall B., Tactical Uses and Future Con-

ceptsin Marinefor CorpsTiltin_Operations,cProprotor Low-bisc-Loadin_Presented at AHS/U.S.VTOL ArmyAircraft Symposium on Operational Characteristics and Tactical Uses

of Vertical Lift Aircraft, November 1967. ii. Asher, Norman I. et al, The Demand for Intercity Passenger Transportation by VTOL Aircraft, Report R-144 Institute for Defense Analyses, August 196B. ].2. Pan American World Airways, Inc., Northeast Corridor VTOL

AeronInvestigauticsation,Board,Docket1969.No. 19078 submitted before the Civil

I 13. Sixty-fiveV/STOL TechnPapersology'andand ThreePlanningPanels,Conferenceto be presentedsponsored abyt the Air Force Flight Dynamics Laboratory, Las Vegas, Nevada, ,,, 23-25 September 1969. !

I 300-099-003 IX-1 00000002-TSF14

RFI .I. HEI .ICOPTER c_,mr,Ai_y

IX. REFERE,NCES (Cont_nu_.d)

14. Bell Helicopter Company Exploratory Definition Final. Report, Model 266 Composite Aircraft Program, Technical Volume Report No. 266-099-217, July 1967.

15, Hafner, R., The Domain of the Convertible Rotor, AGARD V/S_TO_craft, September 1964.

16. Lichten, R. L., Graham, L. M., Wernicke, K. G., and Brown, E. L., A Survey of Low-Disc-Loading VTOL Aircraft Designs, AIAA Paper No. 65-756, November 19_5.

Z 17. Lichten, Robert L., An Analysis of Low-Disc-Loadin_ VTOL Aircraft Types, Presented at AGARD, Paris, France, January 1966.

iB. Fischer,ys. J. Nile, Matthews, Davis W., and Pearlman, Jerome, 'i '_ CostVTOL EffectivenessAlternative Comparison,Intra-TheaterPresentedAir TraatnsportAIAA MilitaryS_stems-A

19. Stepniewski_ W. Z. and Prager, P. C., VTOL - New Frontier _ Aircrof Flight,aft SystemsVerti-Flite,Meeting, AprilOctober1967.1966.

20. Brown, E.. L_ and Fischer, J. N., Comparative Pro_ections of Helicopters_ Compound Helicopters and TiltiD_-Proprotor

Low-Disc-LoPaper No. a67din_939 PresentedVTOL Aircraftat thefor 4thCivilAnnualApplicMeeting,ations, AIAA October 1967.

21. Deckert, W. H. and Ferry, R. G., Limited Fli_ht Eyaluation of the XV-3 Aircraft, Air Force Flight Test Center TR-60-4, May 1960.

22. Anonymous, Pilot Evaluation of the Bell Model XV-3 Vertical Takeoff and Landing Aircraft, U.S. Army Transportation

I Materlel Command Report ATO-TR-62-1, February 1962. 23. Koenig, D. G., Greif, R. K., Kelly, M. W., Full-Scale Wind-Tunnel Investigation of the Longitudinal Characteris- tics of a Tilting-Rotor Convertiplane, NASA TN D-35, I December 19_9.

I 24. ModelBell Helic266 oCpteompositer CompanyAircraExploratoryft Program, DefinitiTechnicalon FinVolumeal Report, Report Nos. 266-099-201 - 223, July 1967.

I 25. Wernicke, K. G., Tilt-Proprotor Composite Aircraft_ Design State of the Art, Presented st the 24th Annual National ! F_rum of the American Helicopter Society, May 1968.

I 300-099-003 IX-2 00000002-TSG01

I

g.

IX. REFERENCES (Continued)

26. Fuselage Weight Estimation by Weight penalty Evaluation Method, Chance Vought Alrcraft Report 10043, November-r955.

27. Helicopter Transmisslon Survey, General Tire and Rubber

Company Report BE83'_2, December 1953. 28. A Low Speed Wind Tunnel Test of The Bell Helicopter 1/5 Scale 300 Force and Pressure Model, LTV Report LSWT311, to be publ_shed.

29. Loftin, L. K., Jr., and Smith, H. A., Aerodynamic

ReChyanoldsracteristicNumberss offrom15 0.NAC7A xAirfi0 °"oilto 9_0Sectixons_i0atG, NACASevenTN 1945, 19h9.

_. I 30. Abbott, Ira H., and Von Doenhoff, Albert E., Theory of Win_ Sections Includin_ A Summary of Airfoil Data, Dover Publications, Inc., _959.

' ' 31. Two Dimensional Wind Tunnel Tests of Nine Bell Airfoil Sections 9t Subsonic Speeds, Unite_ Aircraft Corporation

L I Wind Tunnel Report F930_97-I, December 1967. 32. Livingston, C. L., Rotor Aerodynamic Characteristics Program _ | _, Bell Helicopter Company Report No. 599-004-900, _, | z_ June 1967.

33. Hoerner, S. F., Fluid-Dynamic Drag, Published by the Author, _ 1958. 34. Ribner, Herbert S., Propellers in Yaw, NACA Report 820, 19;'4.

35. Etkin, Bernard, Dynamics of Fli_ht, 3ohlJ Wiley & Sons, Inc., New York, 1959. I I I !

I V

300-099 -003 IX-3 00000002-TSG02

i

I ¢ Eil::LL HI_LICOI_TI_R CO_PANY I_ 'r" I I I I I 1 X. DESIGN LAYOUTS J 1 I I I I I I I I

| 300-099 -003 X-_.

NNNNNNnO T_n£

, I II

i I

WEIGHTS ['JCIRMAL (;i_O£S _-',! ,'d_-L_! .',•, .' _o_,L" 95' EMPiY 68 EJECTION SEA! INCREASE f. NGI NE (')IL. CREW 4 / FUEL 16, TEST EOUIPMENr 4 / ,' MAXIMUM GROS c_ 124 / / / POWER PLANT • / MANUFACTURERt."AODEL PRATT P,_ ; _ ...... WHITNEY i T 30 M_NUTE PO_-R (?..=l(50_ 2,3 POWER LOADING (30 MINUTE)

///" / // MAX CONT POWER (2=9951

DISC AREA/ROTCR 49

-_-,_ 192 0 BLDISACDE LOADINGAIRFOIL (THNORMEORETICALAL GROSSROOTWEIGHT)9N£L

BLADE TWIST EFr-ECTIVE 4,' TIP SPEED HELICOPTER MODE 7, AIRPt ANE MODE 6(

WING SP,.N 3, AREA I 8 ...... WING LOADING (NORMAL G,L)OSS i_IGHT,_5, . ASPECT RATIO AIRFOIL TIP(ROOT N FLAP AREA/SI_E AFT OF HINGE 5 AILERON AREA/SIDE AFT OF HINGE I(

EMPENNAG[ HORIZONTAL I(,I [_ AREA h ASPECT RATIO 4 ELEVATOR AREA TOTAL AFT OF HINGE I 1 VERTICAL TAIL AREA

r//4 RU09ER AREA AFT OF HINGE

,RSION AXIS ASPEGTRAT,O ,

] .... • Io6o

/.- -=_ 1B:4,5" .._ i ...... ! o.ou_oL,NE

-457.5 ...... ,_ SCALE ..... -- -_01(; -- -- ' I

'i ,, !,1 !lfz t I_ _1, O0000002-TSG06

_T£R],5ILC S WEIGHTS NORMAL ('E_SS It;:Eb{-,;.HIt1 MIbSfON_ _500 LR EMPTY 68?6 L8 EJECTION SEA1 INCREASE I;4 L9 ENGINE OR_ 35 L8 CREW 400 LB FUEL 1600 tB TEST EQUIPMENT 475 LB

MAXIMUM GROS_c 12400 LB

POWER PLANT =-_- MANUFACTUR_-RtMODEL PRATT WHITNEY PT6C-40 MAX CONT POWER (2_,995) 1990 SHP ! .30 MINUTE POWER (2=llSOJ 2300 SHP POWER LOADING {30 MINUTE) 4.13 LB, HP

ROTOR DIAMETER 25 FT DISC AREA/ROTCR 491SQ FT 192 0 DISC LOADING (NORMAL GROSS WEIGHT]9.67 LIEVSQ FT BLADE AIRFOIL THEORETICAL ROOT NACA 64.935.,-0,3

BLADE CHORD TIP N14ACINA 64.206..0.3 . .'_, . SOLIDI TY .089 BLADE TWIST-EF.TECTIVE 45 DEG TIP SPEEO HELICOPTER MODE 740 FT/SEC A J#PL,_NI: MODE 600 FT/,SEC

WING SPAN 34.3 FT AREA 176 SQFT WING LOADING (NORMAL GROSS WEIGHT)54.0 LB/SQ FT ASPECT RATIO 6.63 AIRFOIL TIp_'ROOT NACA 64A223 HOD FLAP AREA/SIDE AFT OF HINGE 5.5 SQFT AILERON AREA/SIDE AFT OF HINGE I04 SQFT "

EMPENNAGE HORIZONTAL TAIL AREA 62.5 SO FT ASPECT RATIO 4.1 ELEVATOR AREA TOTAL AFT OF HINGE 17.6 SQFT VERIICAL TAIL AREA 5?.8 SO FT ASPECT RATIO 1.84. RUDDER AREA AFT OF HINGE 7.6 SQ FT

-!

J!60 f8_30

-GUOUhD LINE DESIGN o _._ 3_ LAYOUT SCAL_. I1._1MOOEL 300 | TILTROTOR AIRCRAFT ,

L______,_H _:_ 6_,/2o ._ . .. [_.;.,...t.. "--_ '. 3"_ , .(), L L 00000002-TSG08

300-010-503 SPINNHR SUPPORT- "\ -- 300-010-0,06 COLLECTIVE LEVER ASSY _ _k_

// f-300-010-_05 COLLECTIVE HEAD ASSY

,k

ED 6Y)LLECTIVE Tt_E i"

300-010-t_08 ROTATING CONTROL

300-010-_I0 DRIVE SLEEVE

300.OkO O0k BLADE ASSY _-

300-0[0-500 SPINNER ASSY- / PITCil LINK--

300.010-100 PROPROTOR ASSY _ / 3t _oo-o10-l.Ol YOKE;-./

m 00000002-TSG09 OONNNNNg_T_ _n

_,...._=300-0_.0-0_8 PITCH HORN

j_ CYCLIC TUB_:

.__ _--300-01.0-402 INNER RING AI,SY - SWASHPLATE

_ _ - FIXED CYCLIC _USE INFUT _BER

O -0_0-204 FORK

O _0_0-302 FLAPPING _OP

" ].rT"_--. _ ; /-'-TRANSMI SSION

_[ _-_..]_ _REF 300-96¢-004

--300-010-/¢II¢DRIVE ASSY - SWASHPLATE L--'_%300._ r_io.502 RING ASSY St.;ASHPLATEDRIVE o[ INCHES _l I _, f .,s,o.

PROPROTOR AND CONTROLS L._o_viN_rO_-69#Z"..8 _ _ I ''_ _ 3_0-960-002 I_ULL_U_l ' L'_,,_.,j.,,

00000002-TSG12

..o,<,..i-Sll I }_1). )(JtM' r_',,!ll', .[ ! I - . - ......

I I I I I

00000003-TSA07

COLLECTIVE CObrt ROL - CYCLIC CONTROL 300-960-005 FIXED CONTROL:{

FIXED CONTROL NOUNT FIREWALL SHUT-OFF VALVE, FUEL BLOWER DRIVE SHAFT BLOWER, OIL COOLER BY-PASS ETA RT ER- GENERATOR

EJECTOR, INDUCTION BY-PASS

!

PLENUM, INDUCTION BY-PASS ......

= I FUELFIRE PERXESSTINGUIURE SHERTRANSMI...... TTER - .- ' FUEL FILTER, ENGINE ...... i • OIL FILTER, ENGINE - I

_4 GASPRODUCERCONTROL .....

°_ I ENGINE AIR INLET SCREEN •

FIREWALL SHUT-OFF VALVE, FUEL FIRE EXTXNGUISH_ .....

I CLOSE-OUT BAFFLE, FIREWALL -.

QAS PRODUCER CONTROL • i PRESSURE TRANSMITTER, FUEL ...... TRANSMISSION OIL COOLER BLOWER, OIL COOLER .....

TRANSMISSION OIL COOLER --. BAFFLE, OIL COOLER .... i EJECTOR, INDUCTION BY-PASS

i ENGINE OIL COOLER PL_UM SEAL

I ...... G STARTER GENERATOR I BLOWER, OIL COOLER _- -

ENGINE OIL COOLER

I BY-PASS DUCT, COOLING AIR - COWL SUPPORT

i BLOWER DRIVES_.FT FIRE DRT_CTOR _. FILTER, ENGINE F(_L

I SECTION A-A ZNFILTEDUGTR,IONENGINEBAFFLEOIL" SOREENt INDUCTION ANTI-ICE

!'OLDOUT _1'" ' "......

j. 00000003-TSA09

300-960-007 WING AIR-INLET - UOT SECTION - CONVERSION ACTt_TOR

#/ .- FAIRING \ | TRANSMISSION OIL MANIFOLD J I /// HYDRAULIU RESERVOIR & FILTER

--- :_ . "" S_A. 193.00

.... CYLINDER, COLLECTIVE GONT_OL CYLINDER, CYCLIC CONTk.OL UP • ENGINE EXHAUST STACK _ / HYDRAULIC PUMP

- , BAFFLE, UEC_OR SEPARATOR "WING ETA ! = ...... TUEJERNINCTOGR, VANEXHE,AUSTEXHAUST OOTB RD _ 1. / HYDRAULIC RESERVOIR & F_ CONVERSION DOWN-STOP L93.00 ,/ _ I"UBES HYDRAULIC PUMP I " INLET, COF_ARTMENT COOLING "_ iO °

300-O10-00I BLADE ASSY _ i / ROTATING CONTROLS ---._/ • .- 300-960-002 PROPROTOR AND CONTROLS \ ,/ ....ACTUATOR FAIRING CONVEI ". .-SI'_,NER ..,._...... _ . , \\ ,_

_! W.L. I00.00 " _) . ACTUATOR FAIRING \ ". - CONVERSION ACTUATOR -TF_NSMIESION OIL pL_4P _i " pOWER TURBINE GOVERNOR, HE_.IGOPTER MODE GONVERI j- "_ - SPEED SELECT ACT_TOR, RELIGOI_ER MODE . F_REWALL, INB'D _\" - DROOP CO]_ENSATOR C=_4, HELIGOPTER MODE

j " -- TORQU_ETER & POWER TURBINE SENSOR _GINEAIRIN_ I FIREWALL CYLINDER, CYCLIC CONTROL ] CIT_INDER, COLLECTIVE CONTROL

SPEED SELECT AL_FUATOR ..... POWER TURBINE GOVERNOR

300-960-00_¢ MAIN TRANSMISSION -_• _ INCHE_ 0 TORQUEMETER & POWER TURBI HE SENSOR. , --S-_LE FIRE DETECTOR /

PT6CFI REW-_AOLL ENGINE TFRANIREWASMISSILL, ON UPPEORIL SCAVENGE LINE . --_ - / / _" ,li. i

4-.

_ TRACE AT WING S_A. 193.00 00000003-TSAll

00000003-TSA13

'1 SIDE VIE_/ SHOWN OF ENGINE INPUT GEaR TRAIN 00000003-TSA14

/ CONTROL PIIA,_INP, AL_rUATOR " (R_:F 3nO- 960-O07) J

i]_' _ ! FLAP DRIVI,:ACTIJATOIt

/t ; / /

COLLECTIVE CONTROL TO LEFT ROTOR-

i AILERON CONTROL TO LEPT ROTOR- ,

i / b CYCLIC CONTROL TO L_T ROTOR

!!,. 1 J

1 t

L '

I DIFFERENTIAL COLLECTIVE PHASING LEVERS

t COLLECTIVE RANGE-SHIFt CRANK--.-. I

| COLLECTIVE CONTROL FRON PILOT" _

LATEI

I DIRECTIONAL cONTROL FROM PILOT- CONI CONTROL TO RUDDER (REF 300-960-006) !

I . .. C_NTERLINF AIR_ 00000003-TSB01

DRIVESIIAI_I' INTEROONNEC_ OFARBOX (REF)

AFT SPAR TNANS_IISSION (REI_ GOLLE_'IVE WING DEFLEG_ION [ ISOLA_ ION BELLCRANK \ I

\ \ !/ AILERON SERVO ACTUATOR

< . AI LERON "FLA_ MIX _NO ORANKs 'X ' A // \ "-"1; t - , ,,'_'"_',

...... - _------_-_-- ...... ', \ Ld/ - _"_-_ --_---_--:L i-_" • - -

\- \

__ ...... " ...... - -CONTROL TU_E FAII_.,ECD TYP

300-960-007) _ _-

• --'_ DIFFERENTIAL CYCLIC PHASING LEVERS _ --_

:_[PFERENTIAL CYCLIC PHASING ACTUATOR . AIRSPEE D

:-'_TROL FROM PILOT

:_L FROM PILOT

:_--_EVATOR _R_ 300-960-o06)

t

\ \

' CENTERL INE DRIVESHAFr

_[i_')L1)()ll'l'I "

, ..... ------2. - _ _ [ 11 -- _ L- _ i , ii .... ' _ -_}mm " I I -- i -- T • • V1 O0000003-T£B02

! C_:NTERLINR MA._T :.'0 A_'UATOR _'_

" _ ."_.i / PYLON CONTROLS SUPPORT "I,A._

I:_ i / _oAILEROsI'nONN C.ONT,oLAILERO" "'\i i ,_l_,i / n +_oo +2s.o,,/-15.oo

'J{[/ ) _r -30" +].8 00/-I2.5 ° f¥1_I , I.[,',_i:'•'_ "¢ '__ ' -60° .14.2o/. 5.0°

- " " "LI_,'" "'_- _J_ " "_CYCLIC TORQUE TUBE

/- o =Ioco RoLTDBE

(WING HOUSED) / /,

i _"---CONTROLS SUPPORT / '_- -COLLEC']_Yw. enNTRnT, TIIR_ // / DRIVESHAFT COUPLING (REF)

, ,. -7 .....<" ." ...._'_, {

LE_N (_F 300-96C-007) ; -" "'___i --:} "-t " _--_--,_-/_i

--....-...... - // : "'--AILERON=NT LLINF OE-- _..\ I/ SECTION A-A 65_ ROTATED SEE TABLE C

j / ,DRIVESHAFT (REF)

! / , \ ', l_ #_ CYCLIC OONTROL TUBE- / -AILERON CONTROL LINKAGE ',I -(_)LLEGTIVECONTROL LINKAGE

SEL'fION 8-B ROTATED % 3. 4_ INDICATES BELLCRA_K PIVOT POINTS I '\ ATTACHED TO FIXED STRUCTURE i CONVERSION SPINDLE (R_) 2, COLLECTIVE CONTROLS SHOh%_ IN MAX PIT_[ POSITION - ALL OTHER CONTROLS SHOWN IN N EbTRAL POSITION t. ALL CONTROLS SHOWN IN AIRPLANE MODE _= _ DOTES 00000003-TBB03

I

--',]CONTROLS SUPPORT I{ TABLE C , _

i--FLAPPosITIONAILERON I i " AILERONCONTROL

i +lO0°o I +2+215."0°6o//'1-155+2'0° °

j -30 ° +L8 0a/-12.5 °

-60 ° -14.2°/- S.O °

-- CYCLIC TORQUE TUBE raOTuLECrIVE SWIVEL /---CYCLIC CONTROL TUBE /

_ORT (WING MOUNTED) //_ COLLECTIVE CONTROL TUBE

/ / DRIVESHAFT COUPLING (REF)

/ ---AILERON CONTROL LI_KAGE

\ _/ SECTION A-A

SEE 65TABLE0 C / ROTATED /_/+'DRIVESHAFT (REF)

:_.+ ......

AILERON CONTROL LINKAGE CYCLIC CONTROL TUBE COLLEG_IVE CONTROL I,IN_AOE

SECt ION B-B %_ ,

3. @ INDICATES BELLCPd_K PIVOT POINTS ATTACHED TO FIXED STRUCTURE LAYOUT 2, COLLECTIVE CONTROLS SHOWNIN MAX PITOI _ POSITION - ALL OTHER CONTROLS SHOWN IN l FIXED CONTROLS - WING NEUTRAL POSITION I_ ESSARY _.:6g____.__ t. ALL CONTROLS SHOWNIN AIRPLANE MODE _ _'_60-_)_-_ j'_i_,, _,_" _NOTES_ - ....

IF 1 " i --i "---[--[ ..... " "_I i +I I I III I II i . • 1 !

CONTROL STICK LATERAL OONTROL TO MIXI_O pA_#.AOE.._ -FORE AND AF'J PILOT (REF 300.960-00",) _ , / _. \_ O0000003-TSB06

i J

CONTROL TO MIXING PACKAGE (REF 300-960-005) CONTROL TUBE FAIRLEADS TYP

I I / /

i" _/J I

STA 529. I I STA 478.00 I 300-960-008 FUSELAGE AND EHPE_,

WL 33.00 --

_)T e1",()1TI t ,, O0000003-TSRO7

STA 529.72 I

I i(}-960-008 FUSELAGE AND _dPENNAGE (REF) 2. COLLECTIVE CONTROLS SHOWN IN MAX PITCH POSITION - ALL OTHER _,,.m.._

POSITION LAYOUT CONTROLSS_OW_I_ NEU_RA' I_J DESIGN t. ALL CONTROLS SHOWN IN AIRPLANE |_ FIXED CONTROLS

NOTES o...... m---:_,._-_I300-960-oo6 _* ,,._.,_tP U'T'.7_,, .... :/ O0000003-TSB08

AILERON -

AFT SPAR CAP 7075-T6 AL ALY __ DREXTIVERSUHASI.O_N' . 3/_ IN. HONEYCOMB PANEL •7075-T6 AL ALY .063 INNER & OUTER SKINS RIB WEB & ii AL ALY N

_,_ J 7_075-T6SPARAL_PALY - ___ EXTRUSION

_ .,

PRONT SPAR AL ALY - STIFFENED SHEAR WEE _/_ IN. HONEYCOMB PANEL •7075-T6 AL ALY SECTION 0-O •050 INNER & OUTER SKINS ROTATED 90 ° CCW ;PLICE pLATE

_NE'-,'OOME

I

HI-LOK PASTENE] RIB CAP SPLICE

SEC'IION' D-D

WING PANEL SPLICE SURFACE i TYPICAL UPPER & LOWER 0 5

I O_ .,m-- O0000003-TSBOg O0000C}(]B-T.q R i n

8EL'TION B-B O0000003-TSB11

.... W.S. [93.0 J /

r. I ] / / g. •

,. / CONVERSION ACTUATOR

f / ...... t!

...... _ _ MAIN TRANSMISSION

[ _/ ' REF 300-960-00_ -7075.T6 AL ALY MACHINED

REF 300-960-005

STA 300.0 I- 5°30' / l

COLLECTIVE CONTROL REF 300-960-005

I_ _-- CONVERSIONAXZS

l SECTION A-A

i TRACE W.S. 193.0 FLAP @ STA 300.0

_)J o i.o

,NEYCUMB PANELS SCALE & ODTER SKINS

L°AYOUT,,°,

u W'{NO ASS_94BLY _-69_ I mm ,_.,7,. 1300-960-007

'!

11 , I . I \ -- - ,'I ,,_ . ,1 ...... / ...... _- 1-1...... _...... _,:_......

H " ,I - -.. _ {1._...... ,...... i o' ._.____,i- .... " : ' ', 71_...... --_ ...... -_-_'_-__i "

/ [ _ _ .... _i_V" " I' II ------" b\ i ,, \ :! .... l....

_--- CONTROL TUBE ROUTING TUNNEL O0000003-TSC01

q I

STA _TAoO

ii , r...... I I .... "--'_-" ...... I' =: _ I i O0000003-T.q_n9

BL 0

k:L 87.0

6t ,0 IN.C'LKAP,ANOE

5_,o0 IN. C_EARANCE UNDER WINO

1 i

AL RIGIDIZED SKIN I[ | ALY UPPER BL 7.75 BL '7.75 AL ALY LOWER SKIN B t

SECTION 0-C

I r

BL 0

CONSTRUCI'IO|I _ _ _ _

SRIN - STIFFENER -_

\ ! / Ii

\

V%'R' SKIN OON:

/-- _LEVATOR SKIN - RIB CONSTRU6TION \

\

\ /- --" I/2 IN. HONEYOOMB PANEL / AL ALY SKINS _/ UPPER & LOWER SURFACE

00000003-TSC06

I I

I ,'OLDOU'£lfllAMl_ / 00000003-TSC07 O0000003-TSC08

cf 00000003-TSC09

lain II

\

'_I LAYOUT°"'°" ,o I,_, _ _ __ _T_ I

II FOI,IJ(JU'J_ Ic' /!