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STRUCTURAL CONCEPTS AND AERODYNAMIC ANALYSIS FORLOW RADAR CROSS SECTION (LRCS) FUSELAGE CONFIGURATIONS

David W. Lowry, Melvin J. Rich, Saul Rivera, Thomas W. Sheehy00 Sikorsky Aircraft DivisionILO United Technologies Corporation

Stratford, Conn. 06603I • D D C

4. July 1978 SEP 191918

SFinal Report

LAJ

Approved for public release;C distribution unlimited.

Prepared for

APPLIED TECHNOLOGY LABORATORYU. St. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AVRADCOM)Fort Eustis, Va. 23604 ,.i

,44, , .,. 4

APPLIED TECHNOLOGY LABORATORY POSITION STATEMENT

This report provides a study of fuselage concepts for three low radar cross sectionaircraft configurations. An assessment of these concepts in terms of cost, weight, safety,advanced materials, risk, maintainability and aerodynamic performance was made, includingcomparisons with the UH.60A helicopter.

The three LRCS configurations were developed by the Applied Technology Laboratory.The study was conducted by Sikorsky Aircraft Division of United Technologies Corpora.tion under Contract DAAJ02-76-C-0062.

Mr. Bill W. Scruggs, Jr., of the Military Operations Technology Division served as projectengineer for this effort,

OISCLAIMEri5

The findinqs in this report ar- not to be monstrmed as an nlf,ciml Detiertmarnt of the Armmy position unless sodesignated by other authorited documinernts.

When Oovernmentr drawlrtis ýptieificatirlrs, or other data owr used for any purpose other than In connectionwith a definitely related GOviirnmernt procurement operation, the United Statat Govwnrnmant thereby Incurs norersponsibility not riny ohlir.tiirr whottnvypr' end the fect that thm tjovernrenht may hivin formulated, iurniahed,or in any way supplied lie taid irewmvin%, srricifw.atlon-ý, ur oithe; t'tita ,s i1t to ho reorierdd by ir"pl'catlorn Orotherwlse •1m in tny ,'namner lir:rinsinrt the hrildor or any tither person or cen. uorntion, or ronvoylnq aily rights or13arrnmi$,on, to mnrtim, ? chir,? ust' ur Se, I Anry i nvo thtwl livit f iii tthrilt rtay ,, any Wvgy bit 10Iitad theetO.

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configuration; (5) to identify high tenchnical risk areas; (6) to develop over-all design tranding data for helicopters, usinq the three fuselage configura-tions of conventional and advanced materials; U7) to conduct an analyticalinvestigation of the aerodynamic loads, vertical drag and mission performance 0

the three low radar cross section (LRCS) configurations ; and (8) to compare thefindings with those of the baseline UH-60 helicopter.

Structural concepts developed for the three LECS configurations showed thatextensive reshaping, as exemplified by Configuration 2, would increase fuselageweight from that of the baseline UH-60A fuselage by 223 pounds and cost by3.65 percent. When advanced materials were used Configuration 2 decreasedfrom the baseline fuselage weight and cost by 116 pounds and 3.98 percentrespectivdly. Total aircraft performance capability was degraded primarily bydrag effects. The aerodynamic analysis indicated that Configuration 2 wouldhave .a vertic climb rate at 15 percent of the baseline. Since drag is afunction of ar a and shape, the weight savings of advanced materials hadnegligible eff cts on performance. Weight, cost, and performance penaltieswere less in Configurations 3 and 1 "espectively.

Unclassiffied

SICURtTY CLASSIFICATION OF THIS PAOE(WI"en befI Snfowe.tJ

PREFACE

This study of structural concepts for low radar cross section (LRCS)fuselages was conducted under Contract DPAJ02-76-C-0062 with the AppliedTechnology Laboratory, U. S. Army Research and Technology Laboratories(AVRADCOM), Fort Eustis, Virginia.

The work was performed under the general direction of Mr. Bill Scruggs,Jr., of the Military Operations Technology Division. Sikorsky Aircraftprincipal participants were Melvin Rich, Project Manager; David Lowry,Structural Concepts; George Howard, Helicopter Design; Anthony Dipierro,Weights Engineering; lieville Kefford, Helicopter Design Modeling;Brian Carnell, Survivability; and Calvin Holbert, Reliability andMaintainability.

An analytical investigation of the aerodynamic loads, drag and missionperformance of low radar cross section configurations was performed byMr. Saul Rivera, Aerodynamicist, under the direction of Mr. Thomas W.Sheehy, Aerodynamicist, both of Sikorsky Aircraft.

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TABLE OF CONTENTSPAGE

PREFACE . . . . . . . . . . . . . . . . . . . . . . . 3

LIST OF TLLUTRATIONS ..A.I. ......... ................... 7

LIST OF TABLES. . ....... . . . . . . . . . . . . . . . .. . 11

INTRODUCTION . . . . . . . . . . . . . . . . . . .. .. 13

BASELINE UTTAS FUSELAGE CONFIGURATION. . . . . . . . . . . . . . . 15Description . .. .. .. .0. .. .. .. .. .. .. ..... 15

Baseline Fuselage Wcight and Cost . . . . .......... 16

LOW RADAR CROSS SECTION CONFIGURATIONS . ... . . . . . . . . . . . 21

Configuration 1 .. . . . . . . . . . . . . . . . . 21

Configuration 2. . . . . . . . . . . . . . . . . . . . . . . . 21-

Configuration 3 . . . . . . . . . . . . . . 22PRELIMINARY DESIGN CONCEPTS (CONVENTIONAL MATERIALS). . . . . . . . 29

Weight Analysis. . . .. ... . .. . ......... 29

Cost Analysis. . . . . . ......... . . . . . . . . . 35

APPLICATION OF DESIGN CONCEPTS TO LOW RADAR CROSS SECTION FUSELAGES 39

Configuration 1 ...... ........................ 39

Configurations 2 and 3 ...................... 39

DESCRIPTION OF EVALUATION FOR FAIL-SAFETY, SAFETY AND MAINTAIN- S~~~~~ABILITY . ................. .3

i:Fail-Safety . . . . . . . . . . . . . . . . . . . . . .. . . . 3

Crash Safety/Safety . . . . . . . . . . . . .

Maintainability. . .. .. .. .. .. .. .. .. ....... 43i

MODIFICATION OF CONFIGURATION 2 .:........ .. ..

SELECTED CONCEPTS .B.E............... . . . 48

TABLE OF CONTENTS (Cont.)

PAGE

ADVANCED MATERIAL APPLICATION ........... ................. *.. 49

F'ranes . . . . ............. . . . ........................ . . 49

Beams and Interoostals ...... ......... . . I . . . . . . 55Skins ................... .......................... .... 55Stringers ................. ......................... ... 55

ADVANCED MATERIAL APPLICATION COSTS . . . . . ............... .. 56

HELICOPTER DESIGN MODEL .............. .................... ... 63

Introduction . . . . . . . . . . . . . . . . . . . . . . . 63

HDM Results . . . . . . . . . . . . . . . . . . 63

TECHNICAL RISKS . . . . . . . . . . . .... .............. ..... . .. 68

CONCLUSIONS . . . . . . . . . . . . . . . . . . . . ........... .... ........ 69

REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

APPENDIX A, AERODYNAMICS ANALYSIS OF LOW RADAR CROSS SECTIONCONFIGURlATIONS . . . . . . . . . . . . . . . .S............ . . . 71

LIST OF SYMBOLS . . . . . . . . ... . .................. . . . . .. 138

6,

LIST OF ILLUSTRATIUNS

FIGUNE PAGE

1 PROGRAM STUDY FLOW CHART ................. 14

2 GENERAL ARRANGEMENT OF UH60A ............... 17

3 GENERAL ARRANGEMENT OF LRCS CONFIGURATION 1. . . . . . . .. 23

4 GENERAL ARRANGEMENT OF LRUS CONFIGURATION 2 ............ ... 25

5 GENERAL ARRANGLMENT OF LRCS CONFIGURATION 3 ............ ... 27

6 STRUCTURAL CONCEPTS A, B AND C....... 30

7 BODY GROUP WEIGHT VERSUS BODY GROUP WEIGHT PARAMETERS ... 32

8 ADVANCED MATERIALS APPLICATION FOR CONFIGURATION 2 . ..... 51

9 ADVANCED MATERIALS APPLICATION FOR CONFIGURATION 3 ........ 53

10 PAYLOAD AND RANGE FOR LOW RADAR CROSS SECTION HELICOPTERS.. 67

A-I GEOMETRICAL REPRESENTATION OF THE UH6OA FUSELAGE(BASELINE) . 82

A-2 GEOMETRICAL REPRESENTATION OF THE UH6OA WITH MAIN ROTORPYLON (BASELINE) . . . . . . . . . . . . . . . . . . . . . . 83

A-3 GEOMETRICAL REPRESENTATION OF THE LRCSI FUSELAGE . . . . . . 8(4

A-4 GEOMEWRICAL REPRESENTATION OF THE LRCSI FUSELAGE WITHBASELINE MAIN ROTOR PYLON ..................... 85

A-5 GEOMETRICAL REPRESENTATION OF THE LRCS2 FUSELAGE . . . . .. 86

A-6 GEOMETRICAL REPRESENTATION OF TIE LRCS2 FUSELAGE WITH

BASELINE MAIN ROTOR PYLON ........ .................. ... 87

A-7 GEOMETRICAL REPRESENTATION OF THE LRCS3 FUSELAGE . . . . . . 88

A-8 GEOMETRICAL REPRESENTATION OF THE LRCS3 FUSELAGE WITHBASELINE MAIN ROTOR PYLON ........ ................. .... 89

A-9 COMPARISON OF THE Ui16OA AND LRCSI CALCULATED SURFACEPRESSURES, TOP CENTEMLINE, = 00, $ 00 .... . . 90

A-10 COMPARISON OF THE UH60A AND LRCS1 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, = 00, = 00 ...... .. 91

7

LIST OF ILLUSTRATIONS (cont.)

FIGURE PAGE

A-11 COMPARISON OF THE UH6oA AND LRCSl CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, c = 0o, 4J 00 . ......... 92

A-12 COMPARISON OF THE UH60A AND LRCSi CALCULATED SURFACEPRESSURES, TOP CENTERLINE, a. 40 , 'QO . . . .......... 93

A-13 COMPARISON OF THE UH60A AND LRCS. CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, • m 40, 00 .0O, 94

A-14 COMPARISON OF THE UH60A AND LRCSI CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, a - 40, ' 00. . ............ 95

A-15 COMPARISON OF THE UH60A AND LRCS1 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, as 80, 00 = .0° ..... .. . 96

A-16 COMPARISON OF THE UH60A AND LRCSl CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, a a 80, * 00.. . ......... 97

A-17 COMPARISON OF THE UH60A AND LRCS1 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, a = 80, 4 = 00 . . 98

A-16 COMPARISON OF THE UH60A AND LRCSl CALCULATED SURFACEPRESSURES, TOP CENTERLINE, .= -4, p 0 . .... .... 9

A-19 COMPARISON OF THE UH60A AND LRCSI CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, a -40, 'p .0 .. . . i..00

A-20 COMPARISON OF THE U160A AND LRCSI CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, a -_40, 00 = o. •..... I...

A-21 COMPARISON OF THE UH60A AND LRCSJ. CALCULATED SURFACErn•N-..xs TorENT--L-N°,) p = .. . . ... 1. 2

A-22 COMPARISON OF THE Un6OA AND LRCSI CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, a -6 0 , p = 00 . . .

A-23 COMPARISON OF THE U1160A AND LRCS1 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, -8, p 0°. ......... 10

A-24 COMPARISON OF T}lE Ul16OA AND LHCS2 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, a = 00, 00 . . . ...... 05

A-25 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACE,PRESSURES, BOTTOM CENTERLIWE, Q 00, 00 .°. . .....

A-26 COMPARISON OF THE U1160A AND LIRC2 CALCULATED SURFACEPRESSURES, LATFRAL CENTERLINE, a = 0°, 00 = 0 . .... 107

8


FIGURE PAGE

AI-27 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, - 40, 0 .0.... .... ... 108

A-28 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, a = 0 .0. . . .... . .109

A-29 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, a= 40, 0 = .0 . . . . . 110

A-30 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, a = 80, ' 00 . ..... . I.. i

A-31 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, • = 80, ' 00. . ..... . 112

A-32 COMPARISON OF THE UH6oA AND LRCS2 CALCULATED SURFACE

PRESSURES, LATERAL CENTERLINE, c- 80, 4 00 .... ... .113

A-33 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, -4 -04, 4 - 00 ..... .... . 114

A-34 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, a - -40, 4' 00 . . . . . . 115

A-35 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, -40, = 00 ....... 116

A-36 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, -0 = °, 4 = 00. . . . . . . . i17

A-37 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, a = -80, 00 = 0. . .......

A-38 COMPARISON OF THE UH60A AND LRCS2 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, -- 80, 00. . .... 119

A-39 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, ai= 0 , = 0 00 ........ .... 120

A-40 COMPARISON OF THE UHbOA AND LRCSJ CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, 0 = 0°, .. . ..... . 121

A-41 COMPARISON OF THE UH60A AND LRC83 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, c 0 = 0, 0 = . . ... 122

A-42 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, = 40, = 00 .... .. ... 123

9

... ...... I ~ .~il~l~ 41 .Ahi~lftlL\AL


FIGURE PAGE

A-13 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, = 4 4, 00. . . ... . .

A-44 COMPARISON OF THE UH6oA AND LRCS3 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, ot = 4'1, . = 0°. ......... 5.

A-45 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, 80, 00 = 0.. . . .... .26

A-46 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, 8, = 00 . . . 12..

A-47 COMPARISON OF THE U160A AND LRCS3 CALCULATED SURFACEPRESSURES,. LATERAL CENTERLINE, c = 80, 4 = O0.. ...... 128

A-48 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, = -4 0 , 0O . . . . 129

A-49 COMPARISOIN OF THE UIH60A AND LRCS3 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, = -4 0 , 4 = 00. . . . 130

A-50 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, 0= 04, = .o. ... .31

A-51 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, TOP CENTERLINE, a=-80, 00 . . . . 132

A-52 COMPARISON OF THE U1t60A AND LRCS3 CALCULATED SURFACEPRESSURES, BOTTOM CENTERLINE, -8u, 0). . . . . 133

A-53 COMPARISON OF THE UH60A AND LRCS3 CALCULATED SURFACEPRESSURES, LATERAL CENTERLINE, 008o, 1 3 )1

A-54 VARIATION OF LIFT WITH ANGLE OF ATTACK FOR THE UH6OA,LRCS1, 1RCS2, LRCS3 CONFIGURATIONS . ....... ... 13

A-55 VARIATION OF DRAG WITH ANGLE OF ATTACK FOR nIE UH6OA,LRCSI, LRCS2, IRCS3 CONFIGURATIONS ............ ....... . 136

A-56 VARIATION OF PITCHING MOMENT WITH ANGLE OF ATTACK FOR T.LEUIn6OA, LRCS1, LRCS2 AND LRCS3 CONFIGURATIONS . . . . . . .137

10

LIST OF TABLFS

TABLE PAGE

1 BASELINE UH60A FUSELAGE WEIGHT AND COST DATA . . . . . . . 20

2 BASELINE FUSELAGE SKIN/STRINGER WEIGHT FROM WEIGHT AND 34BALANCE REPORT . ................. .....................

3 CONFIGURATION 2A AND 2B SKIN/STRINGER WEIGHT ..... .... 34

4 CONFIGURATION 3A AND 3B SKIN/STRINGER WEIGHT .... 35

5 CONFIGURATION 2C HONEYCOMB PANEL SKIN WEIGHT .... 37

6 CONFIGURATION 3C HONEYCOMB PANEL SKIN WEIGHT .... 37

7 CONFIGURATION I FUSELAGE WEIGHT AND COST DATA ......... 39

8 CONFIGURATION 2A FUSELAGE WEIGHT AND COST DATA .4

9 CONFIGURATION 2B FUSELAGE WEIGHT AND COST DATA . .

10 CONFIGURATION 2C FUSELAGE WEIGHT AND COST DATA .... . 41

11 CONFIGURATION 3A FUSELAGE WEIGHT AND COST DATA .... . 41

12 CONFIGURATION 3B FUSELAGE WEIGHT AND COST DATA .42

13 CONFIGURATION 3C FUSELAGE WEIGHT AND COST DATA . 42

14 SAFETY AND CRASH-SAFETY RANKING. . . . . . . . . 45

15 SUMMARY OF FUSELAGE WEIGHT AND COST DATA OF CONVENTIONALCONSTRUCTION ........... ...................... 46

16 FLOOR WEIGHT COMPARISON FOR CONFIGURATION 2 . . . . . 47

17 COCKPIT STRUCTURAL WEIGHT ...... ................. 57

18 COCKPIT STRUCTURAL WEIGHT OF CONFIGURATION IA OF CONVEN-TIONAL MATERIALS AND ADVANCED MATERIALS ... .......... 58

19 COCKPIT STRUCTURAL WEIGHT OF CONFIGURATION 2A OF CONVEN- 59TIONAL MATERIALS AND ADVANCED MATERIALS .... .........

20 COCKPIT STRUCTURAL WEIGHT OF CONFIGURATION 3A OF CONVEN- 6oTIONAL MATERIALS AND ADVANCED MATERIALS . . . . . ...

21 CONFIGURATION I ADVANCED MATERIALS FUSELAGE WEIGHT AND 61COST DATA ...... .......................

11

LIST OF TABLES (cont.)

TABLE PAGE

22 CONFIGURATION 2 ADVANCED MATERIALS FUSELAGE WEIGHT AND 61COST DATA . . . .. .. . .. .. . .. . . ... . .

23 CONFIGURATION 3 ADVANCED MATERIALS FUSELAGE WEIGHT AND 62COST DATA ......... . . .............

24 SUMMARY OF WEIGHT AND COST DATA FOR CONCEPT A AND 62ADVANCED MATERIALS.......... . . . . . . . . . ...

25 PRELIMINARY DRAG ESTIMATE ................ 64

26 HELICOPTER DESIGN MODEL CONSTANT GROSS WEIGHT RESULTS. . 65

27 HELICOPTER DESIGN MODEL TRENDING RESULTS .. . .. . 66

A-I UH-60A VERTICAL DRAG CALCULK2ION........ . . . . . . 76

A-2 LRCS I VERTICAL DRAG CALCULATION. . . . . . . . . . . . 77

A-3 LRCS 2 V ETICAL DRAG CALCULATION . . . . . , . . . . . . 78

A-4 LRCS 3 VERTICAL DRAG CALCULATION. . . . . ... .... 79

A-5 SUMMARY OF PARASITE AND VERTICAL DRAG . . .. . . . .. 80

A-6 PRIMARY MISSION PERFORMANCE SUMMARY . . . . . . . . . . 81

12

INTRODUCTION

The objectives of this study were: (a) to develop structural conceptsand determine their effects on weight, cost, fail-safety and maintain-ability for th.ree low radar cross section fuselages constructed of con-ventional materials, and to compare these to a baseline Sikorsky UtilityTactical Transport Aircraft System (UH60A) helicopter; (b) based uponthe results of the study to select the structural concept for the threeconfigurations that is the lowest in weight and cost, and is the easiestto maintain; (c) to assess the application of advanced materials for eachconfiguration; (d) to develop design trending data for helicopters, usingthe three configurations of conventional and advanced materials' and(e) to identify high technical risk areas.

Three fuselage configurations for low radar cross sections were developedby the Applied Technology Laboratory. Mold line drawings of eachconfiguration were provided by the Army for this study. The main rotorpylon fairings and tail surfaces aft of a tail fold hinge for eachconfiguration were the same as those for the baseline UH6OA.

In the initial portion of this study, the weight and costs (percent oftotal) were developed for sections of the baseline UH6OA fuselage.General arrangement drawings of each configuration were developed fromthe mold line drawings. Structural concepts were developed which couldbe applied to each configuration using conventional materials. Anassessment of safety, fail-safety, and maintainability for each configu-ration was performed. The change in structural weight and the percentagechange in cost for each configuration using the concepts developed werecompared to those of the baseline. One concept was selected and appliedto the three configurations.

Having selected the structural concept with the lowest weight change andpercentage cost change for the three fuselage configurations, the effecton weight and costs using advanced materials was developed and applied tothe three configurations.

To evaluate the impact of the results of the fuselage study, designattributes of six helicopters were developed using a Helicopter DesignModel (HDM) computer program.

The six helicopters (three of conventional materials and three of advancedmaterials) were compared to the baseline UH60A. Technical risks wereassessed.

The approach to this study is illustrated on the flow chart of Figure 1.

A detailed analytical aerodynamic investigation of the fuselageconfigurations was undertaken at the conclusion of the above study. Theresults of this additional work are reported in Appendix A.

13

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191

BASELINE Un60A 1USELA3E CONFIGURATION

DESCRIPTION

The baseline fuselage of this study, shown in Figure 2, is that of theSikorsky UH-60A. The fuselage is made up of six major sections:

1. Cockpit (or nose)2. Mid-cabin3. Aft-cabin (or transition)4. Tail cone5. Pylon6. Stabilizer

The cockpit section is a molded fiberglass/epoxy framework above the floor.The framework supports the windshields, doors, and overhead controls.Structure below the floor is of built-up aluminum beams and frames. Thestructure supports the pilot and copilot seats, flight controls andelectronic equipment.

The cockpit section is of double contour with many cutouts and supportsfor a variety of equipment.

The mid-cabin section is of forged and built-up aluminum frames and beams.The mid-cabin consists of an upper section, two side sections, a tubsection and a combination walking cargo floor. The entire sectionsupports the engines, transmission, flight controls, troop seats, litters,cargo and the main landing gears.

The aft-cabin section, double contoured, is of' honeycomb sandwich bulk-heads, aluminum inturcostals, stringers and skins. This section supportstwo fuel cells and a small cargo deck.

The tail cone is constructed of bent-up aluminum frames which supportrolled formed stringers and aluminvm sheet skins. The tail cone is asingle-wrapped contour where the skin/stringer combination is riveted tofloating frames. Floating frames do not have a direct shear tie to theskins. The tail cone supports the tail rotor drive Lshaft, tail landinggear and the tail pylon fold hinges.

The pylon and the stabilizers are airfoil shaped structures of built-up

spars, formed ribs and formed stiffeners covered with aluminum sheetskins. The pylon supports the tail rotor and gearbox. The stabilizersprovide pitch control for the helicopter.

15

BASELINE FUSELAGE WEIGHT AND COST

The structural weights of the six sections of the baseline fuselage wereobtained from the actual weight and balance report for the UH-60A(Reference 1).

The costs of the six sections of the baseline fuselage were obtained froma cost/weight relationship study conducted by Sikorsky Aircraft after thedetail design of the UH60A aircraft had been completed.

The costs were determined for materials and labor for each section of thebaseline fuselage, based on the production of 276 fuselages. The costsreflect 1977 pricing.

Material costs included purchased items such as formed stringers, rolledsheet stock, extruded shapes, forgings and windshields. The totalmaterial costs required to produce each section of the baseline fuselagewere given in terms of dollars per pound of structural weight of thesection. The material costs ranged from a low of approximately $2.00 perpound to a high of approximately $10.00 per pound.

Labor hours required to fabricate each section were determined as follows:

1. Fabrication of individual structural elements that make upframes, beams, bulkheads, fittings, stringers, intercostals,longerons, and floors.

2. Assembly of the structural elements to produce frames, beams,etc.

3. Installation of frames, etc., to produce a structural section ofthe fuselage.

Labor costs were based on $22.50 per hour.

No tooling, engineering, or development costs were considered. It shouldbe noted that tooling costs are relatively small for production heli-copters (less than 3 percent- Reference 2).

i. UH-60A VOLUME 3, PART 1-III WEIGHT AND BALANCE dated September 27,1976.

2. Rich, M. J., INVESTIGATION OF ADVANCED HELICOPTER STRUCTURAL DESIGNS,VOLUME 1 - ADVANCED STRUCTURAL COMPONENT DESIGN CONCEPTS STUDY,Sikorsky Aircraft Div., USAAMRDL-TR-75-59A, U.S. Army Air MobilityResearch and Development Laboratory, Fort Eustis, Va. 23604, May19Th, AD AO262h6.

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Table 1 shows the total structural weight of each section of the baselinefuselage. Material and labor costs, also shown in Table 1, are given interms of percentage of the total fuselage cost.

The weight and cost data developed in this study pertain to the fuselages*1 considered. Weight and cost data for complete helicopbers resulting from

the fuselages are developed by a Helicopter Design Model (HDM) computerprogram discussed in the Helicopter Design Model section of this report.

19

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02

LOW RADAR CROSS SECTION CONFIGURATIONS

Throe low radar cross section fuselage configurations for this 9tudy weredeveloped by the Applied Technology Laboratory. The first configurationslightly modified the nose section from the baseline configuration; thesecond configuration changed the fuselage shape along the lines of atruncated triangular prism; the third extended canted flat side shapingthroughout the fuselage. The tail surfaces and main rotor pylon fairingwere the same as those of the baseline UH60A.

Mold line drawings of the three configurations were used to developgeneral arrangement drawings for each configuration.

CONFIGURATION 1

The general arrangement of Configuration 1, shown in Figure 3, wasdeveloped from mold line drawing No. 20074180. This configuration alters

the baseline fuselage forward of the mid-cabin section (the cockpit).Although this configuration is different from the baseline, the internalstructure must be compatible with the forward cabin to avoid a heavyjoining structure. From Figure 2, Lhe overall lUL&gth is olightlyincreased due to this configuration.

CONFIGURATION 2

The general arrangement of Configuration 2, shown in Figure 4, wasdeveloped from mold line drawing No. 20074135. This configuration isbasically a trapezoidal cross sectior, airframe having sides canted inward3Q0 and made up of flat exterior structural panels. This configurationis wider at the bottom of the fuselage and narrower at the top of thefuselage than the baseline. This configuration is slightly longer thanthe baseline UH6OA, and its overall height is slightly larger than thebaseline. The increased length, width, and height of Configuration 2does not allow an aircraft of this size to meet the air transportabilityrequirements of the baseline. The narrow upper fuselage causes the pilotand copilot seats to be spaced closer to each other, and shoulder roomin the main cabin is decreased. The main cabin floor is approximately6 inches higher than the baseline from the ground. The increased floor-to-ground height causes difficulties for combat troops to enter or leavethe aircraft quickly.

Minor modifications of the mold lines for the transition and tail-conesections were made to properly house the tail rotor shaft of the baselineUH6OA.

21

CONFIGURATION 3

The general arrangement of Configuration 3, shown in Figure 5, wasdeveloped from mold line drawing No. 20074136. This configuration isbasically a flat side cross section airframe having sides canted inward50 and is tapered in width from a narrow cockpit section to a transitionsection as wide as the baseline UH6OA, The tail-cone is a rectangularsection which is narrower than the baseline. The narrow cockpit causesthe pilot and copilot seats to be spaced closer to each other; space forfour-across seating in the main cabin is decreased. The cockpit and maincabin floors are at the same height from the ground as the baseline. Theslope of the windshields may cause problems of visibility for the flightcrew.

Minor modifications of the mold lines for the transition and tail-conesections were made to properly house the tail rotor shaft of the baselineUH6OA.

22

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PRELIMINARY DESIGN CONCEPTS(CONVENTIONAL MATERIALS)

The structural concepts developed are based on current UH60A technologyfor fuselage construction. The extent of the structural concepts isdependent on the complexity of the RCS shapes. Since the primary purposeof this study was to evaluate the effect on weight and cost resultingfrom external configurations, it was assumed that the internal structure(beams, frames, bulkheads and intercostals) is of current UH60A design.It was also assumed that no changes occur for the tail pylon andstabilizer.

Four concepts were developed for the three configurations. The firstconcept was applied to the cockpit canopy framework supporting the wind-shields of the three configurations. This concept is of molded fiber-glass posts and skins as show., in Figure 3 for Configuration 1. Thisconcept is currently used on the UH6OA. The other three concepts areapplied to the external structure of the lower cockpits of Configurations1, 2 and 3 and the external structure for the mid-cabin, aft-cabin, andtail cone of Configurations 2 and 3.

Concepts for the external structure are of three designs. The firstdesign is riveted skin/stringer panels on floating frames (Concept A); thesecond design is riveted skin/stringer panels where the stringer passesthrough the frames and the skins are riveted to the frames (Concept B).The third design is an external structure of aluminum-faced honeycombsandwich panels (Concept C) riveted to frames and longerons.

Concepts A, B and C are shown in Figure 6.

The structural weights for the skin/stringers of Concepts A and B areidentical and were predicted using a fuselage shell program.

WEIGHT ANALYSIS

A Fuselage Shell Program (Y076B) was used to predict the weight of skins,stringers, longerons and stabilizing frames of the UH60A and variants ofthe UH60A for low radar reflectivity. The program uses the followinginputs:

1. Shell geometry and frame spacing

2. Loads (shears, moments and torsions)

3. Initial guesses for skin and stringer gages and stringerspacing

With these inputs, the program iterates for the optimum skin and stringerages, which are assumed to occur when a fully stressed design is reachedapplied stresses are equal to the allowable stresses). The output consists

of a bay-by-bay breakdown of the skin and stringer gages, as well asweights, for the critical flight and ground load conditions.

29

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-...---SHAFT COVER (TYPICAL)

-SKIN

[- FLOATING FRAME

"z" STRINGERS

EXTRUDED LONGERONS(TYPICAL)

CONCEPT A

NOTCHED FRAMES

-CONTINUOUS STRINGERS, CLIPS REQ D

SKIN

HONEYCOMB PANELS

CONCEPT B CONCEPT C

PTnURE 6. STRUCTURAL CONCEPTS A, B AND C

30

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

Iti its present form, the program can analyze rectangular shell crosssections only. Therefore, when used to predict sizes for the baselineU116OA, a rectangular :shell cross section having approximately the sameperimeter aa the actual cross section is used. Using standard Sikorskystringers (used on Ul16OA), the predicted weights for skins, stringers andlongerons are then correlated to the latest UH60A weights to obtain non-optimum factors. These factors are then applied to the RCS configurationsso that a proper weight comparison for all the fuselage cross sections canbe made.

The design of main frames to support local loads requires separateanalysis and is not considered in this study. Instead, weight deltasover the UN60A baseline are estimated using frame weights to maintainshell stability obtained from the Fuselage Shell Analysis Program (Y076B)and based on correlated test information on cylindrical shells subjectedto torsion and bending. Weight increments for each RCS configuration areobtained by taking the weight difference of stabilizing frames andupplying a 10% installation factor. These deltas are +31.8 for RCS Con-figuration 2 and +0.4 for RCS Configuration 3.

Weight deltas for each cockpit configuration are based on configuration

and wetted area changes over the baseline UH60A cockpit which weighs330.3 lb for 150 ft 2 of wetted area, or 2.2 psf. Wetted areas forRCS Configurations 1, 2 and 3 are 146, 158 and 157 ft 2 , respectively,yielding weight deltas of -8.8, +17.6 and +15.4 lb, respectively. Toverify these statistical deltas and to include effects of configurationchanges, a further weight check was made by comparing cockpit componentweights for primary structure (skins, stringers, frames), windshield,doors, cockpit enclosure, floor and supports, windows, etc.

Floor panels are treated separately from floor support beams. Weight

deltas for floor panels are based on floor area changes for each RCS con-figuration using 1.014 psf for the personnel portion (Sta. 247-288) and1.63 psf for the cargo portion (Sta. 288-398), as in the baseline UH60A.Weight deltas for floor support beams are estimated from analytical weightequations derivod from lstress relationships for beam webs and caps.

To verify the total body group weight deltas for each RCS configuration,a statistical weight derivation was made using the statistical body groupequation based on Sikorsky models. The statistical data is shown inFigure 7.

The weight of honeycomb skin panels, Concept C, was determined as follows:

1. The weights of the skins and stringers for the baseline UH60A wereobtained from Reference 1 for each section of the fuselage, and apercentage of skin/stringer weight to section weight was derived asshown in Table 2. The percentage of skin/stringer weight for the base-line cockpit was applied to the cockpits of Configurations 2 and 3.The skin/stringer weights for the other sections of Configurations 2and 3 were obtained from the fuselage shell program and are shown inTables 3 and 4.

31

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32

2. From the preliminary weight analysis of Concepts A and B, a minimumskin gage of .025 in. and a minimum stringer gage of .032 in. wereobtained. For an aluminum skin/stringer panel 12 in. by 20 in.(aluminum density of 0.1 lb/in. ),the weight is:

Skin 12 in.x 20 in, x .025 in.x . Ilb/in. 3 .58 lb.

Stringer (2 standard Sikorsky ilumintunstringers with a cross-sectionalarea of .059 in. 2 each)

2 in, x 20 in, x .059 in.x .1 lb/in, 3 .24 lb.

Total .82 lb.

3. A review was conducted to determine the minimum sheet facing thick-ness for the external honeycomb structure. The minimum gages used bySikorsky Aircraft on military helicopters are .016 in. for the outerface and .008 in, for the inner face. The honeycomb core was con-sidered to be 1/4 in. t, hick at 4 lb/ft 3 (density), the minimumconsidered practicable from manufacturing considerations. For a12-in, by 20-in, minimum gage honeycomb skin panel, the weight is:

Facings 12 in.x 20 in,(.016 in,+ .008 in. )x .1 lb/in.3 .576 lb.

Core .25 x 12 x 20 x 4 -. 134 lb.1728

Adhesive (.06 lb/ft /surface)

.06 x 2 x 12 x 20 - .200 lb.144

Close out (2 edges - 12 in.)

1.25 in.x 12 in.x 2 in.x .008 in.x .1 lb/in.3 .020 lb.

Total .930 lb.

33

TABLE 2. BASELINE FUSELAGE SKIN/STRINGER WEIGHTFROM WEIGHT AND BALANCE REPORT

SECTION TOTAL WETTED SKIN/STRINGER SKIN/STRINGERWE I GHT AREA WEIGHT WEIGHT

(LB3) (K 2 ), (LB) (PERCENT)

Cockpit 330,3 150 39.6 11.9

Mid-Cabin 995.7 350 137.8 13.8

Aft-Cabin 295.3 1.80 136.2 )46. 1

Tail Cone 168.0 1.39 L23.7 73.6

*Tail Surfaces and Miscellaneous 85 Ft

TABLE 3. CONFIGURATION 2A AND 2B SKIN/STRINGER WEIGHT

SECTION TOTAL WETTED ,3KIV/,,I'RINOER SKIN/STRINGERWEIGHT AREA WEIGHT WEIGHT

(LB) ( ,, 2)* (LB) (PERCENT)

Cockpit 3h7.9 158 41.,( 11.9

Mid-Cabin 1135.3 385 156.5 13.8

Aft-Cabin 3514.8 234 198.5 55.9

Tail Cone 173.9 161 143.3 82.4

AI'ail Surfaces and Miscellaneous 118 Ft 2

34

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

TABLE 4. CONFIGURATION 3A AND 3B SKIN/STRINGER WEIGHT

SECTION TOTAL WETTED SKIN/STRINGER SKIN/STRINGERWEIGHT AREA WEIGHT WEIGHT

(LB) ('I2)* (LB) (PERCENT)

Cockpit 345.7 151 41.4 11.9

Mid-Cabin 1017.0 345 1341.8 13.3

Aft-Cabin 324.1 213 173.1 53.4

Tail Cone 168,3 142 126.14 74.6

*Tail Surfaces and Miscellaneous 91 Ft 2

Therefore, the minimum weight honeycomb skin panel is 13 percent heavier

than the minimum skin/stringer panels of Concepts A and B. The honey-comb skin weights are shown in Tables 5 and 6.

The 13 percent increase was applied to each skin/stringer weight of each

configuration (Configurations 2 and 3) to obtain a section weight usingConcept C. The section baseline weight was subtracted and a A weightwas obtained for Concept C.

COST ANALYSIS

The estimated change in cost for each configuration and concept was basedupon the change in costs of material and labor. The material costs in-clude the cost of sheet stock, raw forgings and extruded or rolledshapes. This cost is then given in terms of cost per pound of structuralweight. The labor costs are total labor hours required to fabricate thedetails from materials, to assemble the details, and then to install theassembly. For this study, labor costs are based on the wetted area ofthe structure being fabricated.

The changes in cost for each section of Configurations 1, 2 and 3 were

estimated as follows:

Change in section cost:

Material cost + labor cost - baseline cost

An example is given as follows:

Configuration 2A mid-cabin (Table 8)

35

Material cost change:

o068 (Table 1) 1135.3Table .0775 (Mat'l Cost, %)995.7 (Table )

A cost = ,0775 - .068 - .0095 = .95%

Labor cost change:

385 (TaIe 3) X (..15 + .86 + .191)(Table 1 ) - .4312 (Labor Cost, %)350 (Table.2)

L cost a .4312 - .392 w .0392 a 3.92%

Section cost change:

.0775 + .4312 - .5087 (Section Cost, %)

A cost [ .5087 - .46 (Table 2) : .0487a 4.87%

Percent cost A from baseline is:

.0487 (.46) - .0224 , 2.24%

Concept B is the same baslc design and weight as Concept A, but it isfabricated by a slightly different method. For Concept B, stringer clipsare fabricated, and frame cutouts with joggles are formed during detailfabrication. During assembly, the okin/stringer combination is fittedto the frames,

The material costs, of Concept B are the same as thow;e of Concept A. Thelabor costs are approximateLy 10 percent greater than Concept A for detailfabrication and assembly. Installation costs were assumed to be equal.

The coot changes for the honeycomb sandwiuh skins (structural Concept C)were developed for Configurations 2 and 3.

The cost of honeycomb sandwich s]kins is based on the following:

"1. Detail fabrication, assemble and bond, 1.1i hr/ft,2

2. Materials, $6.50/ft2

For conventional skin/stringers (Concept A)

1. Detail fabrication, 1 hr/ft a

2, Assemble and automntic rivet, .19 hr/ftY3. Material, $2.14/ft•

36

• ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~~~ ~~~~~~~.. .' .. ........................................ iiiii

TABLE 5. CONFIGURATION 2C HONEYCOMB PANEL SKIN WEIGHT

SECTION TOTAL HONEYCOMB PANEL HONEYCOMBWEIGHT SKIN WEIGHT SKIN WEIGHT(LB) (LB) (PERCENT)

Cockpit 353.3 47.1 13.3

Mid-Cabin 1155.6 176.8 15.2

Aft-Cabin 380.6 224.3 58.9

Tail Cone 192.5 161.9 84.1

TABLE 6. CONFIGURATION 3C HONEYCOMB PANEL SKIN WEIGHT

SECTION TOTAL HONEYCOMB PANEL HONEYCOMBWEIGHT SKIN WEIGHT SKIN WEIGHT(L~B) (LB) (PERCENT)

Cockpit 350.9 46.7 13.3

Mid-Cabin 1034.5 152.3 14.7

Aft-Cabin 346.6 195.6 56.4

Tail Cone 185.7 142.8 76.8

37

Based upon an assumed labor cost of $22.50 per hour (Reference 2), the

cost per square foot of honeycomb is

1.11 x 22.50 + 6.50 = $31.48

For skin/stringer panels

(1 + .19) 22.50 + 2.14 - $28.92

The honeycomb panel is 8.8 percent more expensive than the skin/stringerpanel to fabricate and assemble. The installation costs are assumed tobe equal. The material cost for the honeycomb panels is

$6.50/.93 - $6.85/lb

For skin/stringer panels, the material cost is

$2 2 $2.6/lb

Material cost for honeycomb panels is 164 percent more than for skin/stringer panels, The increased material costs are due primarily tothe honeycomb core, which can be between 3 and 10 dollars per squarefoot,

Labor costs for honeycomb panels are T percent less than skin/stringerpanels for the cockpit, aft-cabin and tail-cone sections. The laborcosts for honeycomb panels in the mid-cabin section are 5 percentgreater than conventional construction due to structural inserts andlocal reinforcing required to mount the many components attached to themid-cabin.

38

APPLICATION OF DESIGN CONCEPTS TO LOW RADAR CROSS SECTION FUSELAGES

CONFIGURATION I

The structure of Configuration I could not be changed since the structuremust match the mid-cabin of the baseline and not alter the baseline mid-cabin structure. The change in weight and the percentage change in costfor Configuration 1, from the baseline, are shown in Table 7.

CONFIGURATIONS 2 AND 3

The three structural concepts developed are applicable, as shown inFigures 3, 4 and 5, to the lower cockpit, mid-cabin, aft-cabin and tailcone of Configurations 2 and 3. Tables 8 through 13 show the changes inweight and the percentage change in costs for Configurations 2 and 3using Concepts A, B and C.

The percentage change of the cost of each section of Configurations 2 and3 is based on the difference of the sura of material and labor costs fromthe baseline section cost divided by the baseline section cost.

The percentage change for each configuration fuselage cost is based onthe sum of the cost change for each section,

TABLE 7. CONFIGURATION 1. FUSELAGE WEIGHT AND COST DATA

FUSELAGE SECTION SECTION COST A FROM BASELINE COST A TOTALSECTION A WEIGHT MATERIAL TOTAL LABOR SECTION COST BASELINE

(LB) (PERCENT) (PERCENT) (PERCENT) FUSELAGE

Cockp'it - 8 - .093 - .353 - .318 - o00063

Mid-Cabin 0 0 0 0 0

Aft-Cabin 0 0 0 0 0

Tail Cone 0 0 0 0 0

Pylon 0 0 0 0 0

Stabilizer 0 0 0 0 0

Total - 8 - .00063

39

TABLE 8. CONFIGURATION 2A FUSELAGE WEIGHT AND COST DATA


(LB) (PERCENT) (PERCENT) (PERCENT) FUSELAGE(PERCENT)

Cockpit 18 .15 .54 .35 .05

Mid-Cabin 140 .95 3.92 4.87 2.24

Aft-Cabin 59 1.15 5.00 6.00 1.32

Tail Cone 6 .02 .62 .61a .04

Pylon 0 0 0 0 0


Total 223 3.65 i

TABLE 9. CONFIGURATION 2B FUSELAGE WEIGHT AND COST DATA

FUSELAGE SECTION SECTION COST A FOM BASELINE COST 4 TOTALSECTION A WEIGHT MATERIAL TOTAL LABOR SECTION COST BASELINE


Cockpit 18 .15 .56 .36 .05

Mid-Cabin 140 .95 1.13 5,08 2.31

Aft-Cabin 59 1.15 5.32 6.)12 1.42

Tail Cone 6 .02 .65 .67 .05

Pylon 0 0 0 0 0


Total 223 3.83

40

[J

TABLE 10. CONFIGURATION 2C FUSELAGE WEIG11T AND COST DATA

FUSELAGE SECTION SECTION COST A FROM BASELINE COST A TOTALSECTION A WEIGHT MATERIAL TOTAL LABOR SECTTON COST BASELINE


Cockpit 23 .79 .49 .31 .04

Mid-Cabin 160 2.52 10.00 9.31 4.51]

Aft-Cabin 85 7.41 6.85 6.78 1.24

Tail Cone P5 .85 i.08 1.07 .07

Pylon 0 0 0 0 0


Total 293 5.86

TABLE 11. CONFIGURATION 3A FUSELAGE WEIGHT AND COST DATA



Cockpit 15 .15 .0i .V, .02

Mid-Cabin 21 .14 - .52 - .39 - .18

Aft-Cabin 29 .55 2.99 3.5) .78

Tail Cone .3 .08 .09 .10 0

Pylon 0 0 0 0 0


Total 65.3 .62

141

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

TABLE 12. CONFIGURATION 3B FUSELAGE WEIGHT AND COST DATA



Cockpit 15 .15 .05 .27 .o4

Mid-Cabin 21 .14 - .48 - .35 - .17

Aft-Cabin 29 1.25 3.01 4.35 .80

Tail Cone .3 .08 .10 .11 0

Pylon 0 0 0 0 0


Total 65.3 .67

TABLE 13. CONFIGURATION 3C FUSELAGE WEIGHT AND COST DATA

FUSELAGE SECTION SECTION COST A FROM BASELINE COST A TOTALSECTION A WEIGHT MATERIAL TOTAL LABOR SECTION COST BASELINE(LB) (PERCENT) (PERCENT) (PERCENT) FUSELuAGE

(PERCENT)

Cockpit 20 .79 1.17 .96 .14

Mid-Cabin 39 1.65 9.17 8.47 4.04

Aft-Cabin 51 7.33 3.83 3.91 .71

Tail Cone 18 .91 - .14 .01 0

Pylon 0 0 0 0 0


Total 128 4.89

42

DESCRIPTION OF EVALUATION FOR FAIL-SAFETY,

SAFETY AND MAINTAINABILITY

FAIL-SAFETY

The three configurations meet the following fail-safety requirementsof the baseline UH60A:

1. The ability to sustain limit load with the loss of a single structuralcomponent.

2. Ease of access for inspection of the fuselage.

CRASH SAFETY/SAFETY

A relative evaluation was made of each configuration for crash safetyand safety. For crash safety the criteria were:

1. Noseover due to plowing2. Rollover3. High impact on landing14. Side impact

For safety:

1. Crew visibility2. Egress3. Hazards to protrusion

MAINTAINABILITY

An evaluation of maintainability was made for Configurations 1, 2 and 3based on the ease of maintaining the fuselages due to the overall shapeand size.

For Configuration I (cockpit only), the following is noted:

1. The number of window pieces is reduced from 9 to 6 compared to thebaseline; this should improve maintainability slightly.

2. There is a slight narrowing of the nose section which may cause relo-cation of components or reduce spacing between components. This mayreduce access and degrade maintainability.

The overall ranking is 5.

For Configuration 2, the following is noted:

1. The number of window pieces in the cockpit is reduced from 9 to 6compared to the baseline; however, some panels are larger than thebaseline which may cause them to be more susceptible to cracking.

143

IThe maintainability is decreased compared to the baseline.

2. Relocation of all electrical/avionics components from the cockpitsection is necessary. The wider fuselage should provide good spacefor access, and possibly shorter wiring harnesses may result. Themaintainability is Judged to be equal to the baseline.

3. Deeper tub section may provide space for components with better in-place access. Access to the tub section will depend on the design.The maintainability is judged to be equal to the baseline.

4. Reduced volume in the cockpit above the floor may reduce access in acrowded area. This is a negative effect on maintainability.

5. The cabin side windows may have less chance to fall out, thus fewermaintenance actions compared to the baseline.

6. The fuselage has a larger skin area; however, flat panels are perhapseasier to repair compared to the slightly curved panel of thebaseline.

The overall ranking is 5.

For Configuration 3, the following is noted:

1. The number of window pieces in the cockpit is reduced from 9 to 6compared to the baseline, hlowever, a few panels are larger thanthose of Configuration 2 and the baseline. The larger panels have anegative effect on maintainability.

2. Reduced cockpit space allows less access compared to the baseline.

3. Relocation of electrical/avionics from the cockpit section into amore restricted fuselage will probably reduce access,

11 4. Reduced tail-cone width reduces access to controls and components inthis area.

The overall ranking is 4.•.

The ranking method used was the same as that described in Reference 2.

The results are sjummarized in Tables 14 and 15.

Table 15 is a summary of weight and percentage cost changes compared tothe baseline uselage. Also shown are total wetted fuselage areas andranking for fail-safety, safety and maintainability.

44

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MODIFICATION OF CONFIGURATION 2

A study was made to determine the effect of l,,vering the cabin floor andceiling 6 inches. Lowering the floor 6 inches placed the floor at thesame waterline as the baseline fuselage; lowering the ceiling 6 inchesmaintained the same cargo door and opening as the baseline and Configura-tion 2. The change in weight due to lowering the floor is shown inTable 16 for the lower floor of Configuration 2 and the original floor ofConfiguration 2.

TABLE 16. FLOOR WEIGHT COMPARISON FOR CONFIGURATION Z

sTRUcTuR CONFIGURATION 2 LOWER FLOORA WEIGHT A WEIGHT

(LB) (L.)

Floor Panels 41.7 69.3

Beams & Supports 5.5 2.1

Total 47.2 71.4

Based upon the data of Table 16, lowering the floor 6 inches increasesthe fuselage weight of Configuration 2. Material and labor costs wouldbe increased due to the increase of floor weight and area. As a resultof increased weight and costs, Configuration 2 was not changed for thisstudy.

47

SELECTED CONCEPTS

Based upon the summary of Fuselage Weight and Cost (Tabla 15), Conucrt Awas selected for further evaluation. Concept A (floating frame concept)is the lowest in cost, compared bo Concepts B and C; also, the weight ofConcept A is equal to or less than that of Concepts B and C.

48

ADVANCED MATERIAL APPLICATION

Advanced composite materials can be used in the construction of the threefuselage shapes considered in this study. Studies, as reported in Refer-ences 2 and 3, have shown that the use of composite materials can reduceboth fuselage weight and cost. The fuselages of this study are relativelylightly loaded compared to fixed-wing aircraft. To efficiently useadvanced materials in the fuselages, very light composite skins are usedin the post-buckled stress state. Frames and stringers of the fuselageswould be constructed of stabilizcuI composites to develop the full-strengthcapabilities of the materials.

The stabilized composite structural. members and the thin composite skinscan be molded as a single structural assembly such as sides, top andbottom to form a fuselage section. The fuselages constructed of advancedmaterials are shown in Figures 8 and 9 for Configurations 2 and 3.

The cockpits of Configurations .1, 2 and 3 are constructed from twoassemblies, the cockpit enclosure and the lower cockpit tub. The cockpitenclosure is the framework and skins that support the windshields andwindows. The lower tub supports the seats, controls and equipment.

The enclosure framework is constructed of advanced materials using 7Tpercent Kevlar and 25 percent graphite/epoxy to replace the fiberglass/epoxy used for uonventional construction. The weight savings is 22percent as shown in Reference 2.

A structural weight breakdown for the conventional cockpit structures isshown in Table 17.

Tables 18, 19, and 20 compare the structural weight of the cockpit ofthe three configurations for conventional construction and advancedmaterial.

There are two basic types of frames used in airframe structure. One typeof frame is used to provide the shape of the cross section and to supportthe stringers. This type of frame is usually of thin gage aluminumformed as a "C" section. The other type of frame is used to transfer highconcentrated load to the fuselage shell. This type of frame is made up offorged sections, or built up from extruded members.

The tail cones are made up of the formed frames. The cabin and transi-tion sections are constructed of the built-up frames.

3. Rich, M. J., Ridgley, G. F., and Lowry, D. W., APPLICATION OFCOMPOSITES TO HELICOPTEF AIRFRAME AND LANDING GEAR STRUCTURES,Sikorsky Aircraft Div., NASA Technical Report CR-112333, NationalAeronautics and Space Administration, NASA-Langley Research Center,Hampton, Virginia, June 1973.

I49

The weight of a formed frame is .0128 lb/in. (Reference 2). A foamstabilized frame, which must match the stiffness of a formed aluminumframe, weighs .0115 lb/in. (Reference 2). The weight savings for thetail cone frames is 10%.

50

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

A2001 GRA PHITEf FACES -FOAN CORE F4 OOR

KCWVA#? VA/ER

37GA.J6O OZ'AM

KEVLAR' i(W40WCT WINXOW.MRAMIAISP/or c000%

,HIIGO DOR q- RCA

MWCR4PH/ TE aCE

FOAM 3 TAOI..ZED GRAPR'// ECRqA.H JA/03

WFlOURE A., ADVANCED MATERIALS A.PPLICATION F'OR CONFIflURATTON 2

MGB7 Je/PpoRT aeAMsFOAA4,1SrAI31/ZEO GeMAAONITE

--BOL TEO rAAWl SAOCOE

inv FO a~m STAB/UED GRAPHITE-FRA MES ý S TIR4VWGRS

/-%HAF r COMM9-RM

KZV4AR JAWVSFa4AI SM&IZtEO GR4APH/rt

FRAMES ý.S3TRWAGERS

7AIL CONE

,,54V/AIG t'OOR-. REP

FOAM CORE F4O~Rnma al LGrRAP/4/TE FACE 3

/KEVtAR LiME FOAM J 7MeL YZEC)D4nI~j~4~Q 34J.~Q &APH9E 8EA413

Pl or co ,i.or WINDOW FRAMING

HINGED DOXRS-REA'#~b

FIGUTRE 9. AD)VANCE~D MATERIA~LS APPLICATION V~OR CONFIC;uRATTON 3

/5

- E FRAA14j ICSI'Ii¶PPI&5

IDAGD - ,.-- FOAM JTA8/4 IM D WAP/1/TFRAAIEJ J 6TRIAOERS

MCOREC FL oRAP/-1/TE FACE3

OAMSA 3 AIZED S TA 26. T ~

ý (Z/APH/T - 5H J or rco veR - Rs

WFRAMING

KEVLAR SKIN

oDOORS5 -QREP M C4' SrABJ.LIZEL 6R4PH/TFoqAME., J JTRIAIGERS

:rA/L COV-r

TON P~OR CONFIGU'RATION 3

For highly loaded frames in the cabin and transition section, the weight

savings, using foam stabilized composite frames, is 33% (Reference 3).

BEAMS AND INTERCOSTALS

Beams are similar in construction to heavy frames. The weight savingsfor composite beams is the same as that for composite frames.

Intercostals are similar in construction to bent-up frames.

SKINS

Minimum skin gage is .025 in. aluminum

qult , 500 lb/in. 2024-T 6 in. stringer

F 24000 psi for woven Kevlar fabric at + 4505

u

25% Reduction for Environment (.75 Design Factor)

Fa Design a 24000 psi x .75 Design Factor a 18000 psi

treq. = 500/18000 = .02T7 in.

A layup of + 450 Kevlar 4 ply is minimum skin requirement (.010 in./ply, .050 lb/in,3 ). Percentage weight savings is:

Kevlar skin thickness x Kevlar density

(Al.min.m' ' )/100Aluminum skin thickness x Aluminum density)

.0~4 K x 00o02 x .050 )/100 = 20% weight savings for skins.05A x 00A

STRINGERS

The original composite stringers considered in Reference 3 were of uni-directional graphite and foam. The stringer weight was .0056 lb/in, for a3000-lb load capability. The composite stringer designed and tested asdiscussed in Reference 4 contained 2 ply at + 45Q of Kevlar to provideshear capability for the stringer. The weight of the stringer with theKevlar resulted in a composite stringer being of equal weight with aconventional stringer. These composite stringers are used to stabilize theKevIar skins and to provide axial load capabilities for the airframe.

55

ADVANCED MATERIAL APPLICATION COSTS

SThe costs for advanced materials were based on the material costs of

leferewce 2. These costs are as follows:

Graphite/Epoxy $20/lbKevlar-49/Epoxy $10/lbFoam $ 3/lb

The cost of materials for nonaffected and miscellaneous structures wasthat for mater.als of the baseline fuselage sections.

Labor costs for the fabrication of each section using advanced materialwere based on total labor costs for each section of± conventional oon-struction. The conventional labor costs were reduced 17-1/2 percent togive the labor costs for the advanced materials. The reduction was basedon the following:

. Current studies indicate that cockpit doors and cargo doors constructedof advanced material would reduce labor hours by 39 to 42 percent whencompared to built-up sheet metal doors,

. Studies conducted under a NASA program show that labor hours forproduction are reduced 25 percent with composites (Reference 4).

. Earlier studies (References 2 and 3) showed a reduction of 13 to 14percent for production composite fuselage labor costs compared to"conventional.

The results of the study using advanced material for each configurationare given in Tables 21, 22 and 23. Those results are based on thestructural concept A for fuselage construction of conventional materials,

A summary comparison of fuselage weight and cost changes for both conceptA and advanced materials is presented in Table 24.

4, Adams, K. M., and Lucas, J. J. , STUDY TO INVESTIGATE DESIGN, FABRI-CATION AND TEST OF LOW COSOT CONCEPTS FOR LARGE HYBRID COMPOSITEHELICOPTER FUSELAGE - PHASE I1, Sikorsky Aircraft Div. , NASA TechnicalReport CR-.145167, National Aeronautics and Space Administration, NASA-Langley Research Crntcr, Hampton, Virginia, April 1977.

56

.... ...... ... . . ....W

TABLE 17. COCKPIT STRUCTURAL WEIGHT*

STRUCTURE BASELINE CONFIG. 1 CONFIG. 2A CONFIG. 3AWEIGHT WEIGHT WEIGHT WEIGHT(LB) (LB) (LB) (LB)

Frames 43.0 41.8 45.3 46.3

Skins 32.9 32.3 34.6 34.7

Stiffeners 10.5 10.2 11.4 10.0

Floor & Support 18.9 18.1h 19.9 11.8

Crash Beams 33.6 32.7 35.3 31.7

Seat Beams 13.3 12.9 14.o 11.4

Cockpit

Enclosure 32.2 31.3 33.9 28.6

Windshield 57.5 55.9 60.5 80.8

Windows 15.8 15.4 16.6 13.9

Door 64.8 63.1 68,2 69.5

Steps .3 .3 .3 .3

Paint 2.4 2.3 2.5 2.3

Sealant 5.1 4.9 5.4 4.8

330,3 321.5 347.9 346.1

#Convontional Material

57

TABLE 18. COCKPIT STRUCTURAL WEIGHTOF CONFIGURATION 1A OFCONVENTIONAL MATERIALSAND ADVANCED MATERIALS

STRUCTURE CONFIG. 1A CONFIG. 1CONVENTIONAL ADVANCED MATERIALWEIGHT WEIGHT(LB) (LB)

Frame 41.8 38.2

Skins 32.3 26.1

Stiffeners 10.2 10.2

Floor & Supports 18.8 16,4Crash Beams 32.7 22.5

Seat Beams 12.9 8.9

Cockpit

Enclosure 31.3 24.8

Windshield 55.9 55.9

Windows 15.4 15.4

Doors 63.1 48.4

Steps .3 .3

Paint 2.3 2.3

Sealant 4.9 4.9

321.5 274.3

The material weight for Configuration 1 Advanced is:

Graphite/Epoxy 53.0Kevlar 92.9Foam 39.9Misc. 9.7

Windshield 55.9Nonaffected 22.9

274.3

58

TABLE 19. COCKPIT STRUCTURAL WEIGHTOF CONFIGURATION 2A OFCONVENTIONAL MATERIALS ANDADVANCED MATERIALS


Frames 45.3 41.3

Skins 34.6 27.7

Stiiffeners 11.4 11.4

Floor & Supports 19,9 17.5

Crash Beams 35.3 24.1

Seat Beams 14.o 9.6

Cockpit

Enclosure 33.9 26.8


Windows 16.6 16,6

Doors 68.2 51.2

Steps .3 .3

Paint 2.5 2.5

Sealant 5.4 5.4

347.9 294.8



Windshield 60,5Nonarfected 24.8

294.8

59

TABLE 20. COCKPIT STRUCTURAL WEIGHTOF CONFIGURATION 3A OFCONVENTI•,,AL MATERIALS ANDADVANCED MATERIALS


Frames 46.3 41.7

Skins 34.7 27,8

Stirffeners J0.0 10.0

Floor & Supports 11.8 9.4

Crash Beams 31.7 20.9

Seat Beams 11.4 7.5

Cockpit

Enclosure 28.6 22.3


Windows 13.9 13.9

Doors 69.5 51.5

Steps .3 .3

Paint 2.3 2.3

Sealant 4.8 4.8

:346.1 291.8



Windshield 80.6

Nonaffected 21.3

291.8

6o

TABLE 21. CONFIGURATION 1 ADVANVED MATERIALSFUSELAGE WEIGHT AND COST DATA



Cockpit - 56 61.0 - 17.5 - 6.5 - .91

Mid-Cabin 0 0 0 0 0

Aft-Cabi n 0 0 0 0 0

Tail Cone 0 0 0 0 0

Pylon 0 0 0 0 0


Total - 56 - .91

TABLE 22. CONFIGURATION 2 ADVANCED MATERIALSFUSELAGE WEIGHT AND COST DATA


(1,B) (PERCENT) (PERCENT) (PERCENT) FUSELAGE(PERCENT)

Cockpit - 36 72.3 - 17.1 - 3.7 - .53

Mid-Cabin - 50 44.8 - 13.2 - 7.3 -3.51

Aft-Cabin - 18 547.9 - 11.3 1.1 .20

Tail Cone - 14 208.2 - 16.8 - .7 - .04

Pylon 0 0 0 0 0


Total -118 -3.88

61

TABLE 23. CONFIGURATION 3 ADVANCED MATERIALSFUSELAGE WEIGHT AND COST DATA


(LB) (PERCENiT) (PERCENT) (PERCENT) FUSELAGE(PERCENT)

Cockpit - 3T 93.0 - 17.2 - .74 - .12

Mid-Cabin -153 8.97 - 17.8 -15.10 -7.20

Aft-Cabin - 41 541.0 - 13.8 - 1.50 - .28

Tail Cone - 25 196.5 - 17.4 - 2.10 - .13

Pylon 0 0 0 0 0


Total -256 -7,73

TABLE 24. SUMMARY OF WEIGHT AND COST DATA FORCONCEPT A AND ADVANCED MATERIALS

CONFIGURATLON CONCEPT A ADVANCED MATERIALSA WEIGHT PERCENT A WEIGHT PERCENT(LB) COST A (LB) COST A

Configuration 1 - 8 - .00063 - 56 - .91

Configuration 2 223 3.65 - 118 -3.88

Configuration 3 65 .62 - 256 -7.73

62

HELICOPTER DESIGN MODEL

INTRODUCTION

The design attributes of six aircraft designs, incorporating the six

fuselage concepts, were developed 'aging the Sikorsky Helicopter Design

Model (HDMý computer program described in Reference 2. The designattributes were based on the followinL parameters:

I. Fixed payload

2. Fixed hover performance

3. Fixed range

4. Estimate of vertical and forward drag

Two tasks were performed using the weight and cost changes for sixfuselages shown in Table 24 and preliminary estimates of forward andvertical drag shown in Table 25.

HDM RESULTS

The HDM was used to determine the payload/range of six aircraft at thetakeoff gross weight (TOGW) using the dynamic components of the baselineUH-60A aircraft. Table 26 shows the results for weight empty, fuel, pay-load, maximum cruise performance, vertical rate of climb, range, and theratio of flyaway cost to baseline flyaway cost. The payload range of thesix LRCS aircraft is plotted in Figure 10. The baseline UH-60A was notshown because it coincides nearly with LRCS Configuration 1.

Table 26 also shows that Configuration 2, at the same takeoff grossweight as the baseline UH-60A (16,450 pounds), has the lowest rate ofvertical climb and the lowest cruise speed. This is a result of theaerodynamic draC characteristics of a wide fuselage with sharp corners.

The 1DM was again used to develop design attributes trending solutions of

six aircraft that meet all UH-60A performance requirements except cruisospeed. Table 27 shows the results for takeoff gross weight at a constantpayload of 2644 pounds, weight empty, fuel, main rotor size, main gear-box design horsepower, maximum cruise speed, and the ratio of' flyawaycost to baseline flyaway cost.

Table 27 also shows that Configuration 2, at the same payload as the base-line Un-60A (2644 pounds), results in an aircraft with the largest grossweight, empty weight, and fuel weight, which meets the baseline performancerequirements.

63

.....l.. ...... . .. .

The significant differences in the attributes of Configurations 2 and 3of conventional or advanced materials, compared to the baseline, are dueto the aerodynamic characteristics of the configurations. The weightsavings gained by the use of advanced materials has little effect on thedesign/performance attributes when compared to the greater effects ofaerodynamic design.

TABLE 25. PRELIMINARY DR)AG ESTIMATE

FORWARD DRAG VERTICAL DRAGAT 0 ANG.LE• OF GROSS WEIGHT LESSATTACK, D/q TAIL ROTOR LIFT

CON .IGU'RATI ON 1'FT2) ( PERCENT )

Bal e•,) . i ne 2°6,13 3,37

)",.Conri.lguration i 27 .2U 3, 37

Cotnflguration 2 35., 2 7.20

, Con( gurat ion ~3 3:40 5.6o

I€

6~4

Em

' ~ in fn nr r m Lt\

H 7H H H H H

ujB

t.- 0 0 0Ir\ H HO

~~LHt,

bii

65

- H ~* H H H -

IIHHHHHC

~ ~ cI~~ 4*'i us LI\ NJ~ l111,c o 0 .4 O

H ~ ~ r C'JY\ 'J (

. . . '

66

4500

R~ICS 2

~00O N LYCS3

X]RCS 1 ADVANCI'U)

3500 LHCO" 0 AflVAflC1

LR33 ADVANC1EI)

I 3000

.0 10

Pt., T I0014 -N FULir WARM 1.11-

U 50 ~1.00 150 0K0 0

IRAN(Th NAUICr(AT, MrI:'

rinIMIRM 10. PAYLOAD) AND) FAN~x woi I.,nw LONAIAP Mi~l0" FIO BC 1ON II F;i T r~OTivp

6,r

TECHNICILL RISKS

The three configurations of convention&l construction offer no technical.risks for design and manufacturing.

The use of advanced materials such as graphite, Kevlar and foam forfuselage construction offer a few technical risks, such as:

* Tn absence of proper protective coatings, a possible decrease instrength properties of up to 10 percent for graphite and up to 25percent for Kevlar due to environmental effects of temperature andhumidity can occur.

• A foam capable of curing temperatures of up to 350°F will be requiredfor foail stabilized structures.

* Design data is required for:

1. Interaction strength of composite skin/stringer panels.

2. Crippling strength of composite flanges.

3. Effect of lucal stress concentration in primary structure forrouting of equipment, initial design and retrofit.

• Behavior of composite fuselage structure during a crash.

. Repair of damaged composite structure.

. Wire mesh in the Kevlar skins or external radar reflective coatingsmay be required for lightning protection and to reduce radarreflection of interne-] otructure under the skins.

68

CONCLUSIONS

Based on the results of thLi study, the following conclusions are made:

1. The use of advanced materials can result In both weight and costsavings over the baseline fuselage, even with the most severe changein LRCS configurations presented.

ý?. Without the use of advanced materials, the LRCS Configurations 2 and3 significantly increase both weighlt and cost of the total aircraftcompared to the baseline UI160A.

3, Minor changes to the nose section of Conflguraltion I result innegligible fuselage aLiffrorence to the weight and cost of the fuselage.

i4. Consid•r'ation of the total aircraft attributes show that verticaldrag penalties appear to be of greater magnitude than the structuralweight changes involved with the fuaelages of Configurations 2 and 3.Even with the use of advanced materials, the vertical drag penaltyexceeds any weight savings

Further conclusions from the detailed aerodynamic analysis presented inAppendix A shows:

1. The forward drag estimates for LRCS Configurations 2 and 3 were foundto be higher than the values obtained through the use of' the detailedand more sophisticated aerodynamic analysis presented, The highervalues were the result of the unavailability of prior test data ofthe representative shapes of the two configurations for comparisonpurposes. Also, interference factors which were applied to theoriginal estimates to uccount for interference of' the two shapes

and the main rotor pylon may have been exaggerated.

2. The detailed aerodynamic analysis increased the cruise speed ofConfiguration 2 from 133 knots to 14+0 knots and the cruise speed ofConfiguration 3 from 135 knots to 136 knots.

3. Vertical drag estimates were not significantly changed by the de-tailed aerodynamic analysis, and the low vertical rates of climbfor Configuration 2 and 3, presented in Table 26, were substantiated.

69

IvI

REFERENCES

I. UH-60A VOLUME 3, PART 1-IIl WEIGHT AND BALANCE dated September 27,1976.

2. Rich, M. J., INVESTIGATION OF ADVANCED HELICOPTER STRUCTURAL DESIGNS,VOLUME 1 - ADVANCED STRUCTURAL COMPONENT DESIGN CONCEPTS STUDY,Sikorsky Aircraft Div., USAAMRDL-TR-75-59A, U.S. Army Air MobilityResearch and Development Laboratory, Fort Eustis, Va. 23604, May 1976,AD A026246. I

3. Rich, M. J., Ridgley, G. F., and Lowry, D. W., APPLICATION OFCOMPOSITES TO HELICOPTER AIRFRAME AND LANDING GEAR STRUCTURES,Sikorsky Aircraft Div., NASA Technical Report CR-l12333, NationalAeronautics and Space Administration, NASA-Langley Research Center,Hampton, Virginia, June 1973.

4. Adams, K. M., and Lucas, J. J., STUDY TO INVESTIGATE DESIGN,FABRICATION AND TEST OF LOW COST CONCEPTS FOR LARGE HYBRID COMPOSITEHELICOPTER FUSELAGE - PHASE II, Sikorsky Aircraft Div., NASA TechnicalReport CR-1h5167, National AeronauLics and Space Administration,NASA-Langley Research Center, Hampton, Virginia, April 1977.

5. Sheehy, T. W. , and Clark, D.R., A METHOD FOR PREDICTING H1ELICOPTERHUB DRAG, USAAMRDL-TR-75-48, L'ustis Directorate, U. S. Army AirMobility R&D Laboratory, Ft. Eustis, Virginia, January 1976, AD A021201.

6. Keys, Charles, and Weisner, Robert, GUIDELINES FOR REDUCING PARASITEDRAG, Boeing Vertol Company, Journal of the American HelicopterSociety, January 1975.

7. Delaney, Noel K. , and Sorniuon, Norman E. , LOW SPEED DRAG OF CYLINDERS.*

01' VARIOUS SH1AP'ES, NACA TN-3038, November 1953.

8. Hoerncer, Dr. ing S. F., FLUID DYNAMIC DRAG, Second Edition, Bricktown,New Jersey, IHoerner Fluid Dynamics, 1965.

9. Werner, J. V., and Flemmlng, R. J., YUH-60A QUARTER SCALE WIND TUNNELTEST REPORT, Sikorsky Aircraft, SEIR-70531, May 1, 1973.

10. Barnard, R. u., YUH-6QA/T700 ENGINE IR SUPPRESSOR FULL SCALE PROTOTYPETEST REPORT, Sikorsky Aircraft, SER-70094, June 18, J.976.

11. Prime Item Development Specification for UH-60A Utility TacticalTransport Aircraft System OPQ, RFQ, DAAJ01-77-C-O001 (PGA), Part 1,U. S. Army Aviations Systems Command, St, T.116, Mlmrouri,

ýIovember 1976.

TO

• • I I I I Ij

APPENDIX A - AERODYNAMIC ANALYSIS OF LOWRADAR CROSS SECTION CONFIGURATIONS

INTRODUCTION

An analytical investigation of the aerodynamic loads, vertical drag andmission performance of three proposed low radar cross section configura-tions has been performed.

The Sikorsky Wing and Body Aerodynamic Technique (WABAT), in combinationwith the Sikorsky Automated Paneling Technique (APT),and empirical methodswere employed in the analysis of the aerodynamic loads. WABAT is a three-dimensional potential flow analysis which is capable of calculating theaerodynamic flow and surface pressures about arbitrary lifting bodies.This enables the aerodynamicist to predict body airloads and regions ofhigh dyn.'amic pressure and permits the evaluation of concepts to minimizedrag,

The strip analysis method used to calculate the airframe vertical drag isa semi-empirical approach based on experimental drag coefficients androtor wake flow surveys. This method establishes the flow environmentaround elements of the airframe to yield element drags which are thensummed to yield total airframe drag.

Aircraft performance was based on the Aircraft Trim Adjusted PerformanceAnalysis (ATAP), which uses nondimensional main and Lail rotor perform-ance,.fuselage attitude, airframe lift and drag data, system losses andpowerplant data to compute system power requirements and range character-istics.

PROCEDURES

Using section coordinates from the Government-f.urnished mold lines anddrawings prepared under Task I, the representative geometric models ofthe three LRCS configurations and the baseline UH-60A were developedwith the use of the Automated Paneling Technique (Figures A-I thru A-8).The panel geometries thus generated were input to the WABAT program whichcalculated the potential flow solution for each configuration, yieldingthe surface pressure and forces and pitching moments for each body panel.The geometry and aerodynamic methods and computer programs are describedin greater detail in Reference 5.

5. Sheehy, T. W. and Clark, D. R., A METHOD FOR PREDICTING HELICOPTER HUBDRAG, USAAMRDL TR-75-48, Eustis Directorate, U. S. Army Air MobilityR & D Laboratory, Ft. Eustis, Virginia, January 1976.

71

S..... .......................... ..... .. . ...... .. ..,-,- - - - ---

The lift, drag and pitching moment of each LRCS configuration were cal-culated based on the element surface forces generated by WABAT. Thevariation of lift with angle of attack was determined by calculating thedifference between the total body lift of each LRCS configuration and thetotal body lift of the basic UH-60A at each angle of attack. This diff-erence was then applied to the total body lift of the UH-60A configurationobtained from test data. The same procedure was applied to the calcula-tion of pitching moment.

The total body drag of each LRCS configuration was computed as an in-crement to the basic UH-6OA. The total body drag increment is definedas the sum of the drag increments due to changes in forebody drag, pylon"drag, drag as a result of contraction in the transition region, and dragas a result of wetted area. The drag increment due to contraction wasbased on data presented in Reference 6. The variation of drag withangle of attack was also based on data presented in Reference 6.

The vertical drag of each LRCS configuration was generated using thestrip analysis method. The method requires three basic sets of data toprovide the necessary information for the calculation of drag: airframeelement areas, drag coefficients and cylindrical coordinates describingthe location of each of the airframe centroids. Two-dimensional dragcoefficients based on References 7 and 8 were used for this calculation,In some instances, three-dimensional drag coefficients were used for thoseelements exhibiting three-dimensional effects.

.Using the airframe lift and drag results and the baseline UR-60A power-plant data in combination with the Aircraft Trim Adjusted PerformanceAnalysis program, the mission performance of each LRCS configuration wasgenerated.

SUMMARY OF DATA PRESENTED

Surface pressures generated by WABAT for the proposed LRCS configurationshave been compared to the baseline UH-60A surface pressures over a rangeof angles of attack ( a80 Ix • + 80) along the top ( as w 00), bottom

(as a 1800), and lateral ( c8 * 90c) centerlines at waterline 219 and arepresented in Figures A-9 thru A-53,

(6. Keys, Charles, and Weisner, Robert, GUIDELINES FOR REDUCING PARASITEDRAG, Boeing Vertol Company, Journal of the American HelicopterSociety, January 1975.

7. Delaney, Noel K., and Sorenson, Norman E., LOW SPEED DRAG OF CYLINDERSOF VARIOUS SHAPES NACA TN-3038, November 1953.

8. Hoerner, Dr. Ing S. F., FLUID DYNAMIC DRAG, Second Edition, Bricktown,New Jersey, Ikoerner Fluid Dynamics, 1965.

Lifts, drags, and pitching moments for each of the LRCS configurationshave been compared to the baseline lift, drag and pitching moment over arange of angles of attack (-80 5 a +80) and are presented in FigureA-54 thru A-56. The UH-60A lift and drag data were based on wind tunneltest data obtained from References 9 and 10.

The comparison of the body lifts, without empennage, for the UH-60Aand the three LRCS concepts shown in Figure A-54 demonstrates that boththe LRCS2 and LRCS3 configurations produce approximately 4 square feet ofadditional download at representative cruise angles of attack (-20 to -5o).The LRCS1 configuration demonstrates no significant change in lift com-pared to the UH-60A. While the additional download generated by theLRCS2 and LRCS3 configurations is not insignificant, the primary impacton mission performance is the increased drag of these configurationsshown in Figure A-55. The data shown in this figure include the drag ofthe empennage, all external protuberances, and momentum losses.

The drag increase with angle of attack variation for the LRCS2 and LRCS3configurations is greater because of their difference in cross-sectionalshape compared to the UH-60A or the LRCS1. Although the LRCS1 shows nosignificant change in drag from the UH-60A, the LRCS2 and LRCS3 demon-strate minimum drag values 11% and 17%, respectively, greater than thebaseline UH-60A.

The predicted pitching moments of the three LRCS configurations arecompared with the baseline UH-60A (without empennage) in Figure A-56.The pitching moment of both the LRCS1 and LRCS3 configurations are notsignificantly changed from the UH-60A. The LRCS2 pitching moment trend,however, demonstrates an increase in slope resulting in approximately an83% increase in basic fuselage instability.

In addition to the unfavorable impact of the LRCS3 configuration onstability, the increased drag of both the LRCS2 and LRCS3 configurationswill have an unfavorable effect on the dynamic pressure loss in theempennage region and consequently on the horizontal tail effectiveness.This impact has been estimated by assuming that the dynamic pressure lossof the LRCS2 and LRC03 configurations compared to the UH-60A is propor-tional to the increased drag of the basic configuration not includingthe empennage drag or the drag due to momentum losses. Based on this,the tail effectiveness compared to the UH-60A is reduced by 21% for theLRCS2 and by 28% for the LRCS3.

9. Werner, J. V., and Fleecing, R. J., YUH-60A QUARTER SCALE WIND TUNNELTEST REPORT, Sikorsky Aircraft, SER-70531, May 1, 1973.

10.Barnard, R. S. YUH-60A/T700 ENGINE IR SUPPRESSOR FULL SCALE PROTOTYPETEST REPORT, Sikorsky Aircraft, SER-70094, June 18, 1976.

73

Based on the calculated fuselage pitching moment results and the assess-ment of tail effectiveness, the LRCS2 and LRCS3 configurations willdemonstrate a substantial reduction in static stability compared to theUH-60A and will require redesign of the horizontal stabilator. The LRCS3would therefore be penalized by the requirement for increased horizontalstabilator area and the associated weight and drag penalty.

The vertical drag of each LRCS concept and of the baseline UH-60A arepresented in Tables A-I thru A-4. The VOR/LOC and FM homing ante!nnas ofthe baseline UH-60A were included in the calculation of vertical dragfor each of the LRCS configurations for the purpose of comparison. Dragcoefficients were estimated on the basis of data presented in References

7 and 8. The vertical drag and parasite drags of each configuration aresuimmarized in Table A-5.

The mission performance for each of the LRCS configurations was generatedon the basis of the lift and drag variations shown in Figure A-53 andA-54 and the T700-GE-700 powerplant data used for the baseline UH-60A.The performance was computed with the use of the Aircraft Trim AdjustedPerformance program and is presented in Table A-6. Maximum cruise speedwas determined by using the maximum continuous power rating at a pressurealtitude of 4000 feet for a 950F day. The hover and one engine in-operative (0EI) service ceilings were calculated for a 95OF day using95% of the intermediate rated power and 100% of the intermediate ratedpower, respectively. Endurance was calculated on the basis that thefuel capacity of the three LRCS configurations remains the same as theUH-60A fuel capacity. The endurance mission included the following fourcategories: 8 minutes at ground idle power, 20 minutes at maximumcontinuous power, cruise at 145 knots, and 30 minute reserve at 145 knots.The endurance was calculated at a pressure altitude of 4000 feet for a95°F day. Calculations were based on a gross weight of 16,450 lb Thebaseline mission performance was obtained from Reference 11.

CONCLUSIONS

Comparison of the UH-60A attributes with the calculated results for thethree proposed low radar cr'oss section configurations indicate that:

1. A change in the UH-60A configuration to reflect the LRCSIconfiguration caused a decrase in the drag of the UH-60A at c - 00.This decrease was refleacted through a range of angles of attack(+80 < a < - 80) with the exception of angles of attack of +80 and-80 respectively, where no change in drag was exhibited.

11. Prime Item Development Specification for UH-60A Utility TacticalTransport Aircraft System OPQ, RFQ, DAAJO1-77-C-000l (PCA), Part I,U.S. Army Aviations Systems Command, St. Louis, Missouri,November 1976.

74

Configuration changes reflecting the LRCS2 and LRCS3 configurationshowed increases of 11% and 17% respectively, of the UH-60A drag ata = 00 and as high as 21% and 29%, respectively, at angles of attackof +80 and -8o. These changes do not reflect the required increasein stabilator size which would resu.lt in an additional drag penalty.

2. The LRCSI configuration produces an insignificant change in liftcompared to the UH-60A while the LRCS2 and LRC83 configurations reflectan additional download of 40% of the UH-60A download at (% -- 4.

3. Compared to the UH-60A, the LRCS2 and LRCS3 configurations demonstratea substantial reduction in static stability which will require a re-design of the horizontal stabilator. This is primarily due to areduction in tail effectiveness and an increase in fuselage insta-bility. The reduction in tail effectiveness was estimated as 21%for the LRCS2 concept and 28% for the LRCS3 configuration. Minorchanges in the pitching moment of the LRCS1 and LRCS3 configurations,compared to the UH-60A were calculated, however, the LRCS2 trenddemonstrated an increase in slope resulting in an 83% increase inbasic fuselage instability.

4. The three proposed LECS configurations demonstrated an increase invertical drag of 27% for the LRCSl, 106% for the LRCS2, and 70% forthe LRCS3, of the basic UH-60A vertical drag. This increase invertical drag is a direct result of the additional cross sectionalarea and sharp edges presented to the downwash of the main rotor bythe LRCS2 configuration, In the case of the LRCS3 configuration, thewedge type of cross sectional area of the cockpit, the additionalarea of the main landing gear support structure and the presence ofsharp edges presented to the main rotor downwash resulted in an in-crease in vertical drag as indicated. Finally, the increase invertical drap for the LFCSl configuration is a direct result of thewedge type Lross sectional area of the cockpit.

5. In comparison to the UH-60A, no change in the maximum cruise speedand endurance of the LRCSl configguration was observed, however, aloss of 217 feet in hover ceiling and 12 feet in OEI service ceilingwas exhibited. The performance of the LCS2 and LRCS3 configurationsdid not meet the performance requirements of the basic UH-60A,Decreases of 7 and 11 knots in maximum cruise speed, 805 and 520feet in hover ceiling, 248 and 740 feet in OEI service ceiling, and0.14 and 0.19 hour in endurance were calculated for the LICS2 aidLRCS3 configurations respectively,

75

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TABLE A-5. SUMMARY OF PARASITE AND VERTICA;L DHA(G

PARASITE DRA( 2 0 VERTICAL D)RAGCONFIGURATION @ 0• 0, O I (1 . W

UH-60A 26.58 3.40

LRCSI 26,32 4.33

LRCS2 29.49 7.02

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l,':Fiure A-54. VATIATTON O1 LiPP WITH ANGLE 0, A'.I"ACK FOR v'mIi uI-60A UTTAl'.3,

MCE3) I RCS ],I o ,A L •CON 1 IG[U!R ,I NL

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40

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FigUre A-55. VARIATION OF DRAG WITH ANGLEI OF ATTLACK FOR THE UH-160A UTTAS,LRCS1, UiCS 2, LRCS3 CONFIGURATIONS

136

300 1 . 1 1ZERO ANGLE OF YAW

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100 t7-

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* BSEDON TEST DATA

-5001 1- - -

-12 -8 -4 0 4 8 12ANGLE OF ATTACKa ,DEG

Figure A-56. VARIATTON OF PITCHING MOMENT WITH ANGLE OF ATTACK FOR THEtI-60A Ufi1TAs, LRCS1, TRC32 and TRCS3 CONFIGURAI'IONS

137

LIST OF SYMBOLS

A Elemental area

C p Surface pressure coefficient

D/q Nondimensional Rirframe drag

Fs Design shear strength, lb/in. 2

Fsu Ultimate shear strength, lb/in. 2

GW Gross weight, lb

h/R Nondimensional distance between fuselage surface and rotor plane

1/q Nondimensional airframe lift

MAI Nondimensional airframe pitching moment

NZ Ultimate vertical load factor,specified design limit loadfactor at design gross weighttimes 1.5

PLO Payload, lb

q Dynamic impact pressure, lb/ft2

quit Ultimate shear flow allowable, lb/in.

R Main rotor radius

r/R Nondimensional blade radial station

S Wetted surface area, ft2

t req Required thickness, in.

TOGW Takeoff gross weight, lb

WE Weight empty, lb

X Fuselage buttline

Y Fuselage waterline

Z Fuselage body station

U Angle of attack

138

LIST OF SYMBOLS (cont'd.)

Ma Bodyline angle at Y-219 in.

Ap Pressure change from atmospheric pressure, lb/in. 2

Angle of yaw

- I

I

1-39 •,,•

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