HEAVY LIFT HELICOPTER ADVANCED TECHNOLOGYUSAAMRDL-TR-77-41(i ) HEAVY LIFT HELICOPTER -ADVANCED...
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USAAMRDL-TR-77-41(i ) HEAVY LIFT HELICOPTER - ADVANCED TECHNOLOGY COMPONENT PROGRAM - ROTOR BLADE Boeing Vertol Company iest P.O. Box 16858 Philadelphia, Pa. 19142 D D < September 1977 L7 ,J Final Report for Period July 1971 - July 1975 Approved for public release; distribution unlimited. Prepared for U. S. ARMY AVIATION RESEARCH AND DEVELOPMENT COMMAND S.O4 P.O. Box 209 St. Louis, Mo. 63166 - APPLIED TECHNOLOGY LABORATORY U. S. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AVRADCOM) Fort Eustis, Va. 23604
Transcript of HEAVY LIFT HELICOPTER ADVANCED TECHNOLOGYUSAAMRDL-TR-77-41(i ) HEAVY LIFT HELICOPTER -ADVANCED...
USAAMRDL-TR-77-41(i )
HEAVY LIFT HELICOPTER - ADVANCED TECHNOLOGY
COMPONENT PROGRAM - ROTOR BLADE
Boeing Vertol Company iestP.O. Box 16858Philadelphia, Pa. 19142
D D
< September 1977 L7 ,J
Final Report for Period July 1971 - July 1975
Approved for public release;distribution unlimited.
Prepared for
U. S. ARMY AVIATION RESEARCH AND DEVELOPMENT COMMANDS.O4 P.O. Box 209
St. Louis, Mo. 63166
- APPLIED TECHNOLOGY LABORATORY
U. S. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AVRADCOM)Fort Eustis, Va. 23604
I
APPLIED TECHNOLOGY LABORATORY POSITION STATEMENTr
Due to the termination of the HLH program, reports summa-rizing the strides made in many of the supporting technologyprograms were never rFblished. In an effort to make as muchof this information lable as possible, selected draftreports prepared under contract prior to termination havebeen edited and converted to the DOD format by the AppliedTechnology Laboratory. The reader will find many instancesof poor legibility in drawings and charts which could not,due to the funding and manpower constraints, be redone. Itis felt, however, that some benefit will be derived from I
their inclusion and that where essential details are missing,suff.cient infcrmation exists to allow the direction ofspecific questions to the contractor and/or the U.S. Army.
I.I
DISCLAIMERS
The findings in this report are not *o be construed as oil official oDoart-nent of the Army position unless sodesignated by other authorized documeibts.
Wihan Government drawings, specifications, or other data tre useo for any purpose other than in coisnectionwith a dofinitely related Government procurement operation, the United Statos Governmlene thereby incurs no
res|ponsibilily nor any obligation whatsouver; and tie fact that the Government nit.V have formulated, furnished.
or in any way ;upplied t:t.c ;.:.d drawings. specifications, or nther data is not to be regardec. by implication or
otherwise as in any manner licensing tire holder or any other person or corpo arion, or conveying any rights or
permission, to mninufacture, use, or sell any patented invention tll'&. may in any way be related thereto.
Trade names cited in this report do not constitute an offiial undorsunient or approval or thIe use of suchcommercial hardware or software.
DISPOSITION INSTRUCTIONS
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ýSAAMRD T3R-77-41
4. TITLE (andu sbtlt .-.... .. OVERE;II FT UELICOPTER - ADVANCED TECH- F e ' - ....
/NOLOGY COMPONENT PROGM - ROTOR BJu 7 - .... 75*
.AT_-OR 8. CONTRACT OR GRANT NUMBER(a)
9. PERFORMING ORGANIZATION NAME AND ADORESS 10. PROGRAM ELEMENT. PROJECT. TASK
Boeing Vertol Company AREA & WORK UNIT NUMBERS
P. 0. Box 16858Philadelphia, Pa. 19142
Ii. CONTROLLING OFFICE NAME AND ADDRESS . REPORT DATE
U.S. Army Aviation R&D Command Sep 77P. 0. Box 209 13. NUMBEROF PAGESSt. Louis, Mo. 63166 201____________-Ii. MONITORING AGENCY NAME & ADORESS(l"different from Contolilng Office) 15. SECURITY CLASS. (of this report) ,
Applied Technology Laboratory, U. S. ArmyResearch & Technology Laboratories Unclassified(AVRADCOM) 1S5. OECLASSIFICATION)DOWNGRADING
Fort Eustis, Va. 23604 SCHEDULE
16. DISTRIBUTION STATEMENT (of this Report)
Approved for public release; distribution unlimited.
17. DISTRIBUTION STATEMENT (of the abstract entered In Block 20, It different from Report)
I8. SUPPLEMENTARY NOTES
19. KEY WORDS (Contlnue on recerae side If necessa-ry and Identify by block number)
Rotor Blade MaintainabilityFiberglass ReliabilityFail Safety Controlled CostImproved Rotor Performance
0. AssrR ACT (ContM.~e mverme atb if-,c..eaw and idem!ify by block number)
"This report reviews the development of the Model 301 Heavy LiftHelicopter rotor blade. It describes the design, structuralanalysis, testing, and manufacturinq process of the HLH rotorblade.
FORM
TABLE OF CONTENTS
LIST OF" ILLUSTRATIONS ................ ................. 5
LIST OF TABLES ............. ..................... 10
1.0 INTRODUCTION ............ ................... 11
2.0 SUMMARY ............... ...................... 15
3.0 DESIGN DEVELOPMENT .......... ................ 21
3.1 Design Goals and Objectives ... ......... .. 213.2 Rotor Blade Geometry and Sizing ......... .. 243.3 hotor Blade Structural Concept ........... .. 333.4 Preliminary Design Studies and Support Tests. 363.5 Detail Design of the ATC slade Configuration. 453.6 Detail Design of the Protitype Blade
Configuration ........ ............... .. 85
4.0 STRUCTURAL AND AEROELASTIC ANALYSIS ........... .. 119
4.1 Criteria and Requirements ... .......... .. 1194.2 Limit and Ultimate Loads .... ........... .. 1194.3 Design Fatigue Flight Loads ... ......... .. 1204.4 Material Properties . . . ........... 1204.5 Blade Physical Properties ... .......... .. 1214.6 Ultimate Strength Analysis .... .......... .. 1214.7 Safe Life Fatigue Analysis .... .......... .. 1334.8 Fail Safe Analyses ..... .............. .. 1334.9 Natural Frequency Analysis .... .......... .. 134
4.10 Classical Flutter 1............... 354.11 Rotor Blade Torsional Divergence......... 135
4.12 Pitch Lag Stability . . . ........... 136
5.0 MANUFACTURING DEVELOPMENT ................... 140
5.1 Spar Fabrication ..... ................. 1405.2 Tooling ...... .... ................... 1425.3 Forming of Titanium Cap .... ........... 1425.4 Quality Assurance ........ .............. 143
6.0 DEMONSTRATION TESTS ......... ................ 163
6.1 Full-Scale Component Structural Tests .... 1636.2 Natural Frequency and Stiffness Tes ..... .. 1756.3 Lightning Tolerance Evaluation ........... .. 1776.4 Wind Tunnel Demonstration Test ........... .. 1796.5 Rotor Whirl Demonstration Test ........... .. 193
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A -T -!U=
Page
REFERENCES . . . . . . . . . . .. . . . . . . . . . . . 199
LIST OF SYMBOLS ................ ..................... 201
ly .... ............. .. ........
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LIST O ILLUSTRATIONS
Figure Page
1 General Arrangement - Model 301 HLH ......... .. 13
2 Completed Blade Number 1 .... ............ .. 17
3 Root End View of Completed Blade ... ........ .. 18
4 HLH Rotor Blade ......... ................. .. 19
5 Rotor Airfoils .......... ................. .. 26
6 Maximum Lift Boundaries and Zero Lift PitchingMoment Levels ........... .................. .. 27
7 Maximum Lift Boundaries and Pitching MomentLevels .............. ..................... .. 28
8 Drag Comparison, Wake Probe Data (HLH Reynolds
Number) ............... ................... .. 29
9 HLH Rotor Blade Geometry .... ............ .. 30
10 Chord Trade Study ......... ................ .. 31
11 Rotor Flying Qualities Boundary - Steady Flightlondition ........... ................... 32
12 !:*I Rotor Blade ......... ................. .. 34
13 HLH Spar Assembly ......... ................ .. 35
14 Trade Study - Rotor Blade Concepts ......... .. 40
15 Tension Fatigue Failure of Graphite Composite(Room Temperature) ........ .............. .. 41
16 Tension Fatigue Failure of Graphite Composite(160 0 F) ............. ..................... .. 42
17 Shear Fatigue Failure of Graphite Composite. 43
18 Impact Test - 1 Lb Ball on +450 Graphite andNomex Core .......... .................. ... 44
19 Impact 2est - 1 Lb Ball on +450 Glass andNomex Core ............ ................... .. 44
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Figure Page
20 Design Support Tests ...... .............. 54
21 Coupon Fatigue Tests of 6AL4V Titanium AlloySheet ............... ...................... 56
22 Wraparound Root End ....... ............... .. 57
23 HLH Rotor Blade Aft Fairing .... ........... .. 58
24 Fiberglass Composite Does Not Lose StrengthAfter Operating in Service Environment ..... .. 59
25 Comparison of Fdtigue Strength of 1002S andSP250 Unidirectional Fiberglass EpoxyComposites ............ ................... 60
26 Location of Chordwise Center of Gravity. ..... 61
27 HLH/ATC Rotor Blade Assembly Drawing Tree. ... 62
28 HLH Rotor Blade Assembly .... ............ 63
29 Geometry Two-Pin, Fittingless HLH Rotor Blade. 69
30 Bonded Assembly - Rotary Wing - HLH ......... .. 71
31 Spar Assembly - Two-Pin, Fittingless HLH RotorBlade ............... ...................... 73
32 Prototype Blade Modifications .... .......... .. 91
33 Lightining Protection ..... .............. .. 92
34 Typical Crossply Wrinkles in Spar Heel Area. 93
35 1-Inch Section of HLH Prototype Spar PrecureHeel, Fiberglass and Graphite .... .......... .. 94
36 Comparison of HLH Spar Heels Before and After
Modifications ......... .................. 95
37 Aft Fairing Core ................. 96
38 Aft Fairing Skin ........ ................ 97
39 HLH/Prototype Rotor Blade Assembly DrawingTree .............. ...................... 98
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Figure Lal
40 Blade Assembly Prototype HLH Rotor Blade .. . 99
41 Bonded Assembly - Potary Wing - HLH ......... ... 105
42 &par Assembly Two-Pin, Fittingless HLH RotorBlade ............... ...................... 107
43 Basic Design Requirements ..... ............ .. 122
44 Rotor Blade Moments for Design Limit ManeuverCondition ............. .................... 125
45 Rotor Blade Design Moments for High-Speed LevelFlight .............. ..................... 126
46 Rotor Blade Centrifugal Force .... .......... .. 127
47 Spanwise Distribution of Mass and Stiffness. 130
48 Spanwise Distribution of Blade Axes .......... .. 131
49 Roto- Blade Frequency Spectrum .... ......... .. 137
50 Demonstration nf Blade Torsional AeroelasticStability at High Speed With 14-Foot-Diametez.7Model Rotor ........... ................... 138
51 Comparison of Analytical Pitch Link Load With14-Foot-Diametei Rotor Wind Tunnel Test Loadfor Limit DtSign Dive Speed .... ........... .. 139
52 HLH Rotor Blade Fabrication Sequence ........ .. 145
53 HLH Blade Fabrication Flow Diagram ......... .. 146
54 LeadJ ig-Edge Assembly Flow Chart ... ........ .. 147
55 Spar Assembly Flow Chart ...... ............ 148
56 Root End Fitting Fabrication Flow Chart...... 149
57 Tip Fitting Assembly Flow Chart .... ......... .. 150
58 Fitting Installation Flow Chart .... ......... .. 151
59 Fairing Fabrication Flow Chart ... ......... .. 152
60 Blade Bonding Flow Chart ...... ............ 153
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Figure Page
61 Damper Arm Assembly Flow Chart .... ......... .. 154
62 Final Assembly Flow Chart ..... ............ .. 155
63 Main Bond Tool Being Closed .... ........... .. 156
64 Spar Bonding Fixture Ready for Use ......... .. 158
65 Spar Curing Operation Heating Cycle ......... ... 159
66 Titanium Cap Forming Tool ..... ............ .. 160
67 Tooling Improvements for Titanium Nose Cap . . 161
68 Quality Assurance Flow Chart .... .......... 162
69 HLH Rotor Blade Structural Test Specimens . . 168
70 Full-Size Root End Tests Verified DesignAllowables for Unidirectional Fiberglass . . . . 169
71 Fiberglass Damage Tolerance and SurvivabilityDemonstrated by HLH Root-End Test ........... .. 170
72 Intermediate Section Fatigue Test Fixture... 171
73 Titanium Nose Cap Fatigue Failure ........... .. 172
74 Fatigue Strength of Highly Directional 6AL4VTitanium Alloy Sheet With Effect of MoltenDeposits .............. .................... 173
75 Nomex Honeycomb Prototype Configuration.... ... 174
76 14-Foot-Diameter Model HLH Rotor BladeInstalled in Wind Tunnel ...... ............ 183
77 Hover Figure of Merit HLH/ATC 14-Foot-DiameterRotor Wind Tunnel Test ...... ............. 184
78 Cruise Efficiency HLH/ATC 14-Foot-DiameterRotor Wind Tunnel Test .......... ............. 1R5
79 Flying Qualities Boundary HLH/ATC 14-Foot-Diameter Rotor Wind Tunnel Test .... ......... .. 186
80 Rotational Noise of 14-Foot-Diameter ModelRotor ............... ...................... 187
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Figure Page
81 Comparison of Analytical Blade Bending MomentsWith Scale 14-Foot-Diameter Rotor Wind TunnelTest Data . . . . . . . . . . . . . . . . . . . . 188
82 Comparison of Design Pitch Link Load With Scaled14-Foot-Diameter Rotor Wind Tunnel Test Data 189
83 Comparison of Analytical Pitch Link Load With14-Foot-Diameter Rotor Wind Tunnel Test Loadfor Level Flight Design Condition . . . . . . . . 190
84 Fixed System Control Load Reduction With Damping,14-Foot-DIameter Model Rotor Test . . . . . . . . 191
85 Blade Torsional Load Growth Fixed Load CriteriaFrom 14-Foot-Diameter Model Rotor Test - - . - - 192
86 Rotor Whirl Test . . . . . . . . . . . . . . . . 1.94
87 Dynamic System Test Rig . . . . . . . . . . . . . 195
88 Rotor Hover Efficiency From Whirl Test . . . . .. 196
89 Blade Bending Moments From Whirl Tower and DSTRTests . . . . . . . . . . . . . . . . . . . . . . 197
90 Blade Natural Frequencies Determined From Test 198
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LIST OF TABLES
Table Page
1 Rutor System Description .... ........... ... 20
2 Rotor Blade Design Goals and Objectives .... 23
3 Comparative Properties of Blade Materials 55
4 Lasic Fatigue Loading Schedule ............ ... 123
5 Maneuver Air.;peed Distribution ............ ... 124
6 Material Properties and Design AllowablesSummary ........... .................... .. 128
7 Calculated Weight and Centrifugal Force .... 129
8 Minimum Margins of Safety .... ........... .. 132
9 Nondimensional Natural Frequency ........... ... 134
10 Coefficient of Thermal Expansion Comparisons. 157
11 Comparison of Theoretical and ExperimentallyDetermined Blade Bending Stiffnesses ........ ... 176
12 Comparison of Theoretical and Measured NaturalFrequency ........... ................... .. 176
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1.0 INTRODUCTION
The initial HLH preliminary design concept was formulated andsubmitted to meet the mission requirements and broad designcriteria for the Advanced Technology Component (ATC) Program.This preliminary design concept included an initial estimateof the aircraft size and weight, a definition of the majorsubsystem interfaces, an identification of potential riskareas and, most importantly, an identification and definitionof those components considered to be of advanced technologyand requiring advanced development.
BACKGROUND
The purpose of the ATC Program was to seek maximum reductionof technical and cost risk associated with the EngineeringDevelopment of an HLH system through the design, fabrication,demonstration and test of selected critical HLH components.Engineering Development or full-flight qualification of anycomponent or concept was riot the purpose of this program.
The critical components of the HLH were determined to be therotor blades, hub and upper controls, drive system, flightcontrol system and cargo handling system. The scope of theHLH ATC program was limited to these components, plus theinterface analytical activities necessary to assume the A"Ccomponents would be suitable for subsequent integration withthe complete aircraft. The general arrangement of the HLH isshown in Figure 1.
ROTOR BLADE ATC PROGRAM
The rotor blade was an obvious selection as one of the com-ponents of the HLH Advanced Technology Component (ATC)program. It is a high-cost, long-lead item requiring highreliability and offering a potential reduction in mainten-ance hours.
The blade program was conducted during the period fromJuly 1971 through July 1975. The phases of the programincluded:
Preliminary Design - trade studies andselection of concept
Detail Design - preparation of completedrawings
Fabrication - development of manufac-turing conceptfull-scale blade fabrication
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Demonstration - Structural TestWind Tunnel TestWhirl TowerDynamic System Test Rig
This report presentq a summary or review of each of thesephases which have been reported in detail in the documentsreference herein. Descriptions of the several in.provemenýmodifications which were incorporated into the blade designfor the prototype helicopter are included.
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MA4JOP CHARACTE015TICS.
ROTORWEIGHT (LEz. IAXýETEP (P'T) 92.0 (lESION GROSS WEIGHT *LC'
'riPper &.E(U 750so.0 CSI1GJ PAYLOAD 080O3
CISC LOACING(125F)A'I COw 6.9 ICES,.F '"CLUDE' FUMNELEW .1"00
.LADE A(IEA (3 AT .5 ZQý I'4.0 PILE UEFUL LOAD 9qNLL0E S UJ 06
CXMZTRIC 5OIU.ITI' PATIO .092 ir EMPTY WEi.G'T 400
rE0mETWC D5C. APFA<2 A7 47'J 5 OF r.) QI,295.0 uUA. ALTERNATE GR055 VJE,G)4 1, T0
Pf~ ~JL.IQJIGROUND ANGLES (DEGPEES)NU~EE o ENINS/YPE-70.07O(5) TUP80135H.AFT TURNOVJER GpoUT.JD -H1E ,40' 5 LB
TRANSMISSION' PATINJC (NP.) 770 (w'T EMPTY)ýEpQF4DE NGA5 A H(.
MAXA. SINGLE EF4G'NE RATINGO p*079 TIP TY)I RUD.IE 3~ SIIIYBUF4~Ol&I
INTEGRAL FUEL CAPACITY (6A,) 2,938 (v.EPY
iJl'EGPAL FýEL CAPACITY 1,19,100 rTO Ma IVE S
FPPO AFT
'OLLECTIV'E PITCH .. ,T
DIFFEIIENTIAL COLLEC.TIVE PITCH' 1 5.0' I3.
PPOGPA1IEO LOý4TUOIKNAL LYCLIC PI'TCH' -5.?* -0 .2.0, -5.4* .T ro.0.
4RTMACON'TROI. CYCLIC PITC'H 5.5, 5.0,
DIFFERENTIAL LATEPIAL CYCLIC PITCH 7I] T
LATERAL CYCLIC PITCH o
NOE' -WHEEL /TIRE 51ZE 1 PLY.D I'I- S, I5,50`c
L.~~~PTN MAN-WEL/TO 'rIl 'TYPE IL
M~~iN~ - wHEE -TM $
Figure 1. General Arrangement -Model 301 11LH
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2.0 SUMMARY
The Boeing Heavy Lift Helicopter rotor blade is an applicationof advanced technology encompassing improved airfoil and twistdistribution and composite materials.
The fiberglass rotor blade represents a major advance in rotorsystem fail safety and reliability. The achievements of thedesign are:
* Improved rotor performance using advanced airfoils andoptimized thickness and twist distribution.
*Fail safety of the fiberglass construction and aninherently redundant root end attachment.
* Improved maintainability and reliability with apneumatic delta pressure failure system.
v Controlled cost with a production oriented toolingand fabrication concept.
A photograph of the complete blade is shown in Figure 2.Details of the root end are shown in Figure 3.
The rotor blade structural concept consists of a closed fiber-glass "D" spar terminating in a multiple wraparound root endretention system. The aft fairing uses fiberglass over aNomex honeycomb core. Erosion protection is provided by aticanium nose cap with a nickel leading edge at the blade tip.A pneumatic delta pressure failure detection system isemployed within the "D" spar. This, and the inherently slowcrack propagation of fiberglass and the multiple load pathdesign, provides for long-term detection capability and over200 hours of safe operation after detection.
The radius of the blade is 46 feet7 the chord is 40 inches.The airfoil sections start with the V43015-2.48 at the rootcutout (.25R) which transitions to the V43012-1.58 at 0.4R.This transitions to the new VR-7, which extends from 0.5R to0.85R. The VR-7 transitions again to the VR-8 at the tip.
The rotor characteristics are described in Figure 4 and Table1 for the ATC and the Prototype blade configurations.Manufacturing development and structural testing performedduring the ATC program led to modifications improving the
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design and structural capability of the Prototype blade. Themost important oL these were a precured spar Leel to improvelayup operations and to eliminate wrinkling of the fiberglassstiffeninq of the heel web to increase the aft fairing honey-comb core strength, and a titanium nose cap using highly di-rectional material with reduced cost and improved fatigueproperties. The weight of the ATC blade is 760 pounds andthe prototype blade is 774 pounds.
Demonstration tests verified the design predictions and metthe design objectives.
* A rotor hover efficiency with a figure of merit (FM) of.767 was demonstrated by wind tunnel and whirl towertests compared to the design objective FM of .751.
* Structural tests demonstrated essentially an unlimitedlife for the blade fiberglass spar and root attachment.
e Absolute fail safety of the blade was demonstrated bystructural tests. The root end is capable of sustainingat least 172 hours of high-speed level flight loads withone of four attachment lugs failed. An outboard airfoilsection with the titanium nose cap failed was subjectedto 427 hours at level flight loads, and 109 hours atmaneuver loads without degradation to the remainingfiberglass spar.
* The fatigue strength of the titanium nose caps testedwas lower than the design objectives due to defectivematerial and processing. However, the fatigue strengthwas still greater than level flight stresses, and afatigue life in excess of 1000 hours is predicted forthe prototype helicopter mission. Because of the demon-strated fail-safe characteristics of the composite rotorblade, cracking of the titanium nose cap is not consid-ered to be a flight safety issue.
Airv)rthiness of the blade for flight on the HLH Prototypeaircraft was proven in this HLH/ATC program.
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Figure 3. Root End View of Completed Blade
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TABLE 1. ROTOR SYSTEM DESCRIPTION
Rotor Diameter 92 Ft.
Blade Chord 40.0 In.
Blade Twist (Aerodynamic) -120
Normal RPM 156
Torque Offset (Lead) 4.5 In.
Articulation Hinge Radial Location 26 In.
Blade Attachment Radial Location 66 In.
Damper Arm at Station 66 10 In.
Pitch Axis 25% Chord
ATC Prototype
Rotor Blade Weight:Outboard of Fold Pins Sta.66 740 774 LbOutboard of Elastomeric Bearing 1,131 1,180 Lb
Rotor Blade Static Moment about 17,325 17,960 Ft LbHinge (Including Pitch Housingand Loop)
Rotor Blade Inertia about Hinge 15,640 16,141 Slug-Ft 2
(Including Pitch Housing and Loop)
Centrifugal Force:At Blade Attachment Sta. 66 153,139 158,324 LbAt Hinge Sta. 26 155,299 162,094 Lb
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'i.1
3.0 DESIGN DEVELOPMENT
3.1 DESIGN GOALS AND OBJECTIVES
The major design goals and objectives for the rotor bladecompared to the blade concept achievements are summarized inTable 2.
e Fail Safety
The multiple load path spar results in inherent failsafety since no single failure of a component willcause a catastrophic condition. The fiberglass sparfail safety is due to its high damage tolerance,insensitivity to defects and stress raisers, and softfailure modes. The pneumatic failure detection system,which consists of evacuating pressure in the enclosed"D" spar, provides additional assurance.
e On-condition Operation
The design objective to provide a blade which is fieldrepairable and maintainable and capable of "on-condition"retirement is satisfied by the failure detection system,a damage tolerant structure, repairable fairing, and areplaceable leading edge erosion protection at the bladetip.
e Reduced Maintainability Rates
Maintenance man-hours will be reduced by the detectionsystem which eliminates the need for special inspections.
* Improved Reliability Rates
A major reduction in blade malfunctions will be achievedby the use of the sealed Nomex honeycomb fairing core
which eliminates the extensive metal core-to--sparcorrosion problem.
e Reduced Blade Weight
The blade weight is minimized by the composite fiberglassand titanium spar and the performance benefits fromadvanced airfoil profiles.
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9 Improved Rotor Performance
Improved performance is acccmplished by tailoring theblade airfoil section, thickness ratio and twist ateach spanwise radial section. Variation of theseparameters is facilitated by the composite spar manu-facturing approach. The increased fatigue strengthof fiberglass compensates for the higher strainsassociated with increased airfoil thickness and twist.
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S, 22
.. . . .
TABLE 2. ROTOR BLADE DESIGN GOALS AND OBJECTIVES
GOAL/OBJECTIVE BLADE CONCEPT
Improved Safety * Composite fiberglass/titanium
Survivebility structure
0 Multiple load paths
e Failure detection system
On-Condition Operation * Failure detection system
* Multiple load paths
e Composite damage tolerance
* Replaceable tip nose cap
Reduced Maintainability * Failure detection systemRates
Improved Reliability a Composites
Rates a Sealed Nomex honeycomb fairing
cor a
Reduced Blade Weight * Composite fiberglass/titaniumspar
e Advanced airfoil
Improved Rotor o Advanced airfoil profilePerformance P Tailored secticns permitted by
composite blade
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I:I3.2 ROTOR BLADlE GEOMETRY AND SIZINGi
The use of fiberglass as the primary structural material in
the HLH/ATC blade permits the optimization of blade geometryto an extent not possible with extrudled metal spar blade.onstruction. Blade airfoil section, thickness, and twistcan be tailored along the span to provide the optimum aero-dynamic and structural blade configuration. This tailoringof geometry was first accomplished in the U. S. Army/DoeingAdvanced Geometry Blade (AGB) program. In the AGB, existingairfoil sections were employed along the span to provide anoptimum aerodynamic configuration. Figure 5 shows the AGBcompared to Chinook rotors. Planform taper and a low twistoriented to high-speed flight were also employed in the AGB.In the HLH program, advanced airfoils suitable to the particu-lar blade spanwise aerodynamic environment were developed andblended along the span to provide maximum lift and minimumdrag. The ILII blade, also shown in Figure 5, has a 120twist reflecting the desire for optimum hover performance andno requirement for very high-speed flight. The 121 twistprovides a 1,5 percent improvement in hover figure of meritover the Chinook 90 twist while increasing the blade bendingmoment in forward flight by 10 percent at the 150-knot VHdesign condition.
The HLH/ATC airfoil development program included two-dimensional airfoil developin-it, as well as 6-foot and 14-footdiameter model rotors. Two-dimensional wind tunnel data isshown in Figures 6, 7, and 8. Figure 6 is a compositeplot of airfoils which existed and either were already usedon rot.ors or showed potential for rotor usage. Examinationof this plot shows that no one airfoil provides maximum liftover the entire Mach number ranges (blade span) leading tothe conclusion that the optimization of rotor lift capabilityrequires a family of airfoils suitable for the range of Machnumbers encountered in the rotor environment. This led to theestablishment of the HLH objective shown in Figure 6,and an 11%improvement in CL max over the Chinook airfoil V23010-1.58 atMach number - .5,which represents the lifting areas of therotor on the forward and aft portions of the rotor disk.Figure 7 shows the results of the HLH/ATC airfoil develop-ment, the VR-7 and VR-S airfoils for working section and tipsection, respectively, and the V43012-1.58 in the inboardsection. Actually, a V43015-i.58 was ultimately employed
S~inboard for the best structural, as well as aerodynamic
configuration.
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i 5I1:I i I. . I I I iII I.. ! '
A maximum airfoil pitching moment objective was also estab-lished in order to assure that CL max would not be achievedat the expense of increased control loads. Figures 6 and7 show that the HLH/ATC developed airfoils satisfy theobjective and have lower pitching moments than existing highlift airfoils.
Measured drag data for the HTJH/ATC airfoils in Figure 8show a significant reduction in drag at all Mach numberscompared to the Chinook V23010-1.58.
The HLH/ATC rotor blade geometry is shown in Figure 9. Therectangular planform and squared-off tip were selected becausethey represented the simplest manufacturing approach andbecause a review of existing data showed that little furtherimprovement in blade performance could be achieved over thatpossible with optimum airfoil section, thickness taper, Y1twist.
A 40-inch chord was selected on the basis of chord/weighttrade studies and the CT/a requirement fo• the basic HLHdesign gross weight and alternate gross v?'ight configuraýx'sThe chord trade study results are shown in Figure 10 andindicate that a 40-inch blade chord could be provided with noweight penalty and with no significant effect on other bladeparameters over a 38-inch chord considered initially as theminimal requirement. The 40-inch chord results in an averageCT/U of .077 at the SL/950F design condition. This point isshown in Figure 11 relative to the CT/a limits which were
estimated for the HLH on the basis of 11% increased liftcapability over the Chinook. The over-load gross weightcondition is also shown to have adequate bank angle
capability.
25
I"
BLADE TWIST
'CH-47A 90
0012 AIRFOIL ROOT TO TIP
CH-47B & C 90
23010-1.58 AIRFOIL ROOT TO TIP
AGB 60
LINEAR LINEAR23012 TAPER TAPER
23010-1.58 13006-.7
SHLH 120
V43015 INEA _LINEAR
V43012 ITIOJ SITION
II ' 'I I
0 .1 .2 .3 .4 .5 .6 .65
HOVER MACH NO.Figure 5. Rotor Airfoils
26
TRAILING-EDGE DEFLECTION 00
NACA 64A(4.5)12(.1- 0.8)I.60.8) V23010-1.58
11% HLH OBJECTIVE
1.4. ' .
\ "-..NACA 64A312
MAXIMUM 1.2 V0012 " . C (a A3)LIFT "
COEFFICIENT, HLH'9 MAX OBJECTIVE
VR-1
0.8-
0 0.2 0.3 0.4 0 5 0.6 0.7 0.8
MACH NUMBER, M --
0 .- -- iv
VR--1 \ _ ,V2,o010-1. 58.'
ZERO,,FT -I.0 I -FT-- :-.PITCHINGMOMENT HiLi- OBJECTIVECOEFFICIENT,Cmo -. 08- r
-. 12 NACA 64-A(4,5)12
(a = 0.8)
Figure 6. Maximum Lift Doundaries and Zero LiftPitching Moment Levels
27
I. .. ..
2.0 .0 ---- -, V43012-1.58 HLH REYNOLDS NUMBERS
"TRAILING-EDGE DEFLECTION =00
S'• .,.o._ VR-71.6
1.2 , HLH OBJECTIVE
CQMAX
8 "R.8 V13006-.7
V23010*-1.58 .
.4 I ..
5 6 7 8 9 10 11 12
REYNOLDS NUMBER, io x 10-6
ci _-- _____I .-W -
.2 .3 .4 .5 .6 .7 .0
MACH NUMBEH. M
V43012 " .1.80 -0 ---
• V23010-1.58
-. 02 V13006-.7
ZERO LIFT - - VR-8
PITCHING MOMENTCOEFFICIENT, Cmo -,04
HLHOBJECTIVE vR-7
-. 06 L
Figure 7. Maximum Lift Boundaries and Pitching Moment Levels
28
lq
6TAB = 00
.015-
V23010-1.58
VR-7
S.01V43012-1.58- .010'
VR -8
LLuL .005
Uj
5 6 7 8 9 10 111 12
cc n-REYNOLDS NUMBER, Re X 10-6
1 Ii I I.3 .4 .5 .6 .7 .8
MACH NUM3ER, M
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1I, I
U VR-8U.
S5 6 7 8 9 10 11 12cc= REYNOLDS NUMBER., Re X 10
0
.4 .5 .F .7 .8
MACH NUMBER, M
Figure 8. Drag Comparison, Wake Probe Data (HLH Reynolds Number)
29
O .*NOO - , " ,
, •
*, I 4Q1
2.• ...
• *- 0 0 I .I
S18 1-4r
,II , * ,'
30)
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STATIC DRiOOP CAPABI L-11Y VS CHORD3g MINIMUM ACCEPI ABLE3.5g DESIRABLE
4.0-
0. 3.000
38 40 42CHORD - IN.
FIRST MODE - GRAPH/GLASS8.0- TORSIONAL FREQUENCY VS CHORD --- TI/GLASS1 5ýO MIN ACCEPTABLE - ALL GLASS
6.0- .
5.0
4.0j
38 40 42CHORD
0.7- % STRUCTURAL WEIGHT
WEIGHT TO TOTAL UNIT WEIGHT0.5 ~~~~AT .50 Sn A'sRl 'i L S.ElC r 10N'
38 40 42CHORD - INCHES
Figure 10. Chord Trade Study
31
*LEVEL FLIGHT AVERAGE C1T/o
.16 6ILH OBJECT{IVE
.14 -- WITH ADVANL.ED14 IRIFOILS
11% INCRES
.12 EASE
V23010 AIRFOIL 3' BANK
ANGLE.10 c-L 9 _ • %C P • L T 500,rk7 ,:_,.G .W . " P A I L B A N K
S-L• ANGLE
CAPABILITY
.08 917T-"- z
D.G.W 0
.06 i~l50 KNNOTS
•,, "' .34
TO,
0.0 20 .30 .40 .50
ADVANCE RATIO - .A
Figure 11. Rotor Flying Qualities Boundary -
Steady Flight Condition
32
NI
3.3 ROTOR BLADE STRUCTURAL CONCEPT
The HLH composite rotor blade shown in Figure 12 uses a
unidirectional and crossply fiberglass closed "D" spar as theprimary structural element. The external surface of the sparis covered by a titanium nose cap bonded to the fiberglass.The nose cap provides erosion protection and torsional andbending stiffness. Replaceable erosion protection in thehigh-wear area at the tip of the blade is given by nickelcovering the leading edge of the blade.
The root end attachment features an all-fiberglass wraparoundconstruction in which the spar unidirectional fiberglassmaterial is layed in equal packs from the tip to the root andsymmetrically back to the tip. This feature is illustratedin Figure 13.
The aft fairing is a single box with fiberglass skins and aNomex-honeycomb cote. A pneumatic failure detection isinstalled for fail safety. The blade concept describedevolved from initial design studies and support testing, andis the design fabricated for structural,whirl tower andDynamic System Test Rig demonstration tests and that plannedfor the Prototype HLH.
The above concept was seleted as a resu.t of the preliminarydesign studies and supporting tests discussed in the follnw-ing paragraphs.
33
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3.4 PRELIMINARY DESIGN STUDIES AUD SUPPORT TESTS
The initial blade design concept utilized the fatigue strengthof fiberglass and the high stiffness, and strength-to-weightratios of graphite to form a dual load path compo'.ite "C"spar. The skins were of crossply graphite because of itshigh modulus of rigidity. Design analysis showed that thisconcept satisfied the design fail-safe objectives. The rootend attachment utilized the "coke-bottle" method, which alongwith the "C" spar was successfully demonstrated with all-fiberglass and all-boron advanced geometry rotor blades onthe CH-47 Chinook helicopter. Dual and single spar designswith anl .-ithout condition indicating systems were consideredas illustrated in Figure 14. ) The initial design tradestudies are described in Reference 1.
Design support testing was conducted to evaluate the strengthbehavior of mixed modulus fiberglass/graphite composites,damage tolerance, impact tolerance, failure detection systems,and erosion. The results of these tests and further designstudies had a conside~cable impact on the blade design, andfinally led to the blade structural concept changing from aglass/graphite spar with graphite skins, crack wire detectionsystem and polyurethane/nickel erosion strip to an all fiber-glass spar, pneumatic detection systev and a titanium rtickelnose cap erosion system described in Paragraph 3.1. Thedesign support test results are documented in the HLH/NTCProgram Quarterly Summary Reports -2 and -3 (Reference 2).
The major reasons for the change in the structural conceptare: first, metal was the only material known that would satis-fy the requirements for erosion protection; second, the rapidfailure mode of graphite was undesirable for a primary sparmaterial; and third, the metal and fiberglass construction wassuperior in impact resistance and damage tolerance.
The addition of a metal nose cap provided inherent lightningprotection and improved reliability of the deicing system.With the elimination of graphite crossply, the metal nosecap supplied the necessary torsional stiffness.
The design support tests which led to the ATC blade struc-tural design concept are described in the following sections.
36
I u ni n n nu n ni U "l "I m
3.4.1 Phase I - Material Coupon Testing
The Phase I Material Program included the test of mixed modu-lus and single modulus laminates, sandwich beams and tubes.The results from 216 test specimens were used to establishthe following conclusions:
1. Static modulus, static strength and fatigue stzength ofthe graphite generally met all design goals.
2. Graphite exhibited extremely brittle failure modes.
3. Graphite crack propagration rate was rapid, both staticallyand in fatigue.
Figures 15, 16, and 17 show representative specimensand clearly point out the brittle nature of graphite failurescharacterized by rapid transverse cracks in various locationsin the specimen. Failure across the fibers was expected,but not with the rapidity and severity demonstrated in thesetests. The fiberglass failed as expected with longitudinalsplits developing and failure progressing slowly to a pointwhere the glass could no longer react the load. Anothe'r con-clusion from this testing is that the fiberglass can act asa redundant load path.
3.4.2 Aft Fairinq Damage Tolerance Test
In order to assess the effect of foreign object damage andfield handling damage in service on high modulus graphitelaminates, such as would be used for the fairing skins aftof the spar area, a series of test specimens were fabricatedand evaluated. This is considered to be extremely importantsince aft fairing damage has been the cause for many bladeremovals in service.
Impact test parameters were the same as those used earlier toassess potential field damage- to production units and con-sisted of gravity impact using a 2-inch diameter, one-poundball dropped from increasing heights up to a maximum of 10feet.
Significant differences in skin and core materials behaviorwere found at the 10-foot impact level. The results aresummarized below:
37
Il
1. No glass skin fracture occurred in any of the tests.
2. Aluminum core speuimens sustained more damage than didthe Nomex core specimens with the maximum impact resultingin a 'permanent set" depression.
3. The graphite laminates were fractured when impact balldrops of more than 5 feet were used, with the damagebeing limited to less than the ball diameter in area.
Figures 18 and 19 compare the typical damageresul.ting from a one-pound ball dropped from heightsvarying from 6 inches to 10 feet on to aft fairingsections with Nomex honeycomb core and crossply fiber-glass and graphite skins.
3.4.3 Whirling Arm Impact Testing
A whirling arm impact test was conducted on IR&D funding.The test was conducted by whirling typical blade sections atfull-scale tip speeds and then introducing into the tip pathplane hard-wood dowels of varying diameters. The test wasbased on a linear scaling principle with blades being scaledon an equal weight basis. It was meant to be purely quali-tative and for comparative purposes only. The results of thetest for both 3½-inch and 18-inch chord blade sections indi-cated the excellent damage tolerance of a metal fiberglassconstruction. In these tests, also, the brittle failure modeof graphite was evidenced by the loss of large areas of thetrailing-edge fairing of all impacted specimens.
3.4.4 Crack Wire Failure Detection System Test
Failure detection (crack wire) tests were performed on nine20-inch beam specimens. There was a total of 112 imbeddedwires in these specimens, of which 49 were lost due to han-dling damage or premature failure. Of the nine beams tested,four failures were detected. The results of these testsindicate that: (1) the failure mode of graphite makesdetection by crack wire adequate; (2) the failure modes offiberglass make detection by crack wire undependable; (3)crack wires are highly susceptible to handling damage duringthe manufacturing processes of laminate fabrication, and(4) repair of failed imbedded wires is not possible.
38
[ ___
3.4.5 Erosion and Lightning Protection Study
Concurrent with these material evaluations, studies were con-ducted to investicate erosion protection and lightning protec-tion systems and how their requirements would be affected byblade construction and materials.
Investigatio' of polyurethane, which was the primary erosiont-nci ection system for the inboard 85% radius of the proposedb~i indicated that it would be satisfactory for a pure sandz&dxCsfl environment. However, the rain erosion capability ofurc;:n•.. especially after sand exposure, is very low and did:,ot show promise of being improved in the near future. Poly-urethane was selected originally and carried outboard to 85%radius to avoid subjecting the metal erosion strip to thecritical strain zones on the blade (60%-P9% radius). While itwas known that urethane was inadequate for rain erosion atfull-tip speeds, available data indicated that it was satis-factory to 85% radius. However, after sand exposure, therain resistance of urethane was degraded to make it inadequatefor the HLH blade. Metal at Lhe leading edge of the bladewould be required from 40% radius outboard.
The requirement for lightning protection is that the rotorblade must sustain a 200,000 amp strike without a catastro-phic loss of blade, either aerodynamically or structurally.It is assumed that lightning can strike any part of the bladewith a 90% probability at the tip. To satisfy these require-wlents, the leading edge has to have conductance equivalent to21,000 circular mils of copper from tip to root. Thislightning protection can be achieved inherently with a metalleading edge of sufficient cross-sectional area or a copperrod from tip to root and a metal tip cap connected to all
other metal in the blade. The cross-sectional area for
Sequivalent conductance is .?164 inch2 for copper, 1.64 inch 2
for Ti-6A1-4V and .562 inch for 301 stainless. Graphite orfiberglass aft aerodynamic fairing skins have a probable lossof 2-3 feet 2 without protection. However, undetected damagecan occur in graphite without wire mesh protection (5-10 ampscan damage graphite fiber). Therefore, a glass/graphiteblade must have, in addition to a copper conductive wire, afine grid aluminum weave over the entire blade.
39
I
with condition indicating system
CONCEPT Bwithout condition indicatingsystemk ."
(a) Dual Fiberglass Graphite Spar
CONCEPTC with condiLion indicating system
CONCEPT D
without condition indicating system
SINGLE FITTING
(b) Single Fiberglass Spar
Figure 14. Trade Study - Rotor Blade Concepts
40
V4-r 5207
a 4 * ...- v.........._ _ _ _ _ _ _ _ _
HT 5Žr 07C.C.
Figure 15. Tension Fatigue Failure of Graphite Composite(Room Temperature)
41
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Best Available COPY
P11 - I. 7,
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42'
Figure ~ ~ ~ ~ ~ ~ ~ Bs 16.lbl TesoCatgeFilrPfGrpieCopst
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Best AvaiIab�e CopY
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Figure 17. Shear Fatigue Failure of Graphite Composite
43
Figure 18. Impact Test - 1 Lb Ball on + 450
Graphite and Nomex Core
Figure 19. Impact Test - 1 Lb Ball on + 450
Glass and NoMex Core
i ~44
i
3.5 DETAIL DESIGN OF THE ATC BLADE CONFIGURATION
The initial design studies and supporting tests led to theselection ot the composite metal/fiberglass blade concept.Prior to proceeding with the detail ATC blade design,the bladeparameters that were most affected by the deletion of thegraphite material were reexamined. These included the bladefail safety, torsional rigidity, static droop clearance, andthe failure detection system. The torsional rigidity andstatic droop clearance requirements were met by extendingthe nose cap coverage over the complete -Žxternal spar upperand lower surfaces. This is due to the fact that the metalinherently provided increased torsional and flapwise stiff-ness and minimized the amount of crossply fiberglass. Agreater proportion of the spar weight was then available forunidirectional fiberglass which yields better fatigue strengthand droop prcperties.
The slow failure mode of fiberglass, the primary blade mater-ial, makes failure detection by crack wires undependable, anda differential pressure system was considered to be afeasible alternative.
With the basic concept defined, thu detail blade design wasentered into with a high level of confidence that successfuldevelopment could be accomplished during the ATC program.
The detail design phase was also supported by a considerablenumber of tests, as summarized in Figure 20. Each majoritem of the blade was evaluated before finalizing the designand proceeding with the blade manufacture.
3.5.1 Spar
The spar subassembly is a closed "D" cross section of uni-directional and crossply fiberglass composite.
The change from "C" spar to "D" spar was precipitated pre-dominantly for fabrication considerations. The tooling con-cepL utilized for fabrication of the "C" spar for the CH-47AGB fiberglass and boron blades required three major toolsand five separate blade "cooks." This process was expensiveand not production oriented. A "producible" AGB tool concept.still had three major tools but required only three separatecooks. Major "C" spar fabrication problems were: coretolerances, core insertion, skin-to-spar and spar-to-core
45
- I
bondin;. Since the "IC" spar is cured as a separate entity,and the coic must be stabilized prior to insertion into thespar, the spar inside mold line and the core outside moldline must be held to a very close tolerance if they are tofit and provide a good quality bond. Hence, a tolerance andpressure problem exists in the nose of a "C" spar which isdifficult to overcome. A similar problem occurred in the'H-47 AGB when the precured skins were bonded to the precuredspar-core assembly. Due to the lack of flexibility in themating surfaces, there were large unbonded or void areas overthe span.
The 'D" spar concept reduces the required number cf majortools to one (with supplementary inserts) and coapletelyeliminates the major core/skin-to-spar bonding problems. Thefabrication sequence involves fabrication of a complete spar(titanium and fiberglass) in one cure. This spar will bevirtually void-free and will ensure a superior bond to thetitanium and a repeatable close tolerance airfoil contour.The aft fairing can be bonded to the fully inspected spar asa precured subassembly (which is similar to the fabricationof present day metal rotor blades) or bonded to the spar whilethe fairing is being cured. This latter approach was used toelimInate close control tolerances between two precured com-posite laminates. The "D" spar is the best construction forcost because of these fabrication and tooling advantages.Cost will be further reduced because of better inspectability,better repeatability (tracking considerations), and the capa-bility of replacing the entire aft fai i.ng assembly.
Other important advantages of the "D" spar over the "C" sparis that it gives the highest torsio•nal stiffness per pound ofblade weight and is adaptable to an ISIS-type pneumaticdetection system.
Titanium was selected for the metal nose cap because it hasthe best fatigue strain capability as shown in Table 3.in addition, its creep-forming capabilities are ideally suitedto the complex leading edge of the blade and its erosion pro-perties are better per pound than 301 stainless steel andsurpassed oniy by nickel.
Titanium possesses a 50-percent better saecific torsionalstiffness than does fiberglass crossply, and is equivalsetto unidirectional fiberglass in specific bending stiffaess;
.1 therefure, torsional stiffness can be achieved with titanium
46
-
without sacrificing droop characteristics. By extending thechordwise coverage of the titanium to .35C, the necessarytorsional stiffness of the blade was achieved with a slightweight reduction compared v.o all fiberglass construction.
The thickness of the titanium was determined by torsionalstiffness and erosion considerations. The maximum amount ofunidirectional fiberglass was used in the section cL-imensur-ate with bending stiffness and blade weight requirements.Fiberglass is the basic spar material with its high fatiguestrength, damage tolerance, and "soft" failure modes. Thespecific stiffness of fiberglass is equivalent to that ofmetals; consequently, for the same weight, the blade has thesame stiffness ana frequencies as a metal blade. However,fiberglass has a fatigue strain capability 2.5 times that ofsteel, 3.3 times that of aluminum, and 1.6 times that oftitanium.
The blade alternating strains in steady-flight conditions aremuch further below the endurance limit compared to the othermaLurials. The larga fatigue margin is beneficial for damagetolera.nce ,nd in-service reliability.
The length of the blade eliminated the use of MIL-T-9046 hotrolled 6AL-4V titanium alloy sheet, as this is available inmaximum lengths of 20 feet. Therefore, cold rolled 6AL--4Vsheet to Boeing Specification BMS-7-197 (Reference 3) wasselected. This material was coupon tested to provide fatiguestren. '-h ;:-operties for strength analysis and to evaluate edgetrezAtmenus and the effects of heat treatment and processingrequired for the nose cap forming. The results of these testsare shown in Figure 21.
The endurance limit established from the coupon tests showedtii-anium to be acceptable from a fatigue strength point ofv:. ew.
The use of the hollow "D" spar caused the wall buckling tobe a critical design condition. As a result of the bucklinganalyses, the spar wall thickness was increased withadditional unidirectional fiberglass from .25R to .75R. Abuckling test of a representative spar section was conducted,which confirmed the analyses. Consideration was alsogiven to the deflection of the spar wall, in the chordwisedirection, due to differential pressure from airloading andthe pneumatic failure detection system. The deflections were
47
limited so as to be equivalent to those for the CH-47B/C
helicopter for the same conditions. This was done by addingunidirectional graphite running chordwise with its highspecific stiffness at 900 to the unidirectional fiberglass.The graphite was embedded in the spar inner crossply fiber-glass and the unidirectional fiberglass so as to protect itfrom impact damage.
Although there was no requirement for fabrication of a de-icer blanket in the ATC program, design support tests wereconducted to evaluate the effect of the local blanket tempera-ture on the surrounding fiberglass and adhesive structure.Actual energized blanket tests in a realistic ambient environ-ment and duty cycle indicated thai. the critical bondlinetemperature never exceeded 1100F, and that the blanket locatedbetween the titanium and fiberglass would pose no structuralproblem. Nesting the blanket between the metal cap and thefiberglass assists the deicing process. The heat generatedin the blanket flows outward to the iced surface and is notabsorbed in the blade due to the insulating characteristic ofthe fiberglass.
3.5.2 Root End
The change from the "coke" bottle root end to the wraparoundroot end (Figure 22) culminated a design and analyticaltrade study effort spanning more than one year. Althoughobviously redundant, the dual coke-bottle configurationoriginally proposed was very expensive to fabricate and wouldhave been extremely difficult to protect with any type ofdetection system. Protection of this metal root endconcept would have been mandatory. Further, the disadvantageof having metal components built into the laminate which canbecome potential fatigue problems, and which are not replace-able, reduces the blade's serviceability and increases its
potential cost.
The wraparound root end is redundant since it has four separ-ate load paths into the hub. It has no metal componentsbuilt into the laminate. The metal bushings for the attach-ment hardware are all replaceable. Since all lugs are sep-arated and exposed, visual inspection is all that would berequired for fail-safety. Furthermore, the w:raparound rootend concept is the lightest of all the concepts studied indetail. The cost of metal machining for the root end iscompletely eliminated.
48
VI
A root end design support test (Reference 4 ) was conductedwhich demonstrated the concept's feasibility. Fatigue loadingat amplitudes three times areater than VH high-speed levelflight was sustained for 3 x 106 cycles without failure ofthe primary load-carrying members of the root end. The speci-men endured an additional .92 x 106 cycles of fatigue loadingat three times VH flapwise bending load and maximum lag damperload applied through a lag damper arm at the blade attachmentlocation, Station 66. Bushing fatigue failures created arequirement for a design revision. Sleeves and separatewashers replaced the bushings at the blade retention pins.ISIS leaks around the lag damper arm indicated a need for asealing bulkhead inside the spar.
The ;:oot end was capable of reacting overspeed rpm, flightmaneuver, starting, and ground flapping limit load conditionswithout any apparent damage. The fail-safe testing showedthat the root end is capable of sustaining V11 high-speedlevel flight loads with simulated failures at various spanwiselocations, in three of the four load paths. Axial load equalto the design limit centrifugal force of 250,000 pounds wascarried by the root end in the simulated failed condition.
3.5.3 Aft Fairing
The aft fairing is a single box construction, with fiberglassskins, and Nomex honeycomb core. The fairing subassembly isshown in Figure 23. Fiberglass was selected as the skin materialbecause of the damage tolerance and durability demonstrated inyears of service on the CI-1-47B and C helicopters. Fatiguetesting of fiberglass blade skins returned from service demon-strated that fiberglass properly protected by paint showslittle effect from the service exposure. Fatigue tests ofthe used skins fell within the scatterband of new, unexposedskins as shown in Figure 24.
The Nomex honeycomb core was selected because as a nonmetalit eliminates the corrosion problems experienced with metalhoneycomb. Nomex also provides a substantial benefit in theblade's fabrication concept. When enclosed between two moldsto a fixed dimension, Nomex deflects and provides a backpressure proportional to the deflection. As the temperatureincreases during the cure cycle, the back pressure decreasesto zero and the Nomex sets in this deflected position withlittle or no spring back.
49
Fatigue, static, moisture penetration and migration testsof fiberglass/Nomex specimens were conducted and are reportedin Reference 5.
The trailing-edge wedge and cusp are of fiberglass/graphiteconstruction. The section is constant from Station 138 tothe blade tip except that there is a 2-inch cusp extensionbetween Stations 138 and 220.8. Ninety-degree uni-fiberglassis provided for chordwise trailing-edge stiffness in order tominimize in-flight cusp deflections. The zero-degree uni-fiberglass and HT-S graphite is sized by trailing-edge buck-ling considerations for rotor starting conditions and by theblade chordwise bending stiffness and natural frequencyrequirements.
The cusp stiffness limits the deflection of the cusp, as ittravels around the azimuth, to +.2 inch under external loadingconditions.
The first chordwise natural frequency reduces by approxi-mately .15 when coupled with the drive system. The size andstiffness of the trailing-edge wedge was determined so thatthe coupled frequency was greater than 4.5 per rotor revolu-tion to ensure that, with a 4-bladed rotor, unfavorable 4 perrevolution vibrations would not be transmitted to theairfrfre.
3.5.4 TiR Installation
The tip assembly for the rotor blade provides for a largeadjustable weight capacity. The tip provides the capabilityfor moving the dynamic balance axis forward approximately.75 percent. Advantage has been taken of the more aft locationof the VR-7 and VR-8 centers of pressure by allowing the localchordwise balance axis (and, therefore, the resulting dynamicbalance axis) to fall aft of the conventional quarter chordlocation. Wind tunnel testing has shown that no flutterexists for this configuration, but the .75 percent over-balance capability has been provided as a precaution.
The tip fittings were chopped fiber molding to be precured
and secondarily bonded into the bonded blade subassembly(spar and aft fairing subassembly).
50
3.5.5 Fiberglass Y'aterial Specification
The fiberglass resin system for the HLH blade is SP250-1014S.The CH-47 AGB fiberglass blade was built using an SP1002S resinsystem curing at 350 0 F. The SP250 system which cures at 250'Freduces tooling and fabrication costs. It permits faster curecycles with less heat-up and cool-down times, lesser heatrequirements, and reduced warp in the co-cured spar. Couponfatigue test results (Figure 25) showed that the SP250resin system was at least as strong as the 1002S system.
3.5.6 Location of the Chordwise Balance Axis
Blade design practice places the ce:iter of the blade mass (Bal-ance Axis) aft of the airfoil aerodynamic center, which forconventional airfoils is at .25C. Figure 26 presents the aero-dynamic center for each of the HLH airfoil sections. Theadvanced airfoil permitted the balance axis to be as far aftas .26C, which resulted in a considerz.ble weight savii.g. The14-foot-diameter model rotor was tested in the Boeing VertolWind Tunnel with the balance axis at .257C without evidence ofblade flutter.
3.5.7 Failure Detection System
The failure detection concept for the titanium/fiberglass"D" spar is a pneumatic differential pressure systemutilizing an evacuated spar. Because of the very long lifeafter titanium failure, the pneumatic system will protectonly the fiberglass portion of the spar, and a failure ofthe titanium nose c p will be detected visually.
For normal operation of the titanium and fiberglass actingtogether in the spar, it is not conceivable that the fiber-glass could ever fail before the titanium. In the event offiberglass damage during manufacture or service, it is stillhighly improbable that continued deterioration or propagationof the damage in the fiberglass would result, without causinglocally higher straining of the titanium to the point whereit would fail locally permitting a leak and subsequent fail-ure indication. Tests have confirmed these conclusions. Thetests of glass composites, in combination with steel andtitanium have included undamaged specimens, specimens withprior damage to the metal, specimens with prior damage to theglass, and specimens with simulated bullet damage to both the
51
metal and glass. In all cases, the metal failed first andthe damage did not propagate to the glass.
Results of design support tests with a simulated soar section(Reference 6) and with evacuated elliptical fiberglass tubes(Reference 7) established the feasibility of the applicationof a pneumatic system to composites.
In order to ensure failure of the fiberglass at the bladeoperational stress levels, defects were built in the fiber-glass, producing a spanwise discontinuity of the unidirectionalfibers. The specimen section properties at the defect loca-tion were reduced by approximately 20 percent. Specimenfailures occurred as titanium cracks, debonding between thetitanium and fiber-glass, and propagation of the built-indefect. In all cases, the failure was identified by avacuum leak. Most failures occurred under the beam load clampand were primarily due to the method of loading which producedhigh shear and local secondary stresses. The specimens wereconsidered to be failed when the beam deflections became solarge that loads could no longer be applied. At the termina-tion of each test, all specimens were capable of carrying thetest axial load equivalent to the rotor blade centrifugalforce.
The conclusions obtained from the tests were:
1. Failure of the fiberglass spar under the titaniumnose cap would induce a titaniun. failure or debondingbetween the titanium and the fiberglass, thus providinga vacuum leak path.
2. Following the vacuum loss indication, the bladestructure will be capable of supporting normal flightloads for fatigue cycles equivalent to at least 200 hours.
3. Fiberglass laminates do not inherently leak while underhigh vibratory strains, a necessary requirement for thepneumatic system.
4. Fatigue failures of fiberglass laminates progresslocally through the thickness of uni and are accompaniedby sufficient ,•rossply delamination to permit leakage, anecessary requirement for the pneumatic failure detectionsystem.
52
IrlI
5. Small defects x•Ill not propagate in fiberglasn at strainsof 1000 inc'hes or less.
6. The testing showed that even a large fiberglass defectwould at first cause a titanium failure and delamination,and that fiberglass failure propagation is extremely slow.
Slow propagation was not a requircment for the detectionsystem and no pror-gation time was assumed in the developmentof the 200-hour at (M -2o ) endurance limit criterion. It wasassumed that the remaining structure, after the failure,be ofsufficient section to sustain 200 hours withotut strainsexceeding the (M -2a ) level. The inherent crack propagationcapability of fiberglass makes this criterion much more con-servative than originally anticipated, since initial evalua-tions indicate that the available detection time will exceed200 hours.
3.5.8 Erosion Protection
Whirling arm erosion testing of nickel titanium, and stainlesssteel were conducted at full-scale tip speed (750 fps) andaccelerated sand densities. The results of these tests aregiven in Reference 22. They show that the nickel/titaniumleading-edge system is a substantial improvement over thestainless steel leading edge of the CH-47.
3.5.9 Blade Drawings
A complete list of the blade drawings is given in Figure 27.The blade assembly drawings are included as Figures 28 through31.
53
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3.6 DETAIL DESIGN OF THE PROTOTYPE BLADE CONFIGURATION
The basic structural concept for the blade for the prototypehelicopter is identical to that of the ATC configuration.Rotor blade design improvements that developed out of manufac-turing experience and demonstration testing of the ATC con-figuration were incorporated into the HLP prototype design.These modifications, sumnarized in Figure 32, include the
following items:
1. Lightning protection2. Titanii~m nose cap material substitution3. Tip fitting installation and hardware4. Precured spar heel5. Lag damper arm and sleeve6. ISIS integral spar inspection system7. Internal droop stop wedges8. Outboard spar wall stiffener9. Aft fairing core and skin
The purpose of these changes was Lo reduce the manufacturingcost and to improve the structural capability of the rotorblade.
3.6.1 Lightning Protection
Lightning tests on an ATC rotor blade segment indicated thatthe titanium cap is almost an order of magnitude better thanpredicced in its ability to transmit lightning. Consequently,the coverage provided by the ATC titanium nose cap is muchmore than is needed to transmit a 200,000 amp strike.
It was anticipateC that th.e titanium nose cap would be moreattractive to lightning than the graphite trailing-edge wedge,but this was not the case. Since the graphite acts as a con-ductor, two issues must be addressed as a result of thissituation. The graphite suffers somne microscopic damage whencurrent flows through it (the fiber-to-resin bond breaks down)which would result in a loss of some utrength and stiffnessin the composite. Since the graphite trailing edge was notgrounded in the blade design, the lightning must arc from thetrailing edge to the spar and would choose the path throughthe aft fairing, which would produce internal damage to theblade.
85!
I I II I I l i
The first issue, loss of stiffness in the trailing-edge, isnot safety of flight since the trailing edge wedge has threeindependent unidirectional load paths. Any damage to thewedge, even undetected damage, would result in a change inin-plane stiffness, and a vibration level niiange would occur.
The second issue, trailing edge to spar arcing, could presenta situation where safety of flight would be affected. There-fore, it was concluded that graphite in the trailing edgemust be grounded and protected.
The pigtail arrangement used to ground the inboard end to thetitanium is expensive and requires excessive processing tomake tb' brass-to-titanium-to-wi-e termination.
During lightning testing, an aluminum sheath or covering wasplaced over the trailing edge -rea. This new conductor suc-cessfully shielded the trailing-edge wedge from the remaininglightning strikes. A complete Faraday c-ge is provided by aweave of aluminum. Figure 33 shows the modified lesigncompared to that for the ATC blade.
3.6.2 Titanium Nose Cap
The original Specification (BMS7-197) for the nose cap mater-ial required minimum differences between proprrties in thelongitudinal and transverse directions. Preliminary testsindicated that when the directionality (texture) of thematerial is pronounced, improved high-cycle f.'-i.gue strengthproperties are obtained in the longitudinal direction. Fur-ther tests, reported in Reference 8 , confirmed this phe-.nonemon and showed that the heat treatment for forming thenose cap has no degrading effects. The highly directionalmaterial was selected for the prototype nose caps to takeadvantage of the higher strength. The highly directionalproduct is also easier to fabricate and should prove to bea substantial cost saving.
3.6.3 Tip Fitting Installation and Hardware
Obtaining an acceptable fit of the precured tip weight fittingto the inside mold lines (IML) of the spar proved to be a verydifficult and ex *nsive procedure. The difficulty in obtain-ing an exact fit resulted in t:uestionable bond integrity.
86
_I---
The internal hardware of the ATC tip-weight configurationrequired many parts that were difficult to install. ISISpressure leaked through the upper and lower precured halvesof the tip fitting into the tracking tubes, causing an invalidfailure indication.
The prototype configuration reduced the number of tip hardwareparts and thus simplified the installation. A unidirectionalfiberglass fitting was co-cured with the spar to eliminate theclose fitting requirements and to eliminate the ISIS leaks.This co-cured configuration also reduced the cost of the tipfitting assembly.
3.6.4 Spar Heel
Sporadic wrinkling of the crossply fiberglass in the ID" sparheel area occurred on the ATC blades manufactured (Figure 34).The wrinkling is unacceptable structurally as it caused anearly failure of a spar section during a limit torsion testas reported in Refere.ze 9.
The wrinkling originated during the installation or transferof uncured composite material into the curing mold. Thesolution to this problem incorporated in the prototype designis to precure the heel (Figure 35) as a structural memberin a separate operation prior to the spar assembly.
3.6.5 Laq Dgmper Arm and Sleeve
The root end demonstration test showed that the stresses inthe lag damper arm were higher than calculated. The highstresses caused a failure at the trailing pin hole. Thisfailure was due in part to improperly applied test load. Thefailure origin occurred at fretting between the steel sleeveand the titanium damper arm. The sermetal coating on theinner diameter surface of the sleeve was unsatisfactory foreliminating fretting, showing wear that progressed all the waythrough the coating.
For the prototype design, stress levels were reduced byincreasirg the thickness of the lag damper arm in the criticalarea around the trailing pin hole. Improved fretting protec-tio" was piuvided with a fiberglide coating applied to theinner and outer diameter surfaces of the steel sleeves. Thefiberglide was proven in the CH-47 socket where it was sub-jected to operating bearing press•.res similar to those of the HLH.
87
!II
The number two HLH root end specimen with the new arm and thefiberglide fretting inhibitor was tested at high-speed levelflight (VH) loading for 258 equivalent flight hours. Frettingbetween the steel sleeve and the titanium damper arm waseliminated and the fiberglide on the sleeve was in excellentcondition.
3.6.6 ISIS Integral Spar Inspection System
The root end ISIS bulkhead for the ATC design was located atStation 80, outboard of the chopped fiberglass internal droopstop wedges.
The installation was difficult to inspect and/or repair oncethe sleeves and damper arm were installed. Reconfiguration ofthe droop stop wedges permitted the relocation of the inboardISIS bulkhead to Station 70. The mounting block was eliminated.thereby reducing the number of parts and the overall cost.Repositioning the valve away from the indicator makes theevacuation system failsafe. The weight of this installationis less than the ATC. The installation is more repairablewithout root end disassembly.
'filie initial ly spccified inLernal pressure of 3.5 psia for theevacutLcd :;pzi- was sc.Lcccd on the basis of metal blade ex-pcricnu,- wher-c, because• of the rapid crack propagation andthe tue(I o OoIect crac:k longtLhs of approximately0..10 inch, a completely active system is required. A com-pletely active sysLem has a pressure set to always prcvide adifferential pressure between internal spar and external airfor all fliqht conditions from sea level at -65'F to 8000feet at 100 0 F.
For the prototype blade, an intermediate internal, pressure of7.5 psia was specified based on the following characteristicsof composite rotor blades:
1. Very slow damage propagation.
2. Residual strength of section with extensive damage whichwould obviously leak on the ground or in the air stillprovides 200 hours of safe life.
The 7.5 psia pressure provides a differential pressure forall ground conditions between sea level -65 0 F and 8OGO feet at100*F. The fail-safe test data obtained for the composite
88
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blade failure mode and rate indicate that the continuous ISISsystem could be replaced with a periodic "pump down" groundcheck and still retain the required fail safety.
3.3.7 Internal Droop StoDs
The ATC design used four fittings to react the compressiveloads at the root end due to ground conditions with zero orlow blade centrifugal force. Each fitting was hot bonded tothe internal surface of the spar. The variation in surfacecontour required considerable hand fitting of the blocks priorto bonding. The root end structural tests showed the hotbond to be unsatisfactory. An interim fix using cold bondedEC-2216 fittings capable of receiving the design loads wasused on the whirl tower and DSTR blades. For the prototypedesign, the droop fittings were cured in place with the spar,eliminating the fit and bonding problems experienced.
3.6.8 Aft Fairing
Simulated airloads testing of the ATC airfoil sections re-sulted in premature shear failures of the Nomex honeycomb coreat the bond of the core to the spar heel. The results ofthese tests are reported in Reference 9. The prematurefailure was attributed to the core height and to deflectionof the spar Peel Full-size coupon tests verified the heightand stiffness effects. In addition, the tests showed thatcuring temperature and crushing of the core during assemblydid not degr'<e the core strength.
The prototype design substantially increased the stiffness ofthe spar heel by adding graphite into the heel web at 900 tothe spar direction. Figure 36 compares the ATC and protypedesigns of the heel.
Repeat of the simulated airload tests (Reference 9) verifiedthat the modification met the design load conditions for theprototype helicopter. In addition to the spar heel stiffening,a horizontal stabilizer was introduced into the outboard sec-tion of the prototype fairing to improve its strength. Thecore density of the intermediate section of the prototypefairing was increased to 3 pounds. The 2-pound core behindthe vertical splice was introduced as a weight saving scheme,since the increased strength is not required in this area.
89
These changes are illustrated in Figure 37. The total weightpenalty for the prototype core modificaLion is 7.5 pounds.
Local fairing skin changes were introduced to eliminate skincracks that occurred during final cure due to thermal condi-tions and the pressure necessary to deform the core. Thesecracks only occurred in the three-pound core region where the90* material terminated.
The following modifications (Figure 38) were made to theprototype fairing skin:
1. Extended 900 uni to trailing edge wedge.
2. Shortened inner skin by .5 inch to have core splicecoincide with 0' rib strip.
3. Trailing-edge wedge will have all 900 material added tofairing in subsequent assemblies for additionaltolerance (forward only).
3.6.9 Pendulum Vibration Absorbers
Provis:ions were made on the HLH rotor blade for the installa-tion of pendulum vibration absorbers (Ref. Drawing No. 301-59800). These pendulum absorbers are masses mounted on afiberglass collar (Ref. Drawing No. 301-55117) that is bondedto and clamped around the blade spar between the attachmentpins and the airfoil cutout. These masses are designed tominimize vertical root shear forces by flapping about a hori-zontal axis. 'iwo types of absorbers were designed, one tunedto react 3/rev root shears, and the other tuned to 4/rev.The mounts arc positioned such that either the 3/rev, the4/rev, or a combination of both can be installed at one time.The structural qualification test described in test planreport number D301-10115-23 (Reference 10) was not conductedbefore the program was terminated.
3.6.10 Blade Drawings
A complete list of the prototype blade drawings is given inFigure 39. The blade assembly drawings are included in Figures40, 41, and 42.
90
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94
Figure 36. Comparison of HLH- Spar Heels Beforeand ,flter Modifications
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4.0 STRUCTURAL AND AEROELASTIC ANALYSES
The structural analyses used in the design development and
preliminary structural substantiation of the HLH rotor bladeare contained in Reference 11. The report contains rotorloads analyses, physical properties, natural frequencies, anddetailed stress analyses.
4.1 CRITERIh AND- REQUIREMENTS
The design criteria for the rotor blade limit and fatigueloading are in accordance with the requirements of AR-56,Reference 12, for a crane helicopter except as noted in thedeviations contained in Reference 13, PIDD Revision E. Basicrequirements for the HLH hielicopter are summarized in Figure43. The design maximum level flight airspeed, V1., at thebasic design gross weight is 150 knots.
The design requirement specifies that the fatigue safe lifeshall be equal to or greater than 3600 hours based on meanminus 3 sigma (M -3 a ) allowables and top of scatter measuredloads. The loading schedule used to calculate the designfatigue safe lift is given in Table 4. The sate life isbased on the airspeed distribution for flight maneuvers givenin Table 5.
The failsafety requirement is that thu blade shall have aminimum operating life of 200 hours after a failure Lutectionwith a zonfidence level associated with moan -2 sigma (M -2 uallowables. In the case of a redundant structure, a minimumof 100 hours of safe life is requirod after comp].et•± failureof onea of the load paths using mean minus one sigma (M -l )alluwables. The failsafe life is calculated using the sumeloads and airspeed distribution as in the safe lifecalculation.
4.2 ].LIMIT AND ULTIMATE LOADS
The critical limit load conditions are:
* 2. 5g flight pullup maneuvero Rotor starting
* Rotor brakinge Ground flapping at 4.67 "g" ultimate load
119. .1,
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Figure 44 defines the spanwise distributions of rotor bladeflight maneuver limit loads.
A factor of safety of 1.5 is applied to the limit load todetermine the ultimate design load.
4.3 DESIGN FATIGUE FLIGHT LOADS
The rotor blade design fatigue loads are based on theoreticalpredictions for high-speed level flight condition and ormaneuver load factors from CII-47 helicopter measured flightdata. The L-02 computer program for aeroelastic rotor bladeloads analysis with its nonuniform downwash option was usedto predict the flapwise and chordwise bending moment at thelevel I-Light design condition of 118,000-pound gross weightand 150-kne.t. forward speed (VH) at sea level/95'F. The rootend chordwise moment was established using the lag dampercharacteristics with predicted vibratory lag angles. Thesepredictions for bending moment arc shown in Figure 45. Thecorresponding rotor blade centrifugal force distribution isshown in Figure 46. CII-47 flight lest measured pitch linkload data was combined with the L-02 analysis to establishthe spanwise distribution of rotor blade torsion shown inFigure 45. The maneuver load factors based on flight experi-once and the complete listing of mission profile loads ai'egiven in Reference 11.
4.4 MATERIAL PROPEIRTIEs
Material properties are zequired to establish thi weight andstiffness of the rotor blade,wliich in turn are requirud topredict natural frequencies and loads. Thu fatigue andultimate strengths of the material,; aro aluo requii'ed todesign a structurally adequate rotor blade. The materialproperties and strengjths for the basic rotor blade matu:ialsare sunuuarized in Table 6. Thru 1roj•ertico for the bladecompositv materials and titanium iiose cap are not containedin military specifications and were determined by coupontesting as required to support the design.
Tie proper:ties of the isoiatud matexials are not necesuarilythe usame as when they are combined to form the iotor bladestructure. Thlis; is especially true in the casu. of compositeswhore the combined strength of the elements is often dififu:entfrom the i1ndividual materials. The component testL deucribedin Section 6 ot this :report investigaite the combinod material
120
bf
4 8
strength of the total rotor blade structure that is requiredto demonstrate its load-carrying capability. These tests areused to substantiate the analytically predicted structuralcapability of the rotor blade.
The stress/load cycle (S-N) curves and the stress ratio effectson fatigue strength are defined in Paragraph 11 of rotor bladestructural substantiation report, Reference 11.
The strength of the Nomex honeycomb core was defined duringthe demonstration testing and is discussed in Section 61 ofthis report and in the Full Scale Blade Fatigue Test Report,
Reference 9.
4.5 BLADE PHYSICAL PROPERTIES
The rotor blade physicai properties were developed during thedesign phase to meet the requizements for blade weight andcentrifugal force, loads and frequencics. These propertiesshown in Table 7 and Figures 47 and 48 include the spanwise dis-tribution of:
Weight Pitch InertiaAxial Stiffness Chordwiso Neutral AxisChord Stiffness Shear CenterFlap Stiffness Static Balance AxisTorsion Stiffness
The design loads wore calculated using the properties for theATC blade configuration. The basic structural concept forthe prototype blade is identical to that for the ATC bladea'nd the minor differences in proptort-iJvi will not iignificantlychangje the design load';.
4.6 ULTIMATE ST1RNGoT1h ANAIhY,';I8
'le ultimate loads are obtained by multiplying the limit load:{by the 1.5 ultimate factor of Liafuty. The minimum rmargins ofsafety (MS) calculated for the primary structural componentsaxe shown in Table U. The margin: of !iafoty in• d1,fi.1ie.d bytie following formula: - UltiwaLu SLruiigLli
Ultimwatc Load
These margi.ns use tha blado loads and material :;trengtlhudescribed in Parag.uapl,; 4.2 andI 4.4 olu this report.
1 121
I:L..
DESIGN MAX. ALTERNATE MIN. MISSIONITEM GROSS WEIGHT DESIGN G.W. PROFILE G.W.
Gross Weight 118,000 lb 148,000 lb 73,000 lb
nimit Maneuver +2.5/-0.5 +2.0/-0.5 +2.5/0.5
Load Factor
Center of MOST FORWARD MOST AFTGravityRange 60 in. fwd 40 in. aft I
DESIGNROTOR SPEED POWER POWER
RPM ON OFF
Minimuml 15. 7 14U.1Normal 155.7 -Maximum 155.7 176.9Limit 171.3 i(14.6
Figure 43. Basic Design Ruquirement.,
122
I II II I II I1 II III II1 I I
TABLE 4. BASIC FATIGUE LOADING SCHEDULE
GROSS WEIGHTCONDITION % OCCUR.* (LBS) % TIME
Ground Conditions 1.0Take Off (40C)Steady Hovering 30.0 78,000 45Turns Hovering (2000) 118,000 50Hover Control Reversals (2000) 148,000 5Sideward Flight 2.0Rearward Flight 1.0Landing Approach (765)Forward Flight
20% VH 5.0401 V\ 2.050% V11 2.060% VH 5.070% V11 8.080% V1 9.090% VH 16.8
V11 1.0115% VH 1.0
Climb, T. 0. Powur 3.0Climb, Full Power 4.0Partial Power Descent (900)Turns 5.2
(1000)Control Reversals (815)Pull Up (270)Power to Autorotation (60)Autorotation to Power (60)Steady AuLorotation 1.1Autorotation Turns u. 4
(160)AutorotaLion Control 1w'. (40)Autorotation Landing (40)Autorotation Pull Up (40)Ground-Air-Ground (100)Power Dive 2.5
*B3racketed numbers are occurrences per 100 flight hours.
123
S. . ~~~ ~ ~ .... .......... , , " .... ...... ..J. ... .. *.. 1 . ... •
TABLE 5. MANEUVER AIRSPEED DISTP.IBUTION
FORWARD LEVEL FLIGHT MANEUVER
% VH %TIME % TIME OR OCCURRENCES
iO 5'.0 -
40 2.0
30 2.0
60 5.0 64
70 8.0
80 9.0
90 16.8 33
100 1.0 3
TOTAL 48.8
The t maneuver or occurrences from the basicfatigue schedule are di;txibuted with airspeedas given in right hand column.
'124
lI
. . . . . -. . . . . . . . . . . .-.- - - *
40(,
-- FLAP ~ 1
2_00
-200 __
400 vN
H CHORD
E-4I- I __
0j-
0 ,-ýN _-- 1
Figure-ý 44. Rotor Blade Moments for Dcsign LimitIManeuver Condition
125
FORWARD FLIGHT 150 KNOTSo _ROTOR SPEED 156 RPM
IGROSS WEIGHT 118000 LB
120 AFT ROTOR
1 -'' - ' •'•°C-O-.
80z FLAP
I-I
=4 TORSION
40 -
120 o--- CHORD
oI • -I--.. .. . .I I
0
S40 .......... ..40
S',TORSION__ j-40 .L ,
0 .2 4 .6 .8 1.0
,'i.gurre 45. Rot H: l. 1(.afe De'-.-jqgj MoflmX1-..:; for High Spc : dLLVC.1 1]'L:qht L I)
126
250
200- _ _ _ _ _ _ _ _ _ _
'LIMIT RPM
0 150lolo NORMAL RPM
50
5 156
0 .2 .4 6
Figurc 46. Rotor 1ilade Contri'ugal Force
127
-J--. . . . . . . . ....
fl. Ij' o LZ~ CN
ID e4 0ý o Ico %D (
toC En 0'0' c
9 ~1 -010
V)O 0 K,'ro 4 ID 1 0 \D0 'I
C14 c1 t 0
Lo u 24 1- '0 --
0- a 00 0 0 1 0 :in 00 C0 0)i4,j~~~ Le0D W I D
od -'4 ('4 (N N 7
0 000) 0 0H 0" 0 Co Q 0,0 0 0D 0 C)
CD20 0 - - 0 0 0 . 0 00ý4134(N (~ 0fIn -: 0 H H LP( in 'I-
niniC- -It H- "N ON -H -I ,- HH 4 -4
C-)1-4 m r,0 0 i 0 ('
co Hl H 0 co m0 H ' '
-4 IN 0
": in ul:a1-4 QD-o 0n w ' P
pin.(P( 0 0 ~ 't - (l 0 H
J. 23
TABLE 7. CPT CULATED WEIGHT AND CENTRIFUGAL FORCE
Weight CF - LB
Component LB at Bearinj
BASIC ATC BLADE 747.83 150,000(Ref. 1, D301-10227, Vol. I, Pg. 121)Tungsten Nose Weights* 8.6- 2,920ISIS Hardware 4.05 219
ArvC Blade Total 760.55 153,139Hub Hardware 370.70 10 16)
ATC TOTAL L131.25 163,299
PROTOTYPE CHANGESAft Fairing Core & Skin
ATC -121.41 -28,000Prc'totype 113.16 27,900
ISIS Mounting -3.41 -215Tip Hardware -3.84 -1,430Spar Wall (ISIS Beef Up) 3.00 1,070Precured Heel (Balanced) 21.45 5,860
Prototype Blade Total 774.50 158,324Hub Hardware 370.70 10,160Lag Damper Arm 35.33 1,610Damper Preload -- -8,000
PROTOTYPE TOTAL 1180.53 162,094
PENDULUM ABSORBERS4/Rev Assembly 36.72 2,6754/Rev Mount 10.19 6753/Rev Mount 10.40 7503/Rev Hardware 1.54 105
PENDULUM TOTAL 58.85 41205
PROTOTYPE WITH PENDULUM ABSORBERSBlade 833.35 162,529Total 1239.38 166,.99
* Added to move dynamic balance axis forward.
129
I
____ATC
W!-'-!iT - LB/IN ------ PROTOTYPE
TOTAL WEIGHT: AXIAL STIFFNESS -LB
ATC 760 LB
PROTOTYPE 774 LB 200-
3
r2100
0 0 I I
CHORD STIFFNES 1000W FLAP STIFFNESS
LBAINTCIV\ 106 LB-IN?
8000 r78oLj00
6000 600
4000 400-
2000 200[
- i iooL.
500 TORSIONAL STIFNESS PITCH INERTIA LBLSTI2NE4006106 LB-IN? 2060
300[ 600
100 L200 _ Vi
I * 1 I . ± I 0 '. I I~
0 01 1 i
ROTOR BLADE STATION - r/R (R 552 IN.)
Figure 47. Spanwise Disiribution of Mass and Stiffness
130
106 LB-NN
ATC---------- PROTOTYPE
NEUTR.AL AXIS - IN. SHEAR CENTER - IN.FPITCH AXIS PIj.CJ{ AXIS-1 DAT1_F
-3-3I. j I ] L.L...... 1... I
0 .2 .4 .6 .8 1.0 0 .2 .4 .6 .8 1.0
STATIC BALANCE AXIS - X/C
.28[
.26I--
.24
.22
20
18
L 1LIt III
0 .2 .4 .6 .8 1.0
ROTOR BLADE STATION - r/R(R = 52 IN.)
Figure 48. Spanwise Distribution of Blade Axis
131
........
TABLE 8. MINIMUM MARGINS OF SAFETYC'ýTTICAL
SPAN ULTIMATESTATION STRESS LOADING
PART NO. COMPONENT ULT. CONDITION CONDITION MS
•91-11199-1 Trailing Edge 497 Tension Flt. Loads .18Strip X-PlyF/G
301-11199-3 Trailing Edge 216 Tension Flt. Loads 1.668Wedge 0' Uni.Graphite 276 Compression R-tor .175
Buckling Starting (Lim.)
221 Tension I Rotor 2.17Braking
301-11181 Trim Tab 258 Tension Flt. Loads .12
301-11179 Core 407 Shear Airloads .38
301-11175 Skin 138 Shear/ .46
-:enOs ion
301-11189 Nickel 46! Tension Flt. Loads .18Erosion Strip
301-11174--3 Titanium 276 Tension F!. Loads .02Nose cap I
301-11173-1 Spar Assy. 220 Compression Ground 0Buck I J ng F lapping (Lirw.)
0' Uni. F/G 386 Tension Fit. Loads .18
X-Ply F/G 276 rTonsion Fit. Loads .53
104 Tension Rotor Braking 1.95
3)1 11204 insert 66 Bearing Ground .65Flapping
EA 1628 Adhesive 138 Shear Airloads & 1.47T.E. Loads
104 Shear Nose cap .52Term ination
66 Shear Ground .11(Insert) Flapping i
132
'I
i-I
4.7 SAFE LIFE FATIGUE ANALYSIS
The fatigue criterion specifies that the safe life shall be at
least 3600 hours in order to ensure maximum service reliabili-ty. Maximum flight safety is obtained by retiring the bladesat the time the safe life expires in order to virtually elimi-
nate the possibility of a catastrophic failure during the lifeof the fleet. The safe life is based on top of scatter loadsand mean minus three sigma (M -3a ) allowables. Allowablesare based on coupon test results of the individual bladematerials.
During the initial design, all blade components were sized for
unlimited fatigue life at a load equal to 1.2 times the high-speed level flight (VH) design condition load. The criticalelement for the ATC rotor blade was the fiberglass crossplyskin that h.,d unlimited life for 1.16 times the VH designload. The endurance limits for unidirectional fiberglassand titanium were 1.31 and 1.43 times the VH design load.
Safe life of the titanium nose cap was calculated using theflight spectrum loads including the combined effectL. of alter-
nating tension and shear stresses. The results of this calcu-lation led to safe life prediction of 15,500 hours. The safe
life of the fiberglass crossply was calculated at 185,500hours, indicating that this element is less critical than thetitanium even though the fiberglass crossply unlimited lifefactor is lower.
4.8 FAIL SAFE ANALYSES
Analyses were performed to evaluate the structural adequacy
of the rotor blade after the occurrence of a nartial failure.
In the structurally redundant root end attachment, the fail-safe criterionreqguires that at least 100 hours of safe lifeexist after the complete failure of one load path. The criti-cal lug that normally reacts the highest flight load wasassumed to be failed. The remaining safe life prediction withone lug failed was 175 o 1Lou based on mean minus one sigma(M -I ) allowables and top of scatteL flight loads. 'Theultimate margin of safety for the failed lug condition is.75.
.133
For the outer portion of the rotor blade, the failsafe designcriterion raquires that 200 hours of safe life exist after areadily detectal'le failure has occuirred. The complete ti-tanium nose cap and one-half of the lower zero degree uni-directional fiberglass spar was assumed to be failed, therebyrepresenting a readily detectable failure. With the uni-directional fiberglass failed, the +450 crossply fiberglassmaterial was considered capable of maintaining torsional con-tinuity of the section. This mode of failure was consideredto realistically represent a potential in-service failurethat has been demonstrated during the oval tube testing toinitiate an ISIS system warning while still. providing thebeam continuity required to carry axial and torsional loadingregardless of the spanwise extent to which the unidirectionalmaterial failure has progressed.
The remaining safe life of 1006 hours was calculated. Therequired confidence level for this mode of failure is achievedby using top of scatter loads and mean minus two sigma (M ,- 2 u)allowables. The ultimate margin of safety for the fail&-femode on the outer portion of the rotor blade is .35.
4.9 NAWRAI )?REQUENCY ANALYSIS
The rotor blade flapwise, chordwise and torsional naturalfrequencies aLe predicted using the Leone-YMyklestad method(L-01 computer program). The natural frequencies at normaloperating rotor speed are summarized in Table 9. Thespectrums plotted in Figure 49 define the variation of thenatural frequencies with rotor speed from stationary (0 rpm)to normal at 156 rpm. Comparisons with measnred frequenciesare made in Section 6.2.
TABLE 9. NONDIMENSIONAL NATURAL FREQUENCY
Mode Natural Frequency Per Rev at 15b RATC Prototype
Flapwise 1st 2.67 2.692nd 5.11 5.093rd 8.70 8.524th 13.21 12.82
Chordwise 1st 4.80 4.672nd 12.12 11.52
Torsion 1st 6.46 6.432nd 12.'9 12.77
134
,, -. . . . . .
I,4.10 CLASSICAL FLUTTER
The results of the classical flutter analysis using the L-01computer program indicate that the rotor blade is free fromfluttei up to 1.15 times the limit rotor speed (224 rpm) forforward speeds up to 1-15 VD (209 kts). This favorablecha-racteristic is attributed to the blade shear center loca-tion and the mass center both lying forward of the aerodynamiccentcr. The separation of flap and torsion natural frequen-cies also contributes to the avoidance of classical flap-pitchflutter, The stability conclusion applies to both 00 and26,50 ef h kinumatic flap pitch coupling.
4.11 ROTOR BLADE TORSIONAL DIVERCENCE
AR-56, Paragraph 3.6.2, "Aeroelasticity," states that,"...The rotor blades...shall be free of flutter, divergenceand any other aeroelastic instability at rotor speeds up to1.15 times the design limit rotor speed with and withoutpower at 1.15 VD.' HLH blade motions and loads were analyzedto assure compliance with this requirement, and the I .gh-speeddive condition was tested during the test of the dynamicallyscaled 14-foot diameter HLH rotor model. The results of theanalysis have been previously reported in Reference 11, andthe results of the wind tunnel test are reported in Reference14. Only the conclusions of the analysis and test will besuniumarizcd here and the reader is referred to the referencesfor further detail.
The analysis was performed using the Boeing Vertol C-60 BladeLoad Analysis Computer Program supplemented by a manual calcu-lation to add coupled drag and lift moments about the tor-sional axis due to blade bending. The results showed thattorsiona± loads are not excessive and do not cause unstabletorsional tip deflections. The method of analysis w:as vali-
dated by J.ts application to a CH1-47C high-speed flight testpoint; it c.ompared favorably with the measured data. Analysesof other iILH flight conditions were also performed withoutincident. Sensitivity studies of the rebults to blade stiff-niess and twist were also conducted and showed no indication ofdivergence, but rather a moderate increase in loads. Airfoilcharacteristics and their relation to these analyses were alsoevaluated.
135
Sj
The wind tunnel test results were consistent with theanalytical findings. The tEost condition was flown at the1.15 VDive speed and normal rotor rpm resulting in a 200-knotforward speed, an advance ratio A = .45, and an advancingblade tip Mach No. = .975. The test results are showni inFigures 50 and 51. Figure 50 shows measured modelblade torsion loads vs. P for the dive condition shown bythe triangle test points compared to level flight trim con-dition test results shown by the circle points. It may benoted that thb measured loads at the 1.15 VDive speed are justabout equal to the scaled endurance limit load which in gen-eral represents a good match between fatigue design and loadlevel. Figure 51 shows a comparison of the waveform of thepitch link load measured in the 14-foot model test comparedwith the waveform predicted in the Reference 11 analysis forthe same condition. Remarkable agreement is shown in thiscorrelation. Additional test results are shown inReference 14.
4.12 PTTCH LAG STABILITY
The stability boundaries are determined by the procedures ofReference 15 based on the lag damper critical damping ratioof 26 percent. The forward rotor boundary is less criticaldue to the incorporation of delta -3. The region designated"level flight" includes all gross weight/cg/airspeed condi-tions within the HLH flight envelope. The "maximum g pull-up"yields the most adverse combination of large coning and largel'g angle while the "steep turns in autorotation" combine themost adverse high coning and low pitch angle conditions.
All flight conditions invostigated are well within the cur-rent range of experience, aad provide ample clearance to thestability boundaries.
136
-. . ..l. .--- -.I I. . . -. . ..- .-...~.. : l.... 1 1 I I I II 1 1 i l l I ! I
ROTORI SPEED - RPM
25&
FLSEONA FA
113
[I
0CSCN1L
SCALED WIND TUNNEL TEST DATA6000
I
5000 D :VECT/a = .066
4000
Q0oCD 3000
P rH U JLVEL FLIGHT
Sc7,,!/_ .09
X 2
200- /
0cl1000
0
0 ~~C. .2 .3 .4 .5
ADVANCE RATIO P
Figure 50. Demonstration of Blade TorsionalAeroelastic Stability at High SpewedWith 14-Ft-Diameter Model Rotor
A38
V = 200 KNOTSGROSS WEIGHT 118,000 LBVTIP = 750 FPS
SEA LEVEL/STD
4000
S ,, .,,.-,,c--o
0""ANALYTICAL FULL SCALE
(C-60)8000[-
6--
0 100 200 300 360
BLADE AZIMUTH - DEGREE
Figure 51. Comparison of Analytical Pitch Link LoadWit h- 14-."t1-D).iameter. Rotor Wind Tunnel TesLLoad for Limit 1)3sdgn )ive Speed
3 1
,t. .•& . &, . A-.• fl•d 4' " ".~.b . - .. w ,.. -- V-- ."'" ' ... n" • ' - n. . ... " ' ' 4, " . .. ' " ... . . "4 ... .. .. . -... .
5,0 MANUFACTURING DEVELOPMENT
The HLH advanced composite rotor blade is fabricated withelectrically heated, zone-temperature controlled, match metaldies using internal pressure.
The use of computer-based Master Dimensioning Information wasextremely successful in coordinating the fabrication of thevarious tools required and ensuring that the advanced aero-dynamic rotor blade contours were attained. The compositespar is co-cured with the titanium nose cap. The matchedmetal tools assure airfoil surfaces and blade physical pro-perties that are consistent and repeatable from unit to unit.The HLH rotor blade design possesses many inherent featuresdirected toward Lhe use of automated tooling for high-rateproduction which will result in reductions in unit cost.Nondestructive test techniques have been developed and arenow available to provide the high level of quality assuranceneeded for production of composite rotor blades.
The most important aspects of the tooling, titanium forming,fiberglass fabrication inspection methods, and the results offabricatior, of the initial prototype blades are described inthis sectiorn,
5.1 SPAR FABRICATION
The fabrication involves the assembly of details on a bag andmandrel, installation of a leading edge assembly, and thewiring operation. The detailed manufacturing development ofthe I{LH rotor blade is given in D301-10280-1, Reference 16.
Figure 52 and the flow chart in Figure 53 show the majorevents in the manufacturing sequence of the HLH ATC rotorblade. The steps that go into each major event are defined
in the flow charts of Figures 54 through 62.
5.1-1 Fabrication Results
The fabrication cencept was not intended to be a "productionprocess" when it was devised, but it was intended to be astepping stone to 4he production process. Conceptually, theresults were high',- successful and satisfactory. Laminatequality and integrity in this matched die concept were uniformand excellent, and successful results were achieved in thefollowing areas:
140
a Spar laminate fiber orientation, density and uniformity
a Skin laminate bond to honeycomb
e Bond integrity of co-cured titanium cap and fiberglassspar
v Contour repeatability of matched die tooling
& Consistency of lugs
* Weight control
e Outstanding accuracy of NDT techniques
The two problem areas that emerged were:
e Wrinkling of the "D" spar heel and shank fiberglasscrossply caused by handling of the uncured layup
e Secondary bonding of cured fittings to the spar
The steps taken to eliminate these problem areas are dis-cussed in the following paragraphs.
SCrossplv Wrinkles
The crossply wrinkles in the shank (spar inboard of airfoilfairing) area were eliminated by adding unidirectional fiber-glass to fill the spaces between the spar packs and byimproved control of the lay-up procedure. The solution chosento eliminate the heel crossply wrinkles for a production bladeconfiguration is to precure the heel as a structural memberin a separate operation prior to the spar assembly. Thelayup and cure of this detail is an additional cost; but theheel permits the elimination of another cure and thus paysfor itself.
Seconeary Bonding of Cured Fittinqs
Difficulty was encountered in the bonding of the precured"fiberglass fittings at the root end and tip. The contours ofthe matching parts could not be maintained without time-con-suming and costly hand fitting. This was particularly truefor hot bonds where the thickness of the bond line could notbe controlled. An interim solution for the ATC configuration
141
Ai
was to cold bond (with EC-2216) the fittings in place. Theprototype fittings and intended production approach was tocure the fittings with the spar.
5.2 TOOLING
The blade spar and airfoil section molds are made fromMechanite H.S. The molds are integrally heited and self-contained. A photograph of the complete tjol is shown in
Figure 63. Steel was selected for the tool material for itsdurability and compatibility of thermal coefficient of expan- Ision with those for the spar materials as shown in Table 10.A photograph of the spar curing tool is shown in Figure 64.An electrically heated tool system was selected over liquid(oil) and steam-heated systems. The objection to the latterwas primarily potential contamination of the composition withoil or moisture which would result in poor bonding.
The temperature was regulated by a 60-zone computer-controlledon/off switching system. During the cure of the first toolproving spar, computer control system operation proved thatit was capable of automatically controlling zone temperatureto within required limits as shown in the heat chart inFigure 65.
The tool base has an integral air system used for cooling the
fixtures after the cure cycle.
5.3 FORMING OF TIM TITANIUM CAP
The titanium cap is formed to the outer contour of the bladeand later becomes an integral part of the spar when iL. isbonded to the fiberglass during the fiberglass cure. Theforming of the cap presented a difficult task due to thesharpness of the leading nose radius, blade twist, and airfoilthickness variation. In addition, titanium forming was notco.'mion industry practice. The changing airfoils requirestretching in some areas and shrinking in others throughoutthe 40-foot length.
The initial formi.ng conicept consisted of preforming the lead-ing edge radius on a brake using conventional punch and die.The cap was then formed using male and female ceramic diesand heated to 1A50 0F for 2 hours to produce the desiredshape. The formed parts using this method were good; however,the ceramic dies developed cracks, preventing their use forfurther production.
142
SI5
The second approach changed the tool material to Inconel 802and eliminated the female die. The cap was drape-formed overthe mandrel by attaching weights to the edges and heating to1500OF for a little less than 2 hours. This time and tem-perature kept scaling to a minimum, and the phosphate flourideetching required was kept in the region of .008 inch.Figure 66 shows a creep formed titanium cap with the weightsattached.
These experiences on the ATC blade program provided backgroundfor the improvement of certain areas in the fabrication ofsubsequent titanium nose caps for the HLH Prototype Program.Areas foi improvement included:
(1) Radius of leading edge, and
(2) chordwise bow of the outboard blade section.
Changes to the method and tools are shown in Figure 67. Forexample, 3000 pounds of additional weights have been added toimprove forming of the leading edge'radius. A cap made ofrefrasil, a refractory silicone blanket material, is beingused on the cap's leading edge during the forming operationto help control cool-down of the part. A ceramic femaleupper cap has been added to approximately 10 feet of the out-board blade section to improve nose radius forming.
5.4 QUALITY ASSURANCE
The quality level of the HLH/ATC rotor blades was achievedthrough the control of processes used in blade developmentand by thorough inspections of details, subassemblies, andthe finished product. Specimens were fabricated from thematerials used in the blade construction to check the validityof the inspection techniques. These techniques, which werelater used to inspect the rotor blade itself, gave a highdegree of confidence in the quality of materials and processes.
The critical characteristics of each blade subassembly were
inspected during fabrication, after assembly into the blade,and after blade component specimen tests. Final inspectionof the assembled blades was performed to assure compliancewith design requirements.
143
A Quality Assurance Capability Analysis was made to ensurethat component characteristics were measured and processesadequately controlled during development of the blade.
The basic element of the Capability Analysis was the QualityAssurance Flow Chart, Figure 68, which shows schematicallythe processes involved from the receipt of materials to finalassembly of the blade.
__TNondestructive Testing (NDT). Both ultrasonic and penetratingradiation (X-ray) techniques were used to determine the pres-ence of voids, delaminations, unbonded areas and fiber orien-tation. The ultrasonic inspection was perfcrmed on the "D"spar using a DoIndicator Bond Tester at the root end and heelareas, and then using the Custom VIC4Ahine o'emiautomaticscanning system to inspect the upper and lower airfoil sec-tions. The size and location of all detected indicationsequ-Al to or greater than 1/4-inch diameter weze recorded andkept on file.
144
La
:E :c
U)-
LIIL
145c
K *1
" AlIt ,IN7I:AU~i1.A I IUN
A4 6LMUI. Y
SPAII AS:.;.. Y V
__14 1 IN GINS•I'VN
lIN L -HI1 AI)IIjUN IIIN(,
HUQ1 U.N I,
I-j I• -AU II•AL A~titLMULY
Figurc 53. IIL1I Bladei 1adbricatoli Flow Dijqurm
* LEADING-EDGE ASSEMBLY
0 SPAR ASSEMBLY
* ROOT END FITTING FABRICATION
0 TIP FITTING ASSEMBLY
* FITTING INSTALLATION
0 FAIRING FABRICATION
0 BLADE BONDING 1' DAMPER ARM ASSEMBLY
* FINAL ASSEMBLY
146
I. *1
-CA
L.I /L
CLz
I rto
u in
147
VIIIIA
I All Y
[~A NUIIIAIVO kA. ;:.A,, I NUK
ISMANDIII I -4 m Al" &A,,~ A II B
rII llI~llI NIS INJA0X11 11Hill Ili Jl ANI
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TABLE 10. COEFFICIENT OF THERMAL EXPANSION COMPARISONS
Coefficient of Thermal
Material Expansion 10- 6 In.Qn./0 F
Titanium 4.7
Glass/Epoxy Unidirectional 4.8
Glass/Epoxy Crossplv 7.1
Mechanite H.S. (Tool Material) 5.9
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6.0 DEMONSTRATION TESTS
6.1 FULL-SCALE COMPONENT STRUCTURAL TESTS
This section reviews the results from a series of rotor bladestructural demonstration tests. The structural test results,along with the design load predictions, structural analysisand •.otor whirl demonstration,establish the flight worthinessof tVe HLH rotor blades. The primary objectives ol thesetests were to provide verification of the design limit andfatigue strengths of the HL:I rotor blade full-scale components.The detailed descriptions of the tests are contained in Refer-ence 17. The results of the design support root end test arecontained in Reference 9. Maximunt flight loads and special con-dition ground loads were applied to verify static strength.Fatigue loads were selected to establish endurance limits foruse in the prediction of safe life hours for the rotor bladecomponents. failure mode, failure propagation, fail-safecharacteristics, and the capability of the delta pressureintegral spar inspection system (ISIS) were also investigated.Angular deflection measurements were recorded for verificationof torsional stiffness in the root end and outboard torsionspecimens.
The first set of test specimens were made to the specifica-tions of the HLH/ATC rotor blade (Boeing Vertol Part No.301-11171-1). Results from the initial structural substan-tiation tests were used to improve the 11L11 rotor blade design.These design improvements have been incorporated into the HLHPrototype rotor blade (Boeing Vertol Part No. 301-55101-1).Results from structural demornstration tests of the ILH Proto-type rotor blade are included in this report.
The HLH rotor blade 3tructural demonstration consisted offive separate tests. The test specimens represent portionsof the rotor blade as shown in Figure 69.
163
PIKW; .
Root Ena
These tests were conducted primarily to verify the static andfatigue strength of the zoot-end section of the HLH rotorblade.
"* The root end static strength was demonstrated bysuccessfully sustaining limit and ultimate loads.
"* The 18,000-pound lug load fatigue endurance limitestablished by this test is sufficient to justify asafe life prediction of over 3600 hours (see Figure 70).
" A requirement. for a revised anti-fretting system waaidentified during the initial fatigue testing. Fiber-glide was demonstrated to be a satisfactory solution#or inhibiting fretting of the root end metal hardware.
"* The design development root end test specimen could notretain ISIS vacuum due to leakage in the vicinity of thelag damper arm. During the structural demonstrationtest, a bulkhead, installed immediately outboard of thelag damper arm, was proven sufficient to retain the ISISvacuum in the root end of the blade.
"* A secondary objective of the root end test was to verifythe torsional stiffness. This test indLcated that thetorsional stiffness of the root end is 1.34 tinesgreater than theoretically predicted based on a compari-son of predicted and measured torsional deflectionsbetween Stations 66 and 153.
"* The fail-safe testing demonstrated that the root end iscapable of sustaining at least 172 hours of high-speedlevel fliaht load with one attachLmont lua failed, andan additional 14 hours wifh a major failure simulatedin this test by a 6' x 12:' hole cut through the sectionat Station 104 (see Figure 71.).
164
............
Outboard Torsion
These tests demonstrate the pitching moment static andfatigue strength of the HLH rotor blade.
0 The torsion limit load capability was demonstratedon the outboard rotor blade specimen.
# A requirement for a precured heel to prevent prematurefatigue failures caused by wrinkles was identifiedduring the first torsion specimen fatigue test. Thesecond torsion fatigue test specimen with its precuredheel demonstrated an endurance limit of + 80,000 inchpounds. This endurance limit is sufficient to justifya 6131-hour safe fatigue life for the predicted flightloads. Except for the wrinkled spar heel, no failuresoccurred in either the titanium nose cap or the fiber-glass spar during the torsion testing.
* Torsional stiffness and shear center location of theoutboard section of the blade were verified by thistest.
• The specimen sustained 107 hours of dynamic loadingequal to or greater than VH load with a simulatedtitanium failure. The simulated titanium crack didnot propagate and the fiberglass did not fail.
165
Intermediate Bending
This test was conducted to establish the endurance limit ofthe rotor blade spar structure subjected to vibratory flap--wise bending moment and static CF. Figure 72 shows aspecimen in the test fixture.
0 The fatigue strength of the titanium nose cap demon-strated by the intermediate bending tests is below thesafe life design requirement. In the ATC specimen, thiswas due to shear cracks in the titanium created duringthe rolling process of the raw material. The materialprocessing was changed for the nose cap used for thePrototype test specimen and no failures were experiencedin the Prototype test due to shear cracks. In thePrototype specimen, fatigue cracks developed at moltentitanium deposits on the nose cap. These deposits werecreated during the post-forming cleaning process.(Figure 73 shows a typical fatigue crack.)
* The 16,320 psi mean minus three sigma endurance limitestablished by these tests for the titanium nose cap isnot sufficient to predict a 3600-hour life. The damagedcaps have sufficient fatigue strength to provide a pre-dicted life in excess of 1000 hours for the Prototypehelicopter mission. Coupon tests show that eliminationof defects in the titanium nose caps would result in apredicted safe life of 59,500 hours. (See Figure 74).
* 427 hours at level flight loads, and 109 hours atmaneuver loads were demonstrated during fail-safetesting of the intermediate bending specimen with thetitanium failed. The fiberglass maintained its struc-tural integrity throughout the fatigue and fail-safebending tests.
* Because of the demonstrated fail-safe characteristics ofthe composite rotor blade, cracking of the titanium nosecap is not considered to be a flight-safet-, issue for thePrototype flight test program. Therefore, Prototype bladesfabricated with the same type nose caps as used in thePrototype test specimen are flyable on an "on-condition"basis.
166
Simulated Chor~dwise Airloads
The simulated airloads test was conducted to demonstrate thefatigue capability of the Nomex honeycomb core and the bondedjoint between the spar heel and the aft fairing. The testresults are summarized in Figure 75.
[The fatigue strength of the Nomex core at the spar heeljoint was found to be inadequate for the ATC design con-figuration. Premature failures occurred in the Nomexcore due to core thickness an.I deflection in the sparheel. Neither effect was accounted for in the initialstrength prediction.
* The rotor bladt_ section was redesigned to reduce the spar
heel deflection and to strengthen the core. The sparheel was stiffened using unidirectional graphite with thefibers oriented in the chordwise direction. The coredensity behind the heel was increased and a horizontalsplice was introduced into the core. Fatigue testing ofthe redesigned chordwise airload specimens demonstratedan endurance limit for the Prototype rotor blade fairingthat is adequate for predicting a safe life of over 3600hours.
* No indications of failure occurred in the bond betweenthe fairing skins and the spar heel indicating that thismode of failure is less critical for the HLH design.
Tip Section
Static and fatigue tests were conducted to verify the ultimateCF tension and vibratory flapwise bending moment capability ofthe structural elements concentrated at the tip of the Ji1
rotor blade. The tip structure retains the weights requiredfor dynamic balance and rotor blade tracking.
o The vibratory loads applied to the tip specimen demon-strated a fatigue strength sufficient to establish asafe life prediction of over 3600 hours.
9 The ultimate strengths of the tip retention hardwarecomponents were demonstrated by the successful appli--cation of tension loads equal or greater than 1.52times the design ultimate loads.
167
I
TIP INSTALLLTION
SIMULATED CHORDWISEAIRLOArS
OUTBOAR TORSION
SIMULATED CHORDWISEAIPLC)ADS-,
INTERMED IATE
ROOT END
a
TEST SPECIMEN DESCRIPTION
Root End #1 ATC Specimen#2 Redesigned Lag Damper Arm and Fiberglide
Fretting Inhibitor
Outboard #1 ATC SpecimenTorsion #2 Prototype with Precured Spar Heel
Intermediate #1 ATC SpecimenBending #2 Prototype with Precured Spar Heel
Simulated-Chordwise #1-#5 ATC SpecimensAirloads #6-#ll Prototype Spar Heel and Fairing Core
Tip #1 ATC SpecimenInstallation #2 Prototype Tip Fittings Cured with Spar
I.
Figure 69. HLH Rotor Blade Structural Test Specimens
168
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6.2 NATURAL FREQUENCY AND STIFFNESS TEST
Static loads were applied to a full-scale HLH rotor blade to
determine its flapwise and chordwise stiffnesses. Thesemeasured stiffnesses generally confirm the analytical pre-dictions as shown in Figure 76.
A second objective of the full-scale !)lade test was to deter-mine .lapwise, chordwise and torsional natural frequenciesand mode shapes at zero rotor speed. Those natural frequen-cies and mode shapes were J -ntified by varying the frequencyof a driving force and observing the amplitude and phaserelationship of the blade response. The test results aresummarized in Figure 77.
The measured torsional natural frequency agreed closely withthe theoretical frequency for the test configuration and con-firms the predicted blade torsional stiffness/inertiaproperti-s.
The first and second flap bending frequencies compare accept-ably with the analytical values. The third mode frequency islower than calculated and further evaluation of this mode forin-flight rotating conditions is necessary.
Both the first and second chordwise bonding frequencies arelower than calculated. The differences are attributed to thelower than predicted chordwise stiffness in the area of thefairing cut out. Based on the static results, the rotatingfrequency at normal rotor speed is expected to drop from a,ilculated value of 4.8 per rev to 4.5 per rev which is stillconsidered acceptable.
The natural frequency and stiffness test results are reportedin Reference 18. The proof load portion of this test programwas never conducted because this blade was being held as aspare for the DSTR. It was planned that the proof load testbe conducted following the completion of the DSTR test.
L175Ii•
L:,
TABLE 11. COMPARISON OP THEORETICAL AND EXPERIMENTALLYDETEIhMINED D BiAE iENDLNG 'T-11'FNESS
-6 2STATION FLE'XIRAT. STIFFNESS - IX106 lb ill
/R INS FLALPW1S E CHORDWISE,7/N 3 S 4N 'I THRY S/N 3 S/Ml 4 THEORY
.13 72 1330 1280 1150 * *.158 87 * * 1875 1850 1750"218 120 8313 808 790 1330 1410 1250
29 160 501. * 496 5550 * 6100.40 220 303 288 29u 9250 * 9600
.60 330 i 231 244 235 7200 * 7900
.80 440 222 234 222 10350 * 9700
• No gages at LhUsC 1Ccat:iol,';.
TA13LL 12. COMIARISON 01" THEORETICAL AND MLASUJUýDNATUVAL I"RLQUEN CY
Tes. Rc.aulLzi Natura] lFreI'd-],
Supporit Node Na I uira) 'I sý OilMode SLathii SitLioiI8 Frequency Sctup Hub
IAI. i n. 11-z lhi 11ll
h.1t Flap 420 422 2. 3' 2.41 2.44
2nd FLap 488 2/'1 .488 7.5 7.98 b.07
3r.. Fl.ap :166 192, 366 1i. 16.3 16.8
1sv Torsion 54 9 10.17 10.6 11.4
2nd Torsrion 549 382 29.71 29.6 31.8
1is Chord 407 410 9.23 1.0.4 10.52nd Chord 469 169, 462 23.1 27.8 28.7
176
6.3 LIGHTNING TOLERANCE EVALUATION
Laboratory testing demonstrated the effects of simulated light-ning strikes (to 200 KA) to the HLH rotor blades. The resultsof these tests are contained in Reference 19.
Experience prior to these tests indicated a need to "ground"
the titanium nose cap to prevent arcing across the root end of
the spar to the rotor hub. Therefore, all tests were made
with the titanium cap grounded.
The titanium cap and nickel erosion strip will take strikes
in excess of 200 KA, with damage confined to pitting on the
titanium outer "skin".
High voltage tests confirmed that lightning will strike the
graphite in the blade's trailing edge, with a resulting
decrease in strength of the graphite wedge, which does not
constitute a safety-of-flight failure. Damage to the Nomex
core results from charges arcing to the titanium cap from the
trailing edge. It is concluded, therefore, to cover the
trailing-edge graphite with wire mesh to isolate the graphite
from a strike and to ground the mesh to the titanium cap.
These measures will enable the blade to withstand lightning
strikes in excess of 200 KA from any direction. During test-ing, the graphite in the spar did not attract any current.
For ATC HLH blades, the following lightning protection
measures were taken:
1. The titanium cap was electrically grounded to the
rotor hub through the lag damper bracket by a #6 wire
brazed to a 1.00" x 20.00" copper plate which was bonded
to the underside of the titanium cap. Current will arc
to the copper plate around its perimeter.
2. Aluminum strips were placed at inboard and outboard
ends of the blade to electrically ground the trailing
edge graphite to the titanium cap.
177
For Prototype 1LH blades, the following protection wasincorporated:
1. A wire mesh covered the trailing-edge graphite, top andbottom, to form a "Faraday Cage" to prevent penetrationof current.
2. Wire mesh was also used to electrically ground thetrailing-edge cover to the titanium cap and to groundthe tit~anium cap to the rotor hub.
These measures prevent lightning from penetrating the trailing-edge graphite, thus protecting the Nomex core from any arcinqdamage. Where the titaniumi cap is electrically grounded to therotor hub, there is a minimum weight penalty and no aerodynamuiccompromise, and the blades will take repeated strikes with norepai.+ to the mesh required.
' I
.17.. .
,+,il ..........I
6.4 WIND TUNNEL DEMUNSTRATION TEST
A 14-foot--diameter HLH rotor demonstration model was tested inthe Boeing V/STOL 20 ft x 20 ft wind tunnel. Testing was per-formed at full-scale tip speed of 750 ft/sec over a completerange of full-scale operating conditions which include thedesign hover condition (CT/u = .082) at a tip Mach number of.65 and forward flight trim conditions up to the maximumcruise speed of 150 KTAS ( P = .34) and the high-speed divecondition at 200 KTAS ( /.= .47). The model rotor and sup-porting rotor test stand structure installed in the windtunnel are illustrated in Figure 76.
The primary objectives of this rotor test (BVWT 115) were todemonstrate the performance capabilities of the HIIL rotorsystem, to obtain rotor blade loads, and to evaluate the con-cept of stall flutter damping. To accomplish the loads anddamping objectives, both blades and control system wqerestatically and dynamically scaled to the full-szale HLH-Irotor system including:
e Dynamically scaled blades (five natural mcdes)* Dynamically scaled control system mass and inertia"* Scaled spherical elastomeric bearing retention system"• Variable swashplate support stiffness" Variable swashplate damping"* Dynamic control load measurement capability
A detailed discussion of the test results is presented in thefollowing paragraphs. The complete test results are containedin Reference 14.
Hover Figure of Merit
Hover performance for the HLH/ATC 14-foot-diameter rotor wasmeasured out of ground effect. The resultant hover efficiency,or Figure of Merit (FM) is sunmarizvd in Figure 77, whichpresents FM as a function of rotor thrust coefficient (CT/Oat the design tip Mach number of .65. Correcting the 14-footrotor test results for Reynolds number and blade instrumenta--tion resulted in a FM of .751 at the design CT/o = .0827.Further correcting the FM for surface roughness to a ,smooth"condition could yield an FM as high as .781. It is believedthat the full-scale FM lies within this range (.751 - .781).The instrumentation and roughness corrections were determinedfrom two-dimensional tests of a section of the model bladeconducted at the University of Maryland Wind Tunnel (UOWT 667)in June 1973.
179
Cruise Efficiency
Lift to equivalent drag ratio (L/De) for thc 1.4-foot-diameter
HLII/ATC rotor system is presented in Figure 78 as a function
of advance ratio, for model and full-scale Reynolds numbers.This chart presents the test results at conditions corres-ponding to the HLH forward rotor at 118,000 pounds, mid-center of gravity, with the external load (fe = 250 ft 2 ) at sealevel, standard temperature; it compares these results to the
adjusted 6-foot rotor data for the same conditions. At thedesign cruise speed of 130 KTAS ( P = .292), the 14-footrotor test results indicate an L/De of 8.10 (corrected tofull-scale Reynolds number and blade instrumentation) comparedto a goal of 7.31 and an L/De of 8.13 obtained by scaling upthe 6-foot rotor test results. A further correction for sur-face roughness to a "smooth" condition could result in anL/De of 8.89.
Figure 78 relates the 14-foot rotor test results to the BoeingVertol forward flight power required theory (A-79 ComputerProgram, see Reference 14).
Flying Qualities Boundary
The flying qualities boundary derived from the 14--foot-diameterrotor test is presented in Figure 79. 'Ihe criteria for thisboundary is based on a specified reduction in rotor liftcurve slope with increased thrust. Comparison of the 14-footrotor results with the 6-foot rotor boundary indicates animprovement with the larger scale (Reynolds No.) of the14-foot rotor. The band of possible corrections to full-scaleReynolds Number includes the originally established goal forthe advanced HLII rotor, which had been based on an 11% improve-ment over the 23010 airfoil.
Acoustics
Rotor noise data was obtained during hover (OGE) and forwardflight (P sweeps) test conditions. Rotational noise harmonicdata was recorded to compare with the current predictionmethod used for the full-scale HLH. The modified Heron IIprediction (HLH/ATC program - Ref. 7th Quarterly Report)provided good correlation with the recorded model data as
shown in Figure 80.
180
4!
1,I
Rotor Blade Loads
The wind tunnel test data confirms the theoretical load pre-dictions and provides a basis for scaling to the full-scaleHLH rotor. The measured flapwise bending moments are inagreement with the model rotor analytical predictions asshown in Figure 81. The chordwise bending moment correla-tion between theory and test indicates a conservatism in thatthe theory envelopes the test data as shown in Figure 81.
Figures 82 and 83 display a comparison of the measuredpitch link load trends ana a waveform comparison at Y = .344and CT/a = .093. The characteristic nose-down torsionalmoment is seen on the advancing side of the rotor; however,the magnitude is lower than predicted, and there is a nose-down perterbation around 230* azimuth that the theory doesnot predict.
Stall Flutter Damping
The addition of stall flutter damping in the nonrotating,fixed system controls generally reduces fixed system loadsbut has no effect on pitch link load peak to peak or on thestall flutter spike within the range of conditions tested.The Boeing Vertol analog analysis generally confirms the testresults in that rotating system control loads for a bladehaving a torsional frequency near 3/rev are generally insensi-tive to fixed system damping, while fixed system loads arereduced.
One difference between the analog results and the test datais that the analog prediction shows a reductioi in all threeactuators, while the test data indicates a reduction in onlytwo of the three actuators. However, a comparison of themaximum fixed system control load without damping to themaximum load with damping shows approximately a 50% loadreduction with the addition of damping. See Figure 84.The 4/rev fixed system load was used as a basis for determin-ing actuator loads since the loads are dominated by 4/rev.
• 181
Lj
Blade Torsion Load Growth
The load growth characteristics measured during this 14-footmodel test are different between low P and high P . BelowP = .325,the load growth is caused by the inception of stall
flutter on the retreating blade; while above M = .325, theload break is caused by the torsion load growth c-n the ad-vancing blade.
Figure 85 presents a blade torsion load limit envelopebased on an alternating pitch link load of 4000 pounds(full scale). The test points shown for the 14-foot HLHrotor are compared with the line determined from the 6-foot-diameter rotor test of the CH-47C rotor (Reference 14).IVC is seen that the 14-foot HLH rotor exhibited results atleast as good as the 6-foot CH-47C rotor. The rotor designcondition of 150 knots at 118,000 pounds gross weight isbelow the 4000-pound limit established, indicating lower pitchlink loads than the predicted value used for component design.
Aeroelastic Stability
The model rotor was "flown" out to 200 knots in a simulateddive (M1,90 = .975) to check for signs of aeroelaetic insta-bility. Figures 50 and 51 show the resultant pitch linkload trends and a comparison of the predicted pitch linkwavaform versus the test waveform. No unusual load growthtrends were encountered, and the correlation of the waveformsis excellent.
182
Figure 76. 14-Foot-Diameter Model HL11 RotorBlade Installed in Wind Tunnel
183
AL
OGE
M = . 65
FULL SCALE REYNOLDS NUMiBER
DESIGN GOAL
j..76
TEST RESULTS
o .72 CORRECTED FOR•- \/ EXTERNALCINSTRUMENTATION
-: .68
0
0
. 6 4.a.06 .08 .10 .12
THRUST COEFFICIENT - CT/O
Figure 77. Hover Figure of Merit HLH/ATC1 4 -Foot-Diameter FoLcr WindTunnel Test
184
FULL SCALE CONDITIONSGW 118,000 LB.FWD ROTOR 2f 250 ft
VTTP = 750 FPS
SEA LEVEL/STD
TEST RESULTS CORRECTED FOREXTERNAL INSTRUMENTATION
EIN
[3 6
z
A-79 (NUD) ''IItEC-RY
H
ZD0 4
I-.-
2L
.20 .24 .28 .32 .36
ADVANCE JATO 0
Figure 78. Cruise Efficiency HLII/ATC 14-Foot-DiameterRotor Wind Tunnel Test
185
i i i i -
V TIP = 750 FPS
X/qd o= .20
BAND OF TEST RESULTSCORRECTED TO FULL SCALE
.16
.14 DESIGN GOAL
H
E-4
U-4
.0
I'I.08
.1 . 3.
ADVANCEL RATIO-
Figure 79. Flying Qualities Boundary HLH/ATC14-Foot-Diameter Rotor Wind TunnelTest
186
|I
CT/O = .083
MTIP = .65
120
MEASURED, EQUIVALENT TO 50 FT0 FROM FULL SCALE HLH
0
MOIFE HEON
0
© 80 P•REDICTED ATC
M MODIFIED HIERON II
0
0 60
m 40
z
20
0 5 10 15 20 25
SOUND HARMONIC NUMBER
Figure 80. Rotational No ise of 14-Foot-DiameterMcdel Roto-:
187
iCT/o .093
120 / .34
/-'-.T • RANGE OF
80
,\ETD
L02TANALYSIS
/ r FLAP •" FOR TEST
40CO DITION
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160
S 101,02• :1 2 0 A N A l ,Y S. Y. "
H t~t)NIOIT.I ON
r4
80
_ 0 .2 .4 .11.0 40
Fiqure 81. Comparison of Analytical Blade Bei~z1-:nq MomentsWith Scaled l4-Foxit-Diameter Rot-or W4ind Tunnel
TestData138
'< [
SCALED WIND TUNNEL TEST DATA
CT/h = .093
V TIP= 750 PPS
UDE;SI(;N LOAI
4000
o
I. I
, 300020 H
0
U o
1. o
-A4]4•i a
2000
C) R
•eDsITION
-18
0 , A
0 .1 .2 .3 .4
ADVANCE PJAT IO Ii
, Figure 82. Comparison of Design Pitch Link Load WithScaled 14-Foot-Diameter Rotor Wind TunnelTest Data
189
C.,/ - .093
.34
4000
2000
0%
-2000
ANALYTICAL MODELa ROTORS'AJVD TO FULL S IM-
-400-4000 ANALYTICAL FULL ")CAU'
(C-GO)
-6000 L
0 200 200 3uO 360
B~LADlE AZMUT - DEGREE
Figure 83. Comparison of Analytical Pitch Link LoadWith 14-Foot-Diameter Rotoxr Wind TunnelTest Load for Level Flight Design Condition
190'
190
Z //•\I
80
6J0 NO DAMPIkRS
4.) 40
Q/
0•) .0() .07 .013 .09 .10 .Ii
LIPP COFFI•""CUENT CT/
Figure 04. Fixed Systom Control Load Reductionrl With Damping, 14-1"oot-Diametei: ModelA•' Rotor" Tout
L191
r 191
.0 0 0 .. ..... ....11lJ
II -
CONDITION AT WHICH PITCHLINK LOAD EQUALS
1.4 + 4000 LB
.12-
4±.10
U'. .008 DESIGN CONDITIONFI 150 KNOTS
8 I118,000 LB OW
r-4
H LE, • 06
0 ,i .2 .3 4 .5ADVANCE iRAT IC
Figure 85. Blade Torsional Load Growth Fixed Load A
From 14-Foot-Diameter Model Potor Test
1'4
u!.
6.5 ROTOR WHIRL, DEMONSTRATION TESTS
Whirl Tower and DSTR tests of the full-scale HLII/ATC rotorsystem demonstrated the hover performance capability of therotor and verified its functional and structural adequacy.Photographs of the Whirl Tower and DSTR facilities are shownin Figures 86 and 87. The results of these tests arecontained in References 20 and 21.
The results from the whirl Cest that pertain to the rotorblade are summarized by the following statements:
1. The hover performance figure of merit objective of.751 was exceeded. Depending upon corrections forground effect, the measured figure of merit liesbetween .767 and .795. Conservatively, taking thelower 3evel of .767, the measured performance repre-sents a 3,000-pound increase in payload capabilityfor the HILII over the .75]. Igure of merit objective,(see Figure 89).
2. Stress and motion surveys indicate that the rotor per-formed as expected. Rotor blade frequencies closelymatched predicted values (see Figures 89 and 90).
3. Rotor blade tracking was accomplished utilizing theinboard and outboard tabs to control the pitch linkload range, as well as the blade track. Two methodsof measuring blade track (which could be used in flight)were evaluated and provide comparable results withinthe limits of the tracking criteria.
4. Rotor over-speed tezts up to 125; design rpm wereconducted demonstrating the rotor structural adequacy.
193
Vi"1' 'II
Af
1.94/
Figure 87, Dynamic System Test Rig
195
Mt = .65
.84
FM =.795.• ',FMIGE
CORRECTIONSCATTER BAND
IN,.76
0
.68 /
0
.6.8.10 .12
THRUST COEFFICIENT - CT/a
Ii
III
0- DSTR
AM N WHIRL TOWERFLAP MOMENT
80 .1-
/ LOAD RANGEI/ •,- -. - PREDICTION
S 40 '1
.2 .4 .6 .8 1.0
r/R
CHORD MOMENT
E 160 '
80 _ " , .LOAD RANGE
----------- . ---. " PREDICTION
0 r/.2 .4 .6 .8 1.0 r/R
Figure 89. Blade Bending Moments From Whirl Towerand DSTR Tests
197
[I
NORMALOPERATINGRPM
160 FLAP BENDING
70
1200 6
5SI
800 4 Q
400 2Q
09
N1200U
z DESIGN PREDICTION
CY 8C0 4
400 2s2
0 40 80 120 160 200ROTOR SPEED RPM
Figure 90. Blade Natural Frequencies Determined From Test
J 98
REFERENCES
1. D301-l0062-1, IJ-H/ATC Rotor Blade Confirmation TradsStudy, October 1971
2. D301-10100, HLH Quarterly Progress Reports
3. BMS 7-197, Boeing Material Specification for 6A14VTitanium Alloy Sheet
4. T301-10146-2, Test Results Report - HLH/ATC Rotor BladeRoot End Design Support Test, October 1974
5. T301-10209-1, Test Results Report - HLH/ATC Core andSeal Development Tests (PE-511), June 1973
6. T301-10261-1, Test Results Report - HLH/ATC Rotor BladeFatigue Test of Simulated Spar Section (PE-616)December 1973
7. T301-10234-1, Test Results Report, HLH/ATC Rotor BladePneumatic Failure Detection System Evaluation -
Eliptical Tubes, August 1973
8. T301-3.0246-I, HLH/ATC Titanium Alloy (6A14V) LeadingEdge Material - Effect of Crystallographic Texturingon Fatigue Strength, September 1973
9. D301-10239-2, Test Report - HLH/ATC Full-Scale RotorBlade Fatigue and Static Tests (PE-629), January 1976
10. D301-10115-23, HLH Helicopter Main Rotor System PendulumAbsorber Prototype Flight Performance Demonstration TestPlan, September 1974
11. D301-10227-1, HLH/ATC Rotor System StructuralSubstantiation Report, Volume I, Rotor Blade Assembly,July 1973
12. AR-56 - Structural Design Requirements - Helicopters
13. S301-10000 Rev. E, Prime Item Description Document -
Heavy Lift Helicopter
199
.
REFERENCES
14. T301-10229-1, Test Report - 14.-Foot Diameter Model HLHRotor Demonstration and Stall Flutter Damping KindTunnel Test, September 1973
15. "Pitch/Lag Instability of Helicopter Rotors", by Pei Chi
Chou, Rotary Wing Aircraft Dynamics Session, IAS 26thAnnual Meeting, January 1958
16. D301-10280, HLH Rotor Blade Manufacturing TechnologyDevelopment Report, May 1974
17. D301-10239-1, Test Plan - HLH/ATC Full--Scale Blades(PE-628) October 1973
18. D301-10219-2, Test Results Report HLLH-ATC Rotor BladeBending Proof Load, Stiffness and Natural FrequencyTest (PE-627) May 1.974
19. D301-10240-2, Lightning Tolerance Evaluation of liLHComposite Rotor Blades
20. D301-10201-2, Test Report - 1iLH/ATC Rotor WhirlDemonstration Test (PE-294) July 197.4
21. D301-10259-2, Test Report - H-LH/ATC Dynamic: System TestRig
22. T301-10243-1, Test Results Report - HLH/A'IT Rotor BladeWhirling Arm Rig Erosion Tests
200
IIr'l II------------------------
LIST OF SYhs.T,!)
AGB advanced geometry bladeATC advanced technology compone,.tBIM blade inspection methodc rotor blade chordcg center of gravity'CT/O thrust coeffi(-kz'tDSTR dynamic system tL.t rig
fe equivalent drag area
FM figurc of merit (hover performance factor)g load factorGW gross weightISIS integral spar inspection systemKt stress concentration factorL/De lift to equivalent drag ratio
(fwd flight performance)M.S. margin of safetyOGE out-of-ground effectR rotor blade radius measured from centerline of
zxttionR stress ratior rotor blade stationRPM rotor speedTE trailing edgeVH maximum forward level flight design speed
VD limit dive speed
x blade chordwise distance from leading edgea coefficient
advance ratiouI micro inches (10-6 inches)
P density
standard deviation
rotor speed
roduced 'om
BePt ..-jabje Cop
2 0 1 .... . - - 1. 4333-77
