Hypersonic cruise vehicle wing structure evaluation - Volume 3, sections 23-27

CONTENTS Page Section Volume 1 V Summary vii Introduction 1 -i Trajectory Analysis 2 -i 2 Vehicle Loads 3 -i Aerodynamic Heating Analysis 4-i 4 MEI terials Evaluation 5-i .

5 Material and Process Development Testing 6 -i 6 Structural Acalysis Model 7 -i 7 A Plane Strain Analysis for Determining Thermal Stresses 8-i ' Structural. Internal Loads (Air and Thermal) 9 -i 9 Internal Temperature Analysis 10-i 10 Optimization Procedure for Panels of Monocoque Structure Volume 2 11 Optimization Procedure for Panels of Circular-Arc 11-1 Corrugation Shear Webs 12 Optimization Procedure for Panels of Semimonocoque 12-1 Structure 13-1 Primary Structure Sizing and Weights 14-1 14 Panel Flutter 15-1 15 Vehicle Flutter 16-1 Sonic Fatigue 17-1 1 7 Fatigue 1 8 4 18 Creep 19 -i 19 Optimization Procedure for Heat Shields 20-1 20 Heat Shield Sizing and Weights 21-1 Leading Edge Analysis 22-1 22 Tot3.1 Wing Weight Analysis Volume 3 23 -i 23 Cost Analysis 2 4 4 24 Performance Analysis 25-1 25 Reliability Analysis 26-1 26 Interactim Analysis 27-1 27 Structural Element Testing iii SUMMARY An analytical and experimentai evaluation was perforrned for several promising structural concepts to provide the basis of minimuin total-system-cost for selection 8 hypersonic vehicle wing.

of the best concepts for the design of Results, procedures, and principal justification of results are presented in reference 1.

Detailed substantiation data are given herein. Each major analysis is presented in a separate section. Vehicle loads and temperatures are given with each structural analysis that influences weight. In addition to the weight analysis, fabrication cost, perforniance penalties (surface roughness drag), reliability, and total-system-cost analyses are presented.

Reference 1. Plank, P. P.; Salmta, I. F.; Davis, G. W.; andHichie, C . C.: Hypersonic Cruise Vehicle Wing Structurie Evaluation, NASA CR-1568, 1070.

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INTRODUCTION The utility of a hypersonic cruise vshicle depends upon a low structural mass fraction in a high-temperature envirownent. Unfortunately, this requirement exceeds the limits of state-of-the-art structurzs. The only hypersonic structures flown to date have been the X-15 research airplane and the ASSET unmanned lifting reentry test vehicle, b t h of which are unsuitable for cruising flight.

For the past several years, the NASA kngley Research Center and other agencies have been investigating proniising structural concepts, such a s those discussed in references 2, 3 , and 4, and the 1967 Conference on Hypersonic Aircraft Technology (ref. 5) was devoted to the subject.

An evaluation was performed of i_i-omi:;ing wing structure concepts to the same in-depth analyses, including a l l known envil mmental structural considerations that could affect the four evaluation factors: weight, cost, performance, and reliability.

These factors were then interacted in a total-system-cost study f o r a system range- payload capability of 205 billion ton-miles to provide the basis for selecting the best structural concept for the wing structure of minjmum total-system-cost.

are reported in reference 1. This Results of this structural evaluation reference also includes the procedures and principal justification of results, whereas this report gives detailed substantiation of the results iu reference 1.

Principal analytical and test efforts are presented in separate sections. This report is bound as three separate volumes.

REFERENCES Heldenfels, R. R. : Structural Prospects for Hypersonic A i r Vehicle ICAS 2.

paper, 1966.

Plank, P. P.; and MacMiller, C. I.: Analytical Investigation of Candidate 3.

Thermal-Strwtural Concepts Applicable to Wing, Fuselage, and Inlet Structure of a Manned Hypersonic Vehicle. AFFDL-TIC-66-15, 1966 (conf).

4. Plank, P. P. : Hypersonic Thermal-Structural Concept Trends. SAE paper 660678, 1966.

5. NASA-SP-148 (Conf). Conference on Hypersonic Technology, Ames Research Center, 1967.

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ACKNOWLEDGEMENT This investigation was conducted under NASA Contract No. NAS1-7573, Research and Development Program for Developrcent and Validation of Structural Concepts for a Hypersonic Cruise Vehicle Wing Structure.

The study was originated at the Lockheed-California Company, Burbank, California and completed at Lockheed Missiles & Space Company, Sunnyvale, California.

P. P. Plank was the Program Technical Manager, I. F. Sakata and G. W. Davis the Project Engineers, and C. C. Richie was head of structural concept optimiza- tion. The other contributors to the program are acknowledge at each section.

Four Lockheed-California Company personnel acted in an advisory capacity.

They are L. W. Nelson, Structure Division Engineer, Structures Division; E. J. Himmel, Department Manager, Stress Analysis; W. J. Crichlow, Depart- ment Manager, Advanced Materials and Structural Mechanics; and M. G. Childers, Manager , Physical Sciences Laboratory Development Engineer.

Dr. M. S. Anderson, L. R. Jackson, and J. C, Robinson of the Structures Research Division, NASA Langley Research Center, Hampton, Virginia, were the Program Manager, Technical %presentative of the Contracting Officer (TRCO), and Assistant TRCO, respectively, for the project.

Section 23 COST ANALYSIS b Y E. W. Reed, D. E. Sherwood, and I. 3 ' . Sakata 23 -i PRECEDING PAGE BLANK NOT FILMED.

23-1 INITIAL PANEL SCREENING COSTS 23-1 HEAT SHIELD COSTS 23 -2 LEADING EDGE COSTS 23 -2 WING S E G M E N T COSTS Substructure Costs 23 -3 23 -4 Monocoque Concept Costs Semimonocoque Concept Cost6 23 -7 S t a t i c a l l y Determinate Zoncept Costs 23 -9 Wing Segment Cost Summary 23 -9 23-10 TOTAL WING COSTS 23-12 Labor and M t e r i a l Cost Estimating Relationship 23 -I2 Wing Structure labor Costs 23-12 Wing Structure m t e r l a l Costs 23-12 Tooling Cost Estimating Relationship T o t a l Wing Cost S m r y 23-13 23 -13 VEHJCLF: PRODUCTION COSTS 23-nii PRECEDING PAGE BLANK NOT FILMED, TABLES I n i t i a l screming costs of s t r u c t u r a l panels 23 -1 23 -2 Sulumary of heat shield cost evaluatim factor data f o r 100-vehicle production run 23 -16 Leading-edge cost evaluation f o r wing evaluation area w i t h 23 -3 100-vehicle production run 23 -17 23 -4. Weights f o r main wing manufacturing segment 23 -18 Waffle concept substructure manufacturing costs (dollars) 23 -20 23-5 23 -22 Waffle panel fabrication and i n s t a l l a t i o n costs (dollars) 23 -6 Waffle substructure, panel, and heat shield fabrication, 23-7 assembly and installation costs ( d o l k r s ) 23 -23

23 -8 Summary - Waffle concept main wing segment costs f o r various

23 -24 aspect r a t i o s Monocoque concept substructure manufacturing 23 -9 cost (dollars) 23-25 23 -10 Monocoque concept panc'' fabrication and instal.lat,ion 23 -26 costs (dollars) 23-11 Monocoque concept substructure, panel, hcat shield (including insulation) fabrication, assembly and instal-lation costs (dollars) 23 -12 Monocoque 2oncep-t main wing segment cost 23-13 Monocoque concept substructure manufacturing costs Nonocoque concept panel fabrication and.i n s t a l l a t i o n costs 23-14 23-1.5 Monocoyue concept t o t a l manufacturing c o s t s Semimonocoque substructure manufacturing cosl.:; ( d o l h r s ) 23-16 Semimonocoque substructure, panel, heat shield (including 23-17' insulation)fabrication, assembly and i n s t a l l a t i o n costs ( d o l l a r s ) 23-33 23-18 Semimonocoque panel fabrication and. instalLczti,on c s s t s (dollars) 23 -34 I .

23-V Page 23-19 Semiaonocoque concepts main wing segment costs 23-35 23-20 S emimonocoque concept substriicture manufacturing costs 23 -36 23-21 Spmimonocoque concept pane1 fabrication and i n s t a l l a t i o n costs 23 -37 23-22 Semimonoc ,:. le concept heat shield/insulation %bricat,lon and installa' I costs 23-38 Semimonocoque concept t o t a l manufacturing costs 23-23 23 -39 23 -24 23 -40 S t a t i c a l l y det ermiriet e aiibbsti*ucturemanufacturing costs S t a t i c a l l y determinate panel f'cLbrication ar,d i n s t a l l a t i o n 23-25 23 -41 costs ( d o l h r s ) 23-26 S t a t i c a l l y determinate concept total- substrwture, panel, heat -shield fabrication, assembly, and iastallat ion costs 23 -k2 (dollars )

23-27 Summary - s t a t i c a l l y determinate concept main wing segment

costs 23 -43 23-28 S t a t i c a l l y determinate concept dist-4buted substructure 23 -44 &sst . d y and installatiozz costs wne1, heat -shield fabrication

Summary - s t a t i c a l l y determinate concept t o t a l mnuf'acturing

23-29 costs 2 3 4 5 23 -L 6 23-30 Structure concept manufacturing coste for main ving segment

Summry - structure concept manufacturing costa 23-47

23-31 23 -48 23-32 Total wing weight summary k c h i n e parts estimated Labor,and material costs f o r overall 23-33 king structure 23 -49 Labor and material cost estimrzt.ing relationships i'or overall 23-34 wing structure costs 23-50 23. '2 S t a t i c a l l y determinate concept slip-joint assemhly costs 23-35 23-53 Total wing structure labor costs 23-36 23-37 Total wing structure material cists * 23 -54 I Tooling cost e s t i m t i n g relationship f o r overall wing 23-38 st 1-uc t u e 23 -55 23-39 Overall tooling cost estimating relai;ionship 23 -56 23-40 Total wing structure mnukcturing costs baseline 550,OO-lb alrplane (100 vehicles) 23-57 23-kl Vehicle cost e s t i w t i n g factors 23-58 23-42 Overail wing cost e s t i m t b g factors 23 -60 23-43 Tatal vehicle proiluction costs (100 vehicles) 23-61 PRECEDING PAGE BLANK NOT FILMED.

ILLUS WATIONS page 23 -1 U n i t heat shield cost versus number of a i r c r a f t 23-62 23 -2 Hypersonic crulse airplane - manufacturing segments 23 -63 anufacturing segment, w i n wing 23-64 23 -3 23 -4 Structure concept arrangement 23 -65 C ircular-arc corrugation web element fabrication 23 -66 23 -5 Typical chordvise r i b assembly 23 -6 23 -67 Typical spanwise beam segment 23 -68 23 -7 23-8 Typical intersection d e t a i l 23 -69 23 -9 Typical web and beam cap intersection detail 23 -70 23-10 Right wing leading-edge cap 23-71 23-11 Binel aspect r a t i o study, monocoque waffle primary structure concept 23-72 23-12 mill wing segment weight variation with aspect r a t i o 23 -73 Labor cost vs aspect r a t i o 23-13 23 -74 23-14 b t e r i a l costs vs aspect r a t i o 23-75 23-15 Tooling costs vs aspect r a t i o 23 -76 23-16 Total manufacturing cost vs aspect r a t i o 23-77 23-17 Wing geometry and reference areas for baseline vehicle 23 -78 3 - 1 8 S t r u c t w e general assembly 23 -79 PRECEDING PAGE BLANK NOT FILMED.

SYMBOLS Costing zones defined i n figure 23-17 A, B, c, D,E x and y distances between simply supported edges of plate a , b Panel aspect ratio a/b BIs Butt l i n e CEX? Cost estimating relationship ECM ELectrochemical m i l l i n g \=W Gross weight L Length S Surface area Wing station Sta Subscripts Denotes leading edge le P Denotes panel S Denotes sub structure T Denotes total

Section 23 COST ANALYSIS The basis f c evaluating and rating the structure concepts i s minimum t o t a l systeri cost; therefore, manufacturing cost; information f o r t h e various vehicle compnsnts of t h e baseline vehicle (gross weight = 550 000 lb). The cost estimates 'exdressed i n 1968 dollars) were determined f o r engineering purposes only, using current labor r a t e s and material pi-ices. The data gener- ated a r e consider-.d sufficiently accurate t o provide v a l i d cost i-nfomation so t h a t a r e l a t i v e couparison of concepts can be made.

F a c i l i t i e s and process developnent; costa vere not included. It was further assumed that clean-room conditions would b? available f o r fabricating the vehicle components, that suitable controlled-atmosphere furnaces and process baths have been installed, and that required special equipment and machine tools w i l l have been developed and installed.

I n i t i a l Panel Screening Costs Comparativ?. costs for t h e candidate s t r u c t u r a l panels and heat shield as applicable, were detemined on the basis of t h e aforementioned combinations, premise. These costs were determined by a detailed production cost analysis of typical panels sized f o r representative hypersonic cruise vehicle loads and included recurring and nonrecurring costs encompassing material, labor, an& tcoling f o r 1000 production units. The costs presented i n tab17 23-1 include panel closeouts and applicable manufacturing methods using Rene 41 and Ffaynes 25 alloys. The semimonocoque spanwise concept panel costs r e f l e c t representative values f o r the s t a t i c a l l y determinate concept. The manufacturing methods f o r t h e monocoque waffle and honeycomb are discussed e a r l i e r i n the monocoque weightt, Q-ction (section 13). The semimonocoque concepts r e f l e c t production techniques discussed i n section 27.

Heat Shield Costs CoBt information was determined for two refurbishable and two permanently attached heat shield conn,e@x discussed i n d e t a i l i n the heat shield sizing and weights section (section 20). A11 heat shield concepts were evaluated on the tubular panel (size: 92 i t c h x 46 inch), The r e m b i s h a b l e heat shield included t h e following: 1 . Corrugated skin with multiple supports 2. Flat skin dimpled-stiffened clip-supported 23 -1 The permanently attached heat shields included the two variations of the modular heat shield concept: 1. Modular, simply supported 2. Modular, cantilevered was conducted i n sufficient d e t a i l t o estimate t h e tool- This cost study ing required f o r fabrication and assembly. Figure 23-1 indicates r e l a t i v e costs of the four concepts evaluated, including labor and material. Table 23-2 p.esents costs i n dollars per square foot for 100 vehicles. The r e s u l t s of t h i s cost evaluation indicate t h a t the corrugated heat shield i s lowest i n cost.

Leading Fdge Costs The evaluation of leading edge concepts was made f o r both the continuous and segmented designs. The leading edge cost data encompassing labor, material and t o t a l cost requirements considering 100 vehicles i s presented i n table 23-3 i n terms of $/lb and $/linear foot. These data indicate t h a t the segmented leading edge concept provides the lower cost.

Wing Segment Costs Manufacturing costs of the wing structure concepts wzre determined on t h e basis o f detailed analysis of a typical manufacturing segment using 1968 labor The detailed cost analysis included (1) substructure r a t e s and material prices.

(2) panel fabrication, assembly, and installation; fabrication and assembly; and (3) heat shield fabrication, assembly, and i n s t a l l a t i o n with tooling re- quirements amortized over 100 vehicles.

To f a c i l i t a t e costing, t h e vehicle structure was divided into typical manufacturing segments as shown i n figure 23-2, w i t h the detailed analysis confined t o the main wing manufacturiiig segment. A typical arrangement and geometry f o r the main wing manufacturing segment i s shown i n figure 23-3.

This segment (one-half shown) consists of 1874 square f e e t of planform area located between Station 2136 and Statl.on 2506. The segment is further divided into 3 zones (A, B, and C). These zones represent typical types of structures The basic found i n the segment as determined by detailed structural analysis.

elements consist of the substructure, structural panels, and heat shields (including insulation). This latter was costed i n d e t a i l f o r t h e heat shield cost evaluation presented e a r l i e r and the r e s u l t s applied t o each structure The distribution of the substructure costs t o t h e concept, as applicable.

various zones i s based on the volume contained by each zone (i.e. the product of surface area and depth). These distribution factors a r e 44.5 percent, 31.1 percent, and 24.4 percent f o r zones A, B, and C, respectively. The structural panel costs are distributed on the basis of planform area with distribution factors of 33.4 percent, 27.9 percent, and 38.7 percent for zones A, B, and C, respectively. The heat shield costs are distributed i n proportion 23-2 to the area of applicability of the heat shields. The basic arrangement for the six structural concepts is presented in figure 23-4, showing ribs, spars, number of intersections, etc. The weights for each zone are based on unit weight results of the detailed structuzzal analysis presented in section 13 and swnmarjzed in table 23-4.

Substructure Costs. - The substructure costs consist of (1) chordwise ribs, (2) spanwise spars, (3) a leading edge spar, and (4) a breakline spar.

The assembly costs f'or the substructure are based on the number of spar-rib intersections in the main wing segment and a costing factor used to account f o r the type and complexity of the joint involved.

To determine substructure costs for all concepts, a detailed cost analysis of the monocoque waffle concept was made. These data were then applied as applicable ( i . e . , linear feet of spar, rib) to determine the appropriate cost f o r the fabrication and assembly of the substructure for the main wing manufac- turing segment.

For the substructure of the monocoque concepts, the chordwise ribs were considered continuous from Station 2136 to Station 2506 ( 3 0 . 8 f e e t ) . The rib assembly consisted of continuous caps of 30.8-foot length, with the web Each of the web elements was elements w i n g between the spanwise elements.

considered to be fabricated in a sequence of operations as shown in figure 23-5.

Fabrication of the caps was based on the assumption that the material was purchased as coil stock and slit to appropriate width. This stock would be The chordwise rib PabricatioI straightened, formed, and cut to a 30.8-foot length.

involved joining of the segmcated webs and continuous caps by melt-through welding the cap to the webs, using a tracer-controlled gantry-maunted welded The fixture for this operation was also used as the assembly fixture for head.

the weks and c a p as illustrated in figure 23-6. After the melt-through weld- ing operation, the overlapping edges of the webs are spotwelded for the depth of the beam.

Fabrication of the spars whs planned -to follow a procedure similar to that described above, except that the web had a sligbtly different configmation and the length of the spar segment., AS assembled, was a function of the spacing of the chordwise ribs (figure 2 3 - 7 ) .

The substructure assembly was fabricated by loading the chordwise ribs and spanwise spar segments into a horizontal fixture, and locating these elements at appropriate places to maintain contour and sparlrib spacing.

Figures 23-8 throu.gh 23-10 present typical intersections used for this study.

The various substructure elements were secured at the intersections by resist- ance welding supplemented in certain areas by mechanical fasteners. An esti- mated total of 14 200 resistance spotwelus and 1 015 mechanical fasteners were required in the study area. Appropriately designed splice plates weie added the upper and lower spar/rib at each intersection. An additional 6 800 to resistance welds were required to secure the splice plates. It vas assumed that the substructure would Se aged and oxidized as a unit prior to fit-up and assembly of the structural panels.

23 -3 The structural panel costs, including fabrication assembly and installation, were determined f o r each concept based on panel d e t a i l s presented i n the primary structure sizing and weights section (section 13). The manufacturing cost f o r

the monocoque waffle i s based on electrochemical milling (a), 'rstresskin"

and t h e manufacturing techniques discussed panels for t h e noneycomb sandwich, i n section 27 f o r the semimonocoque and s t a t i c a l l y determinate concepts.

Monocoque Concept Costs. - The wing substructure cost? f o r t h e minimum-

weight waffle concept (AR = 1.8) and the honeycomb concepx here obtained from t h e following data developed f o r t h e monocoque waPfle aspelat r a t i o study.

Aspect r a t i o study: Aspect r a t i o s of 1, 2, 3, and 4, as well a s 1.8 and ' 3.6, were investigated t o determine the s e n s i t i v i t y of t h i s parameter with respect t o weight and cost. A schematic for t h e various aspect r a t i o s studied i s shown i n figure 23-11. The substructure, panel, and t o t a l weight variation w i t h aspect r a t i o For a l l aspect r a t i o s evaluated, 5 shown i n figure 23-12.

a constant chordwise r i b spacing of 22.3 inches was assumed (b = 20.0 -t 2.3).

The aspect r a t i o of 1.8 and 3.6 minimizes the complexity a t the breakline spar intersection by providing repeatable panels and substmxture details.

The substructure costs were developed i n d e t a i l f o r t h e aspect r a t i o of These costs were then factored t o develop costs f o r each of the other 1.8.

aspect ratios. Substructure fabrication labor ar,d material costs were factored as a r a t i o of l i n e a r f e e t of s t r u c t u r a l elements t o the l i n e a r f e e t i n the 1.8 aspect r a t i o . Substructure assembly labor and material costs were factored as a r a t i o of the number of structure intersections. Tboling costs f o r the spanwise, diagonal (one-third high point), and leading edge beams were assumed t o be constant. Tooling f o r the chorcwise members was factored by the r a t i o of l i n e a r feet of structure, compensating f o r the impact t h a t the similarity of the r i b s within the fuselage area would have on t h i s tooling cost. Sbbs.tructure t o be constant, since t h e major part of assembly tooling costs were assumed t h i s cost r e s u l t s from the massive assembly f i x t u r e required- t o mate the various structure elements.

The monocoque waffle substructure manufacturing costs a r e presented i n table 23-5. These data show the increase i n t o t a l cost with the decrease i n aspect r a t i o due t o t h e increase i n number of spars, as well a s assembly com- Rlexity. Further comparison of substructure costs f o r aspect r a t i o 1.8 t o 2.0 indicates t h a t the increase i n spanwise beam cost exceeds tli.. cost due t o t h e thub t o t a l substructure complexity of the substructwe assembly, for AR = 2.0; costs for AR = 1.8 i s l percent t o 2 percent greater than f o r A 3 = 2.0. Sub- structure weight shows a similar trend with the lowest weight coming from t h e aspect r a t i o of 4.

The waffle panels were assumed t o be machined from p l a t e stock, u t i l i z i n g electrochemical milling (ECM) equipnent with a power of 20 000 amperes avail- able a t the cutting surface. A cutting r a t e of 0.1 in,/in.2/1000 amperes was used t o establish ECM machining costs, A study of a panel layout used f o r t h e 23 -4 0 0 45 pattern indicated that a minimum of f i v e ECM tools would be required by 45 f o r a regular sized panel. Special panels, such as occur along t h e leading edge beam, would require additional tooling. After ECM machining of t h e panel pockets, a secondary machining operation w a s performed t o remove t h e r i s e r s i n t h e flanged attaching areas.

After aging, panels were f i t t e d t o t h e sub- structure, trimmed t o size, drilled, and assembled, using Rene’41 plate nuts and flush screws on the lower surface w i t h H i - b k fasteners for the upper surface attachment. Table 23-6 presents the panel fabrication and i n s t a l l a t i o n with tooling costs amortized over 100 units.

costs, Wnimwn-cost r e s u l t s from the aspect r a t i o 1 panels, which also r e s u l t i n minimum panel weight. Com- parison of panel manufacturing costs f o r AR = 1.8 and AR = 2.0 indicates t h a t , although the panel fabrication cost i s less f o r the former, t h e instal- l a t i o n costs due t o the increased l i n e a r f e e t f o r attachments more than o f f s e t t h e gains for panel repeatability.

The t o t a l manufacturing cost variation with aspect r a t i o w a s obtained. by combining the information f o r the substructure (table 23-5) w i t h panel fabrica- t i o n and i n s t a l l a t i o n data (table 23-6) and heat shield data (table 23-2). The elemental costs f o r the substructure, panel, and heat shield/inzulation are presented f o r t h e various aspect r a t i o s i n t a b l e 23-7. The t o t a l cost varia- t i o n with aspect r a t i o indicates a decreasing cost trend for the greater aspect ratios. This difference, however, i s small, indicating minimum-weight con- siderations t o be more impartant than cost for the waffle concept.

A swmary of cost data f o r the waffle concept aspect r a t i o study i s pre- sented i n table 23-8. Labor, material, and nonrecurring costs are itemized separately t o show the effect of each on t o t a l cost. The data a r e presented i n dollars, dollars per square foot, and dollars per pound. For the minimum- weight arrangement (AR = 1 . 8 ) , labor costs account f o r approximately 31 percent of the t o t a l cost, w i t h material cost accounting f o r 65 percent. Tooling costs amortized over 100 u n i t s account f o r 4 percent of the t o t a l cost.

The effects of number of vehicles 0: these costs a r e presented i n figures 23-13, 23-14, 23-15, and 23-16. The decrease i n labor costs with increase i n aspect r a t i o i s indicated i n figure 23-13. That t h e material cost increases with aspect r a t i o is evident iii figure 23-14. The tooling cost variance with t h e number of vehicles is small for a l l numbers considered and becomes almost negligible with 100 o r more vehicles, as shown i n figure 23-15. Total manu- facturing cost variance with aspect r a t i o i s shown i n f2gure 23-16. When 100 or more vehicles a r e considered, t h e decrease i n labor cost w i t h aspect r a t i o i s o f f s e t by the increase i n material cost wlth aspect r a t i o , resulting i n approximately the same cost f o r a l l aspect ratios.

Cnncept costs: The substructure fabrication and labor cost for the minimum-weight waffle concept (AR = 1.8). which i s used a s t h e basis f o r deter- mining abstructure c o s t s for the other arrangeaents, i s presented i n table 23-9.

The t o t a l honeycomb concept substructure cost f o r the main wing manufacturing segment i s appmximately 60 percent of the waffle concept cost. This cost i s attributed t o a 50 percent reduction i n l i n e a r f e e t of r i b s and spars, coupled w i t h the reduced substructure amembly costs due t o t h e lower number of rib- spar intersections.

23 -5 The monocoque concept pinel fabrication and installation costs are presented i n table 23-10. The waffle concept costs a r e the r e s u l t s of the aspect r a t i o study (table 23-6). %e honeycomb concept costs r e f l e c t basic

il

panel costs (as purchased from Stresskin Products Co., Santa Ana, California) w i t h subsequent panel processing and i n s t a l l a t i o n accomplished i n a major fabrication and assembly area. This processing entails the machining of the core t o accept the inner closer channel, channel fabricakion, sptwelding, and chem-milling the face sheets t o the spscified thickness. Panel installatSon includes locating the panels i n the substructure, locating the cover stri s drilling, deburring, f i n a l trimming t o size, installating the plate nuts ?a: specified) and installing the flush fasteners. &ne1 fabrication cost reflec&s the major impact of labor f o r the honeycomb, whereas t h e material cost provi&d the major cost f o r the waffle panels.

Panel installation costs a r e a f'mctiog.

. I of the substrwture grid arrangement, accounting for the added complexity involved i n the honeycomb closeout design. Total panel costs indicate honey- comb costs t o be approximately 60 percent of the E C M waffle panels w i t h a weight reduction.

36 percent I.

The combined substructure, panel, and heat shield fabrication, assembly, and installation costs for the monocoque concepts are shown i n table 23-11.

L

This table summarizes the information on tables 23-9 and 23-10, i n addition to providing heat shield data, Total cost comparison indicates t h a t the honey- cor& concept i s approximately 62 percent of the waffle concept. The summary on table 23-12 presents the main wing manufacturing segment costs i n terms of lqibor, material, and tooling f o r the subs%ructure, panels, and heat shields.

For the waffle concept, labor accounts f o r approximately 31 percent and material approximately 64 percent of t h e t o t a l , with 5 percent for tooling.

For the honeycomb design, labor i s approximately 45 percent of the t o t a l cost, 48 percent for materials, and 7 percent for toollng. .For both concepts, the tooling cost i s insignificant.

I

To provide cost information f o r each zone (A, B, and C ) of t h e wing, appropriate distribution factors are applied t o the t o t a l costs previously

I mterial, and tooy-ing for each

calculated. Substructure costs f o r labor, zope are presented i n table 23-13e These costs with appropriate weights and aneas, a s indicated, provide unit costs for w x h zone. The average cost f o r i thte waffle concept is $90 per squeye foot, w. kh unit cosi; varience between $57 per square foot for t h e outboaru & + : a ta? $120 per square foot f o r the cepter area. Honeycomb concept uniz +OS!,G ($/f@) are approximately 63 percent of the waffle costs.

Panel fabrication and installation costs, including heat shield informatiun, i s provided i n table 23-14. The distribution factor i s a function of area; thus the resulting unit costs ($/ft2) are constant for each concept. Since and weight f o r the honeycomb concept &re approxitnately 64 percenk both the cost of the waffle concept, the resulting unit coats &re similar with the average being approximately $90 per p u n d .

23-6 A summary of the monocoque concept t o t a l manufacturing costs for each For the main wing segment, the honeycomb zone i s presented i n table 23-15.

concept cost i s .approximately $500 per square foot compared t u the waffle con- cegt cost of $779 per square foot. The cost difference i s attributed primarily t o material cost which i s directly associated w i t h might; thus, the importance of minimum weight i s emphasized.

Semimonocoque concept costs. - The wing substructure costs f o r the span-

wise-tubular, spanwise-beaded, and chordwise convex beaded/tubular concepts were obtained from the detailed costing information developed for the monocoque waffle eoncept reported earlier.

"he substructure fabrication labor and material costs were factored as a r a t i o of t h e linear feet of structural elements, as indicated i n figure 23-4 and table 23-16, t o t h e linear f e e t i n the AR = 1.8 monocoque arrangment (table Substructure assembly labor and material costs were factored as a ratio.

23-5.

of t h e number of structure irtersections (tablas 23-5 and 23-16) considering the complexity of the join'; involved (figure 23-4). Tooling costs f o r the spanwise, breakline, and lead5ng edge spars were assumed t o be constant. Tool- ing costs f o r the chordwis? ribs were factored by the ra-t;io of linear f e e t of stmcture, c0mpensatir.g f o : * %:le impact that the similarity of the spars within the fuselage area would have on t h i s tooling cost.

Substructure assembly tocling costs were assumed t o be constant, sinoe the major part of t h i s cost r e s u l t s from the massive assembly fixture required t o mate the various structural elements.

The lowest substruckwe cost is associated w i t h the spanwise concepts, the chordwise concept i s 26 percent costlier, due t o its closely spaced spars which r e s u l t i n a large number of spars and rib-spar joints.

The fabricated sheetmetal structural panels were costed i u d e t a i l with variations Xppropriate t o the uniqueness of each panel concept. For example, t h e tubular panel designs were assumed t o be formed i n two halves with each requiring a three-stage fcrming operation w i t h two interstage mieals. An- nealing was assumed t o be performed i n a controlled-atmosphere furnace with subsequent bath cooling. Panel halves, as formed, were assembled with blanked doublers and spotwelded t o form a complete structural panel. Heat shield com- ponents were added as appropriate, and the complete assembly aged and oxldized.

Panel assemblies were f i t t e d t o the substructure, trimmed, drilled, and as- sembled. Typical panels with appropriate heat shields were .osted for each zone (A,B, and C ) of the wing surface, considering changes i n mater5al usage and shape of panel. Other panel manufacturing was accomplished i n a similar

manner with appropriate variations f o r the particul.ar panel concept (i.e. ,, one

formied panel f o r the beaded concept, or specie1 forming tools along the leadiug edge of the panel for the chorawise concep'c). Although veccy high i n i t i a l tool- i n g costs are required for %he beaded panels, this panel r e s u l t s i n lowest Gost, with approximately 21 percent of the fabrication cost attributed t o labor, 6s percent f o r materials, and 15 percent for tooling. Since the material costs are related d i r e c t l y t o weight, the Lmportance of minimum weight is indicated by %hese data, The data also indica2;e at approximately 60 percexn't; of the panel i - 1

:l I

fabrication and i n s t a l l a t i o n costs a r e f o r installation. Comepix w i t h larger substructure grids r e s u l t i n fewer fasteners anc! minimize i n s t a l l a t i o n costs.

Cost comparison of the t o t a l panel fabrication and i n s t a l l a t i o n requirements indicates that the beaded concept i s lowest i n cost, with the tubular concept 2.5 percent greater and the chordwise concept 41 percent greater than t h e beaded concept. The l a t t e r i s attributed t o the impact of closely spaced spws requir-

i]

ing more closeouts per l i n e a r foot of panel, as well as greater insttillation costs due t o the increased l i n e a r f e e t of attachments.

A summary of the smimooocoque concept manufacturing costs =:..compassing qubstructure, panel, and heat shield (including insulation) fabrication, as- sembly, and i n s t a l l a t i o n i s presented i n table 23-17. The substructure and 23-18, respectively. The heat shiexd p n e l cost data a r e from tables 23-16 and (insulation) cost i s based on the data from t a b l e 23-2, w i t h heat shield being used on the exposed wing f o r the spanwise concepts and lower surface only f o r the chordwise concept. Minimum cost r e s u l t s f o r the spawise beaded concept, with the spanwise tubular b d n g approximately 2 percent c o s t l i e r and the chord- wise concept being 23 percent costlier.

A labor, material, and tooling cost swnmary is presented i n table 23-19 for

g

t h e semimonocoque concepts. For t h e lowest cost beaded concept, the labor costs account for approximately 4.0 percent of the t o t a l ; material costs 45 per- cent and amortized nonrecurring costs 15 percent. Total u n i t costs range from @gl/ft2 t o $358/ft2 and $58/lb bo $53/lb f w the minimum-weight t o maximum- considering the basic s t r u c t u r a l elements of the mair wing weight concepts, manufacturing segment.

I

The substmctw*e manufacturing costs f o r eac' of the zones (A,B, and Cj Appropriate distribution factors, weight, and a r e presented i n table 23-20.

geometry data a r e used t o develop unit costs for each zone as indicated, Basic

i

data from tak-e 23-17 and appropriate weight tables a r e used (tabla 23-4), The panel fabrication and i n s t a l l a t i o n costs f o r each zone a r e presented Cost data from t a b l e 23-18 a m used with the appropriate d i s - i n table 23-21.

trtbution factors noted and appropriate panel weight data irom t a b l e 23-4 t o obtain the unit cost data.

The heu: shield and insulation manufacturing cost data f o r each zone is Distribution factors are based on area of applicam prFsented iu table 23-22.

b i h i t y (9. e., zone A represents heat shield on the lower surface only) with weight information taken f r o m tables 23-4 t o obtain the unit cost data for t h e heltt shields.

A summary of' manufacturing costs and u n i t costs f o r each zone i s presented i n table 23-23. These manufacturing coste a r e based on data from tables 23-20 The costs r e f l e c t t h e manufacturing requirements for t h e basic 2+21 and 23-22.

Additional cost factors are essential t o develop mlt structural elements only, costs t h a t would be representative of the t o t a l wing.

23 -8

Statically detemninate concept costs. - The wing substructure manufactur-

ing costs, excluding the impact of the slip-joint at;seml~.ies, a r e presented 3.n table 23-24. The basic arrangement of substructure i s representad by t h e grid presented i n f i g w e 23-4 and includes additional spars as indicated. The d e t a i l i s used with appro- costing infomnatlon for the sxibstructure presented e a r l i e r priate linear feet and number of intersection r a t i o s t o obtain the cost data presented. Comparing the substructure cost data with the s9nwise beaded con- cept indicates a 16 percunt increase i n cost w i t h a L 2 percent increase i n sub- structure weight. The cost increase i s attributed t o an increase i n spar requirement and resulting increase i n t h e effective number of spar-rib inter..

sections. A f w t h e r iteration of the s t a t i c a l l y determinate concept could possibi,ly yield an increase i n spar spacing and a corresponding cost reduction, The impact of the slip-joint assemblies i s not included i n the aforementioned percentages.

The s t a t i c a l l y determinate concept panel manufacturing costs presented i n table 23-35 indicate similar fabrication costs and s l i g h t l y greater i n s t a l l a t i o n cost i n coiiiparison t o the semimonocoque spanwise beaded coacept. h c r e a s e i n installation costs fs the r e s u l t of panel attachment; details i n which additional fasteners are used along the panel spanwise joints, A summary of substructure, panel, and heat shield (including inr;uXation) costs for labor, material, and tooling i s presented i n table 23-26, Total man- ufacturing costs f o r the basic structural elements excluding the slip-joint assemblies are approximately 1 8 percent grecter than f o r the minimum-cost semi- monocoque concept. The impact of labor, malerials, and tooling on the manufac- Labor and material turing cost of t h i s concept i s presented i n table 23-27.

dollars each account f o r spproxima3ely 43 percexl: of the t o t a l , with heat shield cost approxiaately 13 percent of the t o t a l manufacturing coat. The u n i t cosL ($/ft2) of the basic structure (less slip-joint :!+sewblies) i e 18 percent greater than the minimum-cost semimonocoque concept x i t h a dollars-per-pound increase of 12 percent also indicated.

The manufacturing costs encompassing labor, materials, and tooling f o r each zone (A,B, and C ) are presented I n table 23-28*for the substructure, panel "he data w e based on r e s u l t s shown i n table 23-26, using and heat shields.

appropriate distribution factoxa discussed e a r l i e r and weight informtion from table 23-4. Table 23-29 sumrmarizes the labor, material, and tooling ?vAs f o r the nmin wing segment, as well as presenting the unit cost? ($/Lb and q/fk2) for each zone. These data, with appropriate cost factors t o account f o r the slip- j o i n t a s s a b l i e s and other cost iGems t o r e f l e c t t o t a l wing costs, are used a s inputs t o the Snteraction analysis discussed i n section 26.

?dins semen* --- cost nmar,y, - A summary of manufeacturing costs f o r the main

wing &e?~~ic.n.t is precentcd i n :abies 23-30 and 23-31. These costs are t o t a l costs for the combination OS concepts including primary stmctuxe d heat shields/insulation. Total cost, weight and wit cos$s ($/lb, $/ftT are pre- sented fcw each zone (table 23-31) as well as for the t o t a l mia %ring Sqgnent, The s t a t i c a l l y determinate costs do not include the impact of the slip- j o i n t assemblies; and encompass t h e basic elements of the wing structure only.

The minimuw cost concept is the semimonocoque-beaded at $291 per square

q

foot: with the semimonocoque-tubular next a t $5.00 per square foot greater.

i The monocoqueconcepts a r e the costliest, with the waffle md honeycomb being l . 6 8 percent and 65 percent greater, respectively, than t h e lowest-cost beaded concept. It is emphasized that these costs r e f l e c t manufacturing costs f o r only

B

tho basic structure of a representative manufactuTing s e e e n t (main wing) and only provide a cost comparison for a r e l a t i v e r a n k i w of the concepts. It is further noted that the s t a t i c a l l y determinate concept costs do not include the

I]

im’pact of the slip-joint assemblies. Factorn t o account for machined and sheet- metal parts, as well as other machined parts and miscellaneous structures are added 50 these costs t o provide cost data t o obtain t h e t o t a l wing manufacturfing costs discussed later.

Total Wing Costs Cost estimating relationships (CERs) for labor, material, and tooling w e r e ieveloped for each concept, using the detailed main wing segment inanufacturing costs as the bases. These Cms, presented i n $/U, provide a factor which, when nrdltiplied by the estimated weight of the t o t a l wing, res2XLts i n the incre- mental manufacturing cost (i. e., labor, material, tooling) of the t o t a l wing.

Wing geometrlr f o r the baseline vehicle, w i t h reference areas, is presented i n figure 23-17. Wing weights for the baseline vehicle a r e itemized i n table 23?32, and are used t o obtain the manufacturing casts f o r the t o t a l wing.

Experience has indicated t h a t f o r a given material and design concept selection, there is a preitic-kable relstionship between t o t a l vehicle manufactur- ing hours and hours requfred t o perform various a c t i v i t i e s during t h i s period of manufacturing. The main wing segment costing is a s detailed as possible considering the depth of des5gn available. The costing involves the fabrication and assembly of the basic structure ef the w i n g ( L e . , psnels, substructure, heat shield) which for most concepts are sheetmetal components, the exception being t h e ECM waffle concept p n e l s and the slfp-joint assemblies of t h e s t a t i c - a l l y determinate concept. The actual manufacturing of t h e complete wing .rrould introduce other sheetmetal and machined parts, particularly at t h e interfaces between manufactur9ng segments of *he vehicle t i . e. trzng-to-wing, wiiig-to- fuselage, etc.). Tihe developed CERs for the wing include factors which account for these unknown elements.

The detailed costing relat5onshLps developed for the supersonie transport

il

($T), which, from a technological standpoint was quite similar to t h e hypersonic cruise vehicle, indicated an estimsted cumulative average cost at 100 units of 768 100 t o t a l manufact;uring hou-rs. A breakdown of t h i s %otal Included 357 000 hours f o r machined parts, or 49 percent of the to.GaS. manufacturing lrours. For the purpose of t h i s study, it was aeemed t h a t t h i s relationship would exist f o r the semlmonocogue chordwise concept, since 5% is the most simikir t o %he con- structfon proposed for the SST (i, e. % muLtkspap, chordwise stizfened). Other as:;umptions made t o develop the CEBs f o r ea& concept include the following: To provide a more common basis of comparison of HCV (chordwise concept) t o SST, t h e heat shield/insulation labor and material costs were deleted from the chordwise concept costs t o arrive a t the machined parts cost.

Labor and material costs were increased by 25 percent t o account for additional sheetmetal parts ( 5 . e. stiffeners, clips, etc. ) but were not costed i n d e t a i l .

Calculated weights were increased by 10 percent to account f o r the uncosted items. These sheetmetal parts a r e assumed t o be relatively l i g h t i n weight.

A labor rate of $l2/hr was used t o obtain the time involved i n machining.

A material removal r a t e for Rene'4l of 0.266 lb/hr was used.

A n overall titanium machining material removal rate of 1.33 lb/hr has been developed from actual experience with ti.tanium. An analysis of the comparative mach2nabiltty of Rene' 41 versus 6m-4~ titanium f o r the various t y p s o f machining done during a i r c r a f t manufacture indicates t h a t Rene' 43- is approximately five times as d i f f i c u l t t o machine as titanium. Therefore, a material removal rate of (1.33/5) p o u n d s per hour was assumed f o r Rene'41.

A buy-to-net factor of 2 for sheetmetal costs was used to ac- ccunt f o r losses due t o rejected parts and other scrap.

A buy-to-net factor of 1 1 was used for machined parts.

Net material a f t e r machining was estimated as 10 percent of estimated material removed.

Total raw material purchased vas est5mated as equal t o the product of the net-to-buy factor f o r maciiined parts and. the estimated net materlal, o r 1.10 times the estlmated meherial rsmoved .

Machined parts raw material cost was based on $25/1b.

The machined parts estimated labor ana matex-ial costs a r e presented i n table 23-33. These cost esttmates a r e f o r the semimonscoque chordwise concept for labor and materials based on t h e asfiumptions made above. The total. costs for the machined parts are $262 6 1 . 4 . SLnce the machined parts required a r e primarily a f'unction of substmc.Gure arrangement and complexity, the machined 23 -11 i parts labor and material costs for t h e other concepts are assumed t o be pro- portional t o t h e substructure labor and material costs f o r each concept (table The s t a t i c a l l y determinate concept machined parts and added sheetinetal 23-34).

costs include t h e cost fox t h e special slip-joint f i t t i n g assemblies a t each spar-rib interface, special furselage-to-wing interface f i t t i n g s , and t h e re- quired s h e t m e t a l elements as calculated i n table 23-35. The labor and : %- tcrial costs f o r t h e slip-joint assemblies are $73/ft2 and $39/ft2, rep' tively, based on cost r e s u l t s of table 23-35.

h b o r and material cost estimating r e l a t i o n s h i j . -The t o t a l lab03 anL material costs are determined (table 23-34) f o r each concept by estimating cost increases (25 percent) due t o additional sheetmetal elements and appropriate factors for machined parts labor and material. U r i i t costs ($/lb) are calculated, &sing t h e costed structure weight increased by factors used t o account f o r t h e additional sheetmetal elements as w e l l as machined part weights. The f i n a l labor and material CERs presented i n table 23-34) are.used t o determine t h e cost for manwfacturing t h e complete w i n g . The CERs r e s u l t i n approximately twice t h e labor costs and 2.5 times t h e material cost developed for t h e main wing manufacturing segment.

Wing structure labor costs. -The t o t a l wing structure labor costs f o r The labor f a c t o r i s t h e each zone (A,B, and C ) are presented i n table 23-36.

r a t i o of t h e labor cost relationship determined for the overall wing t o t h e labor cost f o r t h e main wing manufacturing segment. The labor f a c t o r is used as a multiplying f a c t o r t o determine t o t a l labor costs f o r each zone of t h e wing.

Total labor costs f o r t h e wing concepts vary between 2.36 million dol- lars for t h e minimum-weight beaded concept t o 4.25 million dolIars f o r t h e monocoque waffle concept.

Wing structure material costs. - The total wing structure material costs

for each zone are presented i n t a b l e 23-37.

The material f a c t o r is t h e r a t i o of the material cost relationship determined f o r the overall wing t o -the ma- terial cost for each zone of the wing.

Total material costs f o r t h e wing concepts vary from 3 . 9 6 million dollars f o r t h e minimum-weight beaded concept t o 11.1 million dollars f o r t h e monocoque waffle concept.

Tooling cost estimatbg relationship. - The tooling cost estimates deter-

mined for the min ying manufacturmg segment are used t o calculate t h e tooling unit cost ($/lb) for t h e various structure concepts (table 23-38). Tooling costs for t h e estimated sheetmetal and wchtned parts are based on t h i s same unit cost; thus, t h e t o t a l tooling cost is increased a-pproximately 15 percent above i n i t i a l calculated values. The wing .Gooling CER for each concept (table 23-38) is used t o compute overall toolfng costx.

Tooling CERs f o r t h e fuse-ge and empennage are based on estimates avail- able from t h e SST program. Estimted tooling costs on t h e SST program were 85 hours per pound f o r the wing, 131 hours per ,omd for t h e fuselage, and 185 hours per pound f o r t h e empennage. As i n t h e case of machined parts cosls, these eetlmates are ~ ~ ~ 3 ~ m e d t o be directly applicable t o tho semf.monocoque chordwise concept. Appropriate values for t h e tooling CER for the fuselage f and empennage f o r t h i s concept, based on t h e estimated tooling cost for the SST, are shown i n t a b l e 23-39. This tooling C F R was assumed t o be constant f o r the fuselage and empennage f o r t h e various conceptsI The t o t a l structure tooling cost (table 23-39) f o r each s t r u c t u r a l seg- was determined by t h e product of t h e uent (wing, fuselage, and empennage) tioclling CER and the respective segment weight I The structure CER, as indicated, i s obtained by dividing t h e summation of tooling cost by t h e t o t a l weight.

I n addition t o t h e above tooling costs, f i n a l mate and assembly (FM & A) tooling costs m u s t be included i n the overall tooling costs. Experience on the P-3 program and the estimate f o r t h e SST program indicate t h a t about 19 per- cent of t h e t o t a l tooling costs r e s u l t i n t h i s area; thus, F M 80 A tooling costs are assumed t o be 19/81 or 23.4 percent of t'ne structure CER.

The tooling costs estimated t o this point represent what is usually refer- red t o as i n i t i a l tooling. '&is is the basic tooling required t o produce t h e vehicle prototypes. Once a production program i s begun, additional tooling (production tooling) i s required t o meet an established production rate. O n past programs, a r a t i o of t o t a l tooling ( i n i t i a l tooling plus production t o o l - ing) t o i n i t i a l tooli.ng has been estimated, using a Rand formula, a t about 2.2. Since it appears The SST program extimates indicated R m t i o of 2.82.

that the tooling required f o r t h i s program i s simpler, a m t i o of 2 has been assumed. The resulting data indicate t h a t the overall tooling cost estimating relationship (overall tooling CER) varies between $1416 per pound for t h e waffle concept t o approximately $2000 per pound f o r t h e s t a t i c a l l y monocoque det erminak e conc ept .

Total wing cost summary. -The manufacturing costs for t h e t o t a l wing struct%& f o r each concept were determined for the baseline airplane (GW = 550 000 l b ) , The t o t a l wing costs (table 23-40) are based on t h e cost estimat- ing relationships f o r labor, material, and tooling, and t h e t o t a l wing weights, as shown i n table 23-37, 23-38, and 23-39.

Results given i n Table 23-40 indicate t h a t t h e semimonocoque spanwise tubular is t h e next lowest-cost concept. This concept i s 6 percent heavier than t h e beaded concept f o r t h e baseline vehicle, but t h e wing coat i s only 2 percent greater than the beaded concept. The cost r e s u l t s f o r t h e other concepts, i n order of the cost, are semimonocoque chordwise; monocoque honey- comb s t a t i c a l l y 6.eterminate and monocoque waffle. The cost order is similar t o t h a t calculated f o r t h e main wing segment (table 23-30) except f o r change in order of the honeycomb and s t a t i c a l l y determinate concepts Vehicle Production Costs To determine vehicle production costs, a comparison of t h e overall wing structure cost estimating relationships with those developed f o r t h e SST pro- gram waij made. These data provzde a r a t i o indecating t h e r e l a t i v e complexity of the structural technologies between the SSI! and hypersonic cruise vehicle.

Using t h i s m t i o and value-engineering estimating techniques, cost est *%tin& relationships were developed f o r each of the strueturn1 and subsystem segments 23-13 f of t h e hypersonic cruise vehicle typified by figure 23-18. Development of

other subsystem requirements ( i .e , avionics, controls, etc ) were obtai'ied

from data taken from t h e Electra, P-3, F-104, FX, and SST programs- These relationships, as presented i n table 23-41, are i n terms of labor and material, and were developed from h i s t o r i c a l data accumulated from previous production and development c o n t r w t s .

For each of t h e s t r u c t u r a l concepts studied, it was assumed t h a t these cost estimating factors would remain constant for a l l segments other than the wing and leading edge table 23-42). It was assumed t h a t overall cost differ- enctes associated with the various s t r u c t u r a l concepts would be reflected by t h e application of these cost estimating factors t o varying vehicle segment weights.

Total vehicle production costs (labor and material only), less engines, were developed, u t i l i z i n g these cost estimating factors. Total vehicle costs shown in table 23-43 indicate t h a t t h e semimonocoque spnwise beaded concept i s minimum cost. The semimonocoque spanwise tubular concept is next, being less than 1 percent costlier. The other concepts, i n order of cost, are t h e seminonocoque chordwise, monocaque honeycomb statically determinate, and the lnonoc oque waffle.

The calculated dollars per pound ($/16) main wing manufacturing c o s t s inSormtion f o r each zone, as discussed i n t h e concept cost sections and sum- rnaraized i n tables 23-15, 23-23, and 23-29, a r e used with t h e appropriate cost faators as inputs t o t h e interaction analysis discussed i n section 26.

23 -14 TABLE 23-1 I N I T I A L SCItEHNING COS'I'S O F STRUCTURAL PANELS Primary structure concepts Material cost, Units $/ ft2 I a

M o n o coque Waffle grid unflanged - Rend 4 1

J+5O x 4 5 O Hwnes 25

Waffle grid flanged - Itend 4 1

450 x 450 Haynes 25

Waffle grid unflanged - Rend 4 1

0 ' x 90° Haynes 25

Waffle grfd flanged - Rend 4 1

29 3 0 ' x.9O0 Haynes 25 Yoneycomb sandwich Rend 4 1 35 4 Haynes 25 Truss-core saadwich Rend 4 1 Haynes 25 Semimonocoque b Tubular Rend 41 (spanwise) Haynes 25 Corrugation stiffened Iiend 4 1 74 Haynes 25 Trapezoidal corrugation Rend 4 1 Haynes 25 Beaded lime' 4 1 Naynes 25 b Convex beaded S emimonocoque ]{end 4 1 ( chor dwi s e ) Haynes 25 TrapezoidaL corrugation Rend 41 Haynes 25 4-7 Beaded Rend 4 1 Haynes 25 a = 9,s fez; Panel size: 26 in, x 49.8 i n .

'ref 9 . 0 ft2; ~ a n e ~ size: 30 i n , x 43.2 i n , %e f - - a * 0 3 , qc TABLE 23-2 SUMMARY OF HEAT SHIELD COST EVALUATION FACTOR DATA FOR 100-VEHIC I ; E PRODUCTION RUN cost Permanently Refur bishable attached evaluation factor Corrugated skin Flat-skin dimple- Simply Canti- multiple supports stiffened clip-supported supported levered Material and labor, 24.50 40.10 44.20 36.40 $ per ft2 TABLE 23-3 LEADING-EDGE COST EVALUATION FOR WING EVALUATION AREA WITH 100-VEHIC LE PRODUCTION RUN

Dollar s/ll Dollars/lbear ft 1

Leading-Edge Primary structure Concept b

Total Labor lMateriall Total I

Labor Material

I

Monocoque 22.90 I 77.55 100.45 m 97.40 329.60 427.00

.

Segmenteda Semimonoeoque and Statically deter- 19.90 67.40 97.40 329.60 427.00 87.30 minate Monocoque 387.01 501.34 888.35 51.33 199.13 250.46 Continuous

=I=

Semimonocoque and 497.68 892.47 227.72 Statically deter- 394.79 47.50 180.22 minate I I a2~-in. segments.

bTD Ni Cr .

TABLE 23-4 WEIGHTS FOR MAIN WING MANUFACTURING SEGMENT ~ Ib W, lb/ft2 p, Primary structure concept

- -

B C C A A B

- -

Monocoque waffle concept Panel 6 887 2269 2 294 6.68 2 324 70 81 9.13 Substructure 2.90 2.90 2 493 821 690 1. 69 976 0 Thermal protection 3804 380 1.07 0 0 0 Heat shield (0.92) (330) Insulation (0.15) (50)

-

9 760 3 090 3 370 Total 9.44 3 300 0.71 12003 I

-

Monocoque honeycomb rmcept 4 400‘ 1240 1700 4.76 Panel 4.73 4.83 119oi 330 0.92 Substructure 1.52 1. 52 390 470 380 i 0 380 1s 07 Thermal protection 0 0 0 I (0.92) Heat shield I (3301 Insulation (0. 15)

- -

5 970 ’ 1630 2410 Total 6.25 6.35

6.75 1930 , Seinimonocoque, spanwise tubular 2 631 748 992 2.78 891 Panel 2.89 2.91 1212 404 386 1.08 Substructure 1.37 1.57 1 115 290 653 1.83 172 Thermal protection 0.56 1.13 (1.53) Heat shield [ O b 56) (1.13) (5461 (0 30) Insulation (1071

-

-

4 985 1442 2 031 4.82 5.69 1 4 8 5 .Total 5.61 I

- -

Serpimonaoaque, spanwise be@ed 2 343 878 689 2.46 776 b-1 2.52 2.68 1212 404 386 1.08 422 Substruoture le57 1.37 1 116 290 dB3 1 . 8 3 172 Oe56 1.13 Thermal protection (546: (1.53) ‘ Heatshield (0.66) (la 13) (107: (0.30) fnsulation

- -

4 670 1 3 8 3 1917 1 370 Total

- -

2 2 2 “ S i = 308 ft ; SB . = 257 ft ; Sc = 357 f t ; STOM = 922 ft2..

TABLE 23-4 MAIN WING SEGMENT WEIGHTS (CONCLUDED)

I W, lb/ft2 I w, lb

Primary structure concept C A B C 2

IA(B

Semimonocoque chordwise convex beaded tubular Panel 3.52 3.47 3.39 1 0 7 0 370 1 2 6 0 3 200 Substructure 1.54 880 720 550 I 2 150 2.85 2.78 Thermal protection 0.61 0.61 Heat shield Insulation 0 Total atattically determinate spanwise beaded Panel Substructure Thermal protection 0.56 Heat shield Insulation Total 23-29 Ml-tOoO mcumao rl ' 3 M c u a m rlrld VJ e3 cu rl

I " 9% ---E

4 u : ed - i r R J p' u

f

23 -22 m u I b I M cu

w

I 4 E9 !3 23-23 Y WF' I I I

r- --

t

P I m m a m v3 m

c

23-24 m m TABLE 23-9 MONOCOQUE CONCEPT SUBSTRUCTURE MANUFACTURING COSTa (DOLLARS) (Main Wing Segment) Monocoque Structure concept Waffle Honeycomb 1. Chordwise ribs Linear feet ft 5 02 263 3 930 2 060 Labor $ 11 160 21 300 Material $ Nonrecurring $ 1 290 650 835 050 38 1.36 21 570 Subtotalb $ 2 . Spanwise beams Linear feet f t 273 142 3 220 Labor $ 12 120 6 300 Material $ Nonrecurring $ 313 080 313 080 18 471 11 101 Subtotalb $ 3. Leading edge and breakline beams Linear feet ft 68 380 380 Labor $ Material $ 156 190 156 190 Nonrec rring $ 3 792 3 792

Subtotal E $

4 . Substructure assembly No. of intersections 6 13 6 950 2 410 Labor $ 3 520 Material $ 1217 450 Nonrecurring $ 1 217 450 22 644 15 504 Subtotalb $ 83 043 52 267 5 . Total Cost $ 6. Substructure 2 493 1 190 Weight lb 23-25 TABLE 23-10 MONOCOQUE CONCEPT PANEL FABRICATION AND INSTALLATION COSTSa (DOLLARS) (Main Wing Segment) Monocoaue Structure concept Waffle Honeycomb 1 . Fabrication Labor 65 300 83 130 Material 358 500 151 994'

--

Nonrecurring 186 500 Subtotalb 425 665 235 124 2 . Installation Linear feet 843 473 Labor 139 150 105 990 Material 54 860 32 190 Nonrecurring 53 720 34 033 b 138 520 Subtotal 194 547 3. Total panel cost $ 620 212 373 664 4. Panel weight lb 6 884 4 400 aFor one-half of the main wing segment = 922 ft2 bNonrecurring costs amortized over 100 units.

C lncludes basic honeycomb panel purchased from Stresskin Products Co. , Santa Ana, Calif.

23 -26 TABLE 23-11 MONOCOQUE CONCEPT SUBSTRUCTURE, PANEL, HEAT SHIELD (INCLUDING INSULATION) FABRICATION, ASSEMBLY AND INSTALLATION COSTSa (DOLLARS) (Main Wing Segment) Monocoque Structure concept Waffle Honeycomb 1 . Substructure $ 6 520 14 480 Labor $ 20 530 Material $ 38 790 Amort. NR $ 29 773 25 217 83 043 52 267 Subtotal $ 2. Panel $ 204 450 189 120 Labor $ 413 360 184 184 Material $ Amort. NR $ 2 402 340 373 644 620 212 Subtotal $ 3. Heat shield/insul. $ 3 100 3 100 Lab0 - $ 9 250 9 250 Material $ Amort. NR $ 2 950 2 950 15 300 15 300 Subtotal $ 718 555 441 211 4 . Total cost $ 5 970 5 . Total weight lb 9 760 a For one-half the main wing segment.

23 -27 TABLE 23-12 MONOCOQUECONCEPT MAIN WING SEGMENT COST Monocoque Structure concept Waffle Honeycomb Planform area ft2 922 922 Total weight lb 9 760 5 970 222 030 198 740 Labor $

$/e2 241 216

$Ab 23 33 461 400 213 964 Material $

$/e2 500 232

47 36 $/lb 35 125 28 507 Nonrecurring $

(Amort over $/a2 38

100 units) 4 5 $/lb

I $ 718 555 441 211

Total cost 1 $/fP 480

73 74

I $/lb

23 -28 I I M c- c- s i P- c - c u m r l r l M 0 0 0 o m 3 6 3 m o m O J c u d

'

23 -29

I t - -

0 0 0 0 0 0 0 m o o c o d 0 * c u b ttc I I c r , c*J

r l r l d *

cu O b 0 23-30 23 -31 TABLE 23-16 SE MIMONOCOQUE SUBSTRUCTURE MANUF.ACT UmNG C OSTSa (DQLIARS) (Main Wing Segment) Spanwise Chordwise $tructwrz Tubular C Beaded Convex beaded (U) concept Tubular (U) 1. Chordwise ribs ]Linear feet ft 259 259 146 2 030 2 030 1 144 Labor $ 10 990 10 990 Material $ 6 198 Nonrecurring $ 9Q2 420 902 420 653 070 22 044 22 0 4 13 873 Subtotdb $ 2 . Spanwise beams Linear feet ft 119 119 434 1 400 1400 5 120 Labor $ 5 280 5 280 19 270 Material $ Nonrecurring $ 313 080 313 080 313 OS0 Subtotalb 9 9 811 9 811 27 5 2 ' 3 . Leading edge and breakline beams Linear feet ft 68 68 68 380 380 38 0 Labor $ 1850 I850 Material $ Nonrecurring $ 156 190 156 190 156 190 3 792 3 792 3 792 Sllbtotalb $ 4 . Substructure assembly No. of intersections 186 186 400 4 532 2 106 2 106 Labor $ 2 295 1067 1 067 Material $ 1217 450 Nonrec ring $ 1217 450 1217 450 15 347 15 347 19 000 Subtotal $ $ b 50 994 64 186

I 5 . Total cost 50 994

2 X5G 3 212 1 212 6 . Substructure lb weight "For one-half of main wing area = 922 f t .

bNonrecurring costs amortized over 100 units.

23 -32 TABLE 23-17 SEMIMONWOQUE SUBSTRUCTURE, PANEL, HEAT SHIELD (INCL INSULATION) FABRICATION, ASSEMBLY AND INSTALLATION COSTSa (DOLLARS) (Main Wing Segment) Spanwise C hordwise Structure

-

concept Tubular Beaded Convex Beaded (U), Tubular lL\ 1. Substructure 11 175 5 916 5 916 Labor $ Material 19 187 19 157 29 613 b $ 23 398 Amort. NR $ 25 891 25 891 64 186 50 994 50 994 Subtotalb $ 2. Panel 84 598 126 a30 90 151 Labor $ 69 252 1 100 680 75 199 Material $ Amort. NR $ 3 362 lQ 576 3 230 168 712 164 426 2 m 740 subtotalb $ 3. Heat s'nield/Insulag Area ft 1536 1535 922 16 375 16 375 10 033 Labor $ 32 514 32 514 20 440 Material $ Amort. NR $ 4 458 4 455 4 455 53 344 53 344 34 936 Subtotalb $ PI. u_ 329 862 273 050 268 764 4. Total cost; $ "For one-half or" main wing area = 923 ft2.

bNR. nonrecurring costs amortized over 100 units.

23 -33 TABLE 23-18 SEMIMCNOCOQUE PANEL FABFUCATION AND INSTALLATION COST* (DOLLARS) (Main Wing Segment) Spanwise Chordwise Structure concept Tubular Beaded Convex beaded (U) Tubular (L) 1. Fabrication 16 531 $ 13 448 19 860 Labor 45 939 41 202 58 510 Material $ 301 900 1 024 300 274 900 Nonrecurr%g $ 65 489 64 893 81 119 SI& .otalb 8 2 . Installation Linear feet ft 4 4 6 431 648 7s 620 71 150 106 970 Labor $ 28 050 42 170 29 260 Material. $ 34 260 33 300 48 140 Nonrecurring $ 103 223 99 533 149 621 subtotalb $

-. . - -

164 426 230 740 3 . Total panel cwcs $ 168 712 2 631 2 343 3 200

I

aFor one-hdf o f main wing area = 922 ft .

bNonrecurring costs amortized over 100 units.

23-34 SEMIMONOCOQUE CONCEPTS MAIN WING SEGMENT COSTS Spanwise Chordwise Structure concept Convex beaded (U) Tubular Beaded Tubular (L) area ft2 922 922 922 Planform 4 958 4 670 6 150 Total weight lb Labor $ 112 4L.2 106 889 148 038 122 116 160 $/f?

23 23 24 $/lb Material $ 126 900 120 953 150 741 $/ft2 138 131 164 26 26 24 $/lb Nonrecurring $ 33 708 40 922 31 083 (Amort. over 36 44 34 $/ft2 100 units) 7 9 5 $/b 273 050 268 764 329 862 $ Total $/ft2 29 1 358 cost 55 58 53 $/lb 23-35

* %

I n r t c u o o 4 c u m r t m m o m 3 P) N u a Q)

u

23 -36 m m m G o o 0 0 0 d r ( d m m C D C D W W G o c - e - 0 m L3 m c v m w o m l - i o o o r l m r l m o m c \ l c i l c u

4 a u

23 -37 m o o 0 m o o c r 3 b b m m b b d a m 00 d a m

w m c 5 dc w m m -/aPPI$

m o m m o m w c - 0

z & 22131 w m m

d d qqo 0 d

b b u 3 U 5 M O 0 4 0 0 a Q, w o f - a 0

*

cu

*

dl G J 23 -39 TABLE 23-24 STATICALLY DETERMINATE SUBSTRUCTURE MANUFACTURING COSTSa (Main Wing Segment) Spanwise Structure concept Beaded 1. Chordwise ribs Linear feet f t 227 1 774 Iabor $ 9 620 Material $ 946 180 Nonrecurring $ 20 856 Subtotalb $ 2. Spanwise beams Linear feet f t 274 3 240 Labor $ 12 200 Material $ 313 800 Nonrecurring $ 18 578 Sllbtotalb $ 3. Leading edge ani breakline beams Linear feet ft Irtbor $ 1 850 Material $ 156 190 Nonrecurring $ 3 792 Subtotalb $ 4. Substracture assembly No. of intersections 22 1 2 510 Labor $ 1 270 Material $ 1 217 450 Nonrecurring $ Subtotalb 15 954 5. Totalcost 59 180

~,. Substructure weight Ib 1 360

aFor one-half of main wing segment = 922 ft2 , bNonrecurring costs amortized over 100 units.

23 -40 TABLE 23-25 AND STATICALLY DETERMINATE PANEL FABRICATION INSTALLATION COSTW (DOLLARS) (Main Wing Segment)

-

Structure concept Beaded 1 . Fabrication Labor 13 784 Material 43 500 Nonrecurring 1121 852 Subtotalb 68 502 2. Installation Linear feet 599 98 900 Labor Material 39 000 43 300 Nonrecurring 138 333 subtotalb 206 835 3. Total panel costs $ 4. Panel weights lb 2 a For one-half of main wing segment = 922 ft .

bNonrecurring costs amortized over 100 units.

23 -41 TABLE 23-26 STATICALLY DETERMIN.ATE CONCEPT TOTAL SUBSTRUCTURE7PANEL, HEAT-SHIE LD FABRICATION, ASSEMBLY AND INSTALLATION COSTSa (DOLLARS) (Main Wing Segment) Spanwise Structure concept Beaded 1. Substructure Labor 7 904 24 940 Materia 1 .

b Am- :t. NR 26 336 Subtotal 59 180 2. Panel 112 ('84 Labor Mate rial 82 500 b Amort. NR 11 651 Subtotal 206 835 3. Heat Shields Labor 15 865 Material 30 184 b Amort. NR 4 455 30 504 Subtotal 316 519 4. Totalcost -u a one-half of main wing area = 922 ft2 .

bNonrecurring costs amortized over 100 units.

23-42 TABLE 23-27 SUMMARY STATICALLY DETERMINATE CONCEPT MAIN WING SEGMENT COSTS Spanwi se Structure concept Beaded

-

Planform area ft2 Total weight Ib 4 843 136 453 Labor $ $/ft2 $ /Ib 137 624 Material $ $/ft2 14 9 $/lb Nonrecurring $ 42 442 (Amort. over $/ft2 100 units $/lb 316 519 Total cost

I $/lb

24-43 -1 I I I I I I I

23 -44

r C - J C r J L c o r n o L o c h 0 r?, -45 23 -46 TABLE 23-31

SUNMARY - STRUCTURE CONCEPTa MANUFACTURING COST§

(Main Wing Maiiufacturing Segment)

- I_

$ I $2 Ib I ft I Monocoque 74 793 I I 82 772

-

73 l 779

II -

77 1 481 74 469 72 483

- -

74 480 II 60 291 3 1 4 tubular 50 287 UI II_ 64 285 58 3 10 53 282 2 92 _I _I_ ~ 55 382 54 366 convexbeaded/ tubular 52 332

- L__

54 358

-

I 70 64 362 62 330 _I_ _u_1 66 344 -L __u aPrirnary structure and heat shield,/insulation.

bDoes not include the impact of the slip-joint assemblies.

TABLE 23-32 TOTAL WING WEIGHT SUMMARYa

(Baseline - G. W. = 550,000 lb)

a Pounds.

bDoes not include leading edge.

23 -48 TABLE 23-33 MACHINE PARTS ESTIMATED LABOR AND MATEIUAL COSTS FOR OVERALL WING STRUCTUREa Results Item Operatiodref.

1 Total labor costs - main wing (Table 23-19) $148 038 2 Heat shield/insulation labor costs (Table 23-17) 10 033

3 Labor costs - basic structure 138 005

(a- E)

172 506

4 Estimate total labor costs (1.25 1)

Total material costs - maiu wing) (Table 23-19 150 741

20 448 6 Heat shield/insulation mate rial costs (Table 23-17 7 Material costs -basic structure 130 293

(a-

8 Estimated net material (1.25 m) 162 870

- 9 Buy-to-net factor 10 Total estimated material 325 740 (rnX191)

11 Total labor and material - basic structure $498 246

(B+ flol)

12 Total machined parts labor $165 726

(49/51 m)

13 13 810 h r Total machined parts hours Rate of $12/hr 14 Estimated material removed Rate of 0.266 l b l a 3 673 lb 15 Buy-ta-net factor 11 367 lb 16 Estimated net material Estimated total raw material purchased 4 037 lb 18 Machined parts raw material costs $24/lb 19 Machined parts estimated material costs $ 96 888

20 Total labor and material - machined parts $262 614'

"Estimated cost for semimonocoque chordwise concept based on assumptions specified.

23-49 m o m 0 0 G o O N O O O D m m b c u

' I '

0 0 0 N 0 0 m OD Q) 0 0 In

* 0 N u)

N b Q) CD N 00 0

* 3

u) N c - (0 co P- d cu 0 d I4 Go

3 co CJ (D

rl 4

rl m m I I n 23-51 c, m c1 cu I M cu c, a 'CI n 1 1 1 1

El

Y - . - .

m cd 23-52 I I I I 8 . I

I I

- 1

23 -53 U J W N 23 -56 TABLE 23-41 LWIICLE COST ESTIMATING FACTORS Units Value Costing factors cost per pound of avionics CPAV = 1590 C K S L = labcr cost fer ECS 30 CEC34 = rnaterial cost f o r ECS c;:'L = labor cost f o r elevans cl!N = material cost for elevons labor cost f u r e l e c t r i c a l C m R L = 89 C'ELHM = material ccst f o r e l e c t r i c a l labor cost f o r furnishings and equipment C F E U = C F W = material cost f o r furnishings and equipment labor cost f o r f l i g h t controls CFCL = 75 CFrn = material cost f o r f l i g h t controls labor cost f o r f i n s CFIIG = 153 CFIYlM = material cost f o r f i n s CFSL = labor cost f o r f u e l system 151 CFSM = material cost for f u e l system 289 CFUSL = labor cost f o r body structure CFVSM = material cosi f o r body structure labor cost for hydraulic 120 CHYDL = material cost f o r hydraulic CHYDM = 342 labor cost f o r i n l e t CLNLL = 219 CINLM = material cost f o r i n l e t TABLE 23-41 (Concluded) Ccsting factors Value CiTi'L = labor cost f o r instruments C I f F i Y = material ccst fcx instruments 186

*

CLKL = labor cost fer wing leading edges

*

cLF2.I = material ccst fcr wing leading edges

*

r ; W L , A = labor ccst Z c r wing structures - A

*

CI:GB = lahor c c s t f o r wing structures - B

*

c A n r T w L C = labor cost f a r wing structures - C

*

cbWfA = material cost for wing structures - A

*

CMNME3 = material cost for wing structures - B

*

C M W C = material cost for wing struct:lres - C C T L = labor cost for nose cap CRCM = material cost f o r nose cap *?

WIG = labor cos% f o r landing gear L G J i G i = material cost for landing @til* 29

1-511 .o

IC.?AV = i n s t a l l a t i o n cost per powid o f avionlrl.:: 'L.0 NTRJ number of engines per vehLcle --.

*Reference t a b l e 23 -42.

23 -59 O b *a 23 -60 TABLE 23-43 TOTAL VEHICLE PRODUCTION COSTSa (100 VEHICLES) ~ Primary structure concept Dollars ($) ~~ - Monocoque waffle 51.745 x IO6 46.273 Monocoque honeycomb Semimonocoque spanwise tubular 44.255 44.032 Semimonocoque spanwise beaded Semimonocoque chordwise convex 45.814 beaded, upper; tubular, lower 46.835 Statically determinate spanwise beaded "Labor and material, less engines.

23 -61 \ 1 8 x lo2

-

-.

,-Modular, simply supported I 12 Flat-skin dimple-stiffened clip-supported

k! 4-'

I .

r/

___)____

I , Modular contilevered L Corrugated skin multiple supports I

4 -

2 00 300 400 500 10 100 Numkr of vehiclor F i w e 23-1. Unit heat shield cost versus number of aircraft 2 3 -62

V

\

Q I S TA S T A 2136 2506 L eading-edge spar Chordwire rib Area (ft2) Zone A B C

- 743 includes leacling

edge (Sle = 30 ft ) Total Figure 23-3. Manufacturing segment, main wing 23 -64

I

A ?

4-1/4 in. tooling hcles 1st ' - - !

q & . - .. 47.0 Net length

operation . I I block 2nd operationlblock I- - - ' Amount of material Develop blank, 8.02 in. x 44.4 x 49 inches x 0.298 Ib/in.3 = 12.97 Ib Net part 0.02 in. x 2087 inches x 0.298 Ib/ina3 = 12.44 Ib .* ~ .. . - - .- -- .- - .~ - . _ _ - .

Material cost 2176 inches'x O.O8f'3* $/in,2 = $175.82 - - . - - ._ . - I ____ ..- ._. . ..

-. ..

Production time ECH Ref. Operation Set-up time Run time 10.220 Shear strip, blank and pierce, burr, clean, and ID 3.43 0.348 10.220 Verson {arm 1st block 0.05 0.085 10.220 Verson form 2nd block 0.05 0.085 1.136 Shear ends to net trim 0.13 0.145 (remove added trim and T,H.'s)

--

Total (hours) 3.66 0.663 ~ _ .

(hours) - - ..-- . . - - ~ _I__--

*

Ref: Engineering Cost Handbook (ECH) - - -. . - - .-. _. -. ., . .. .___ --.- - .-. -.

- .- -. __ - .

- - - Figure 23-5. Circular-arc Corrugation Web element Fabrication 23 -66 a , f S

.-

P t

. 1 --

Q D V W fll cn

.-

e

M cu Qutb'd I I Fwd

x

-\ -.

\ \ \

'\

1 ' $ 1

/

5- Spotweld 24 places

i

QQ / //' ' I

!

j x f

I

I f

i I #lo nut and bolt (bottom) L e . . I.- -.__ _I \ #IO Hi-Loks (top)

'\ (gf" (structural pane I fasteners)

I

\

I

'.

I \ \ \ \

I

--dL---- l/-"-Ll,.- Figure 23-8. Typical intersection detail 23 -69 1 .34 Linear dim ./node 44.22 Linear node length +I .OO Flat 1st end x 33 Nodes/panel

-

+1.78 Flat 2nd end 44.22 Linear length/33 nodes 47.00 Inches = Total linear length . - - - _ - _ _ 120° .. .

_ - -.

---, 1 . . , Web circular ,

‘, Arc corrugation -. i

!’:/ -’-

Arc

Arc length = 2*R* 3600. i

120 1

= 2(3.1416)(0.75) 360

I = 1.52 inches -.- . - F i g u r e 23-9. Typical web and beam cap Intersection d e t a i l 23 -70 23 - 7 1 .

23 -72 ' 3 Aspect ratio, a/b Figure 23-12. Main wing segment weight variation with aspect ratio 23-73 , 6 x 10

1 I

nocoque waffle conoept

?I 3

a $ 3 4J rn k Aspect ratio Labor c o s t vs aspect ratio Figure 23-13.

23 -74 . .

x 10 500 Aircraft ~ Monocoque waffle concept 1 2 3 4 Aspect r a t i o Figure 23-14. Material costs vs aspect r a t i o ?

23-75

6 x lo5

Monocoque waffle concept LO Alrcraft I 1 2 Aspect ratio F i g w e 23-15. Tooling costs VB asgect ratio c m 8 10 .- L c U . . .

t

E"

Figure 23-16. Total manufacturing cost vs aspect r a t i o .

23-78

I

23 -79

Section 24 P E R F O ~ C E PSITALYSIS bY

R . S . Pe-&on

CONTENTS METHODS AND PARAMETRIC DESIGN D A T A Uniformly Distributed o r Equivalent Sand Grain Roughness Sheet Metal J o i n t s and Fasteners Two-Dimensional Surface Waviness i n Which t h e Wave Crests 24 -2 are Perpendicular t o t h e Wing Chord Three-Dimensional Surface Bumps or Depressions 24 -2 Surface Corrugations h r a l l e l t o t h e Wing Chord 24 -3 D e f o r m t ion of t h e Priurary Wing Structure 24 -3 24-4 Parametric Design Data Evaluation Approach 24 -4 24 -4 P R I M A R Y STRUCTURES 24-5 Monocoque Waffle Monocoque Honeycomb Sandwich 24-5 24-5 Semimonocoque Spanwise Tubular 24 -5 Seminionocdque Spanwise Beaded Skin 24 -6 S emimonocoque Chordwise Convex-Beaded/Tubular 24 -6 S t a t i c a l l y Determinate 24 -6 Fuel Increment S u m m r y 24 -7 HEAT SHIELlDS 24 -8 LEADING EDGE 24-9 F@FWENCES 24-iii PRECEDING PACE BLANK NOT FILMED.

i 24-10 24-1 Primary-structure coDcept performance evaluation 24 -11 24-2 Eea-t -shield concept performance evaluation PRECEDING PAGE BLANK NOT FILMED.

ILLUSTRATIONS Figure 24 -I Fuel increment required t o compensate f o r uniformly distributed roughness on wing surface f o r constant missior.

range 24-12 24 -2 Fuel increment required t o compensate f o r uniform waviness over wing surface f o r constant mission range 24-13 24-3 Fuel increment required t o compensate f o r uniform three- dimensional waviness over wing surface f o r constant range 24-14 24-4 Fuel increment required t o compensate f o r uniform corruga- t i o n i n wing surface f o r constant mission range 24 -14 24 -5 Fuselage deflections net due t o l i m i t loads along BL 120 (intersection of fuselage and wing), monocoque waffle concept 24-16 24 -6 Wing deflections net due t o l i m i t loads, monocoque waffle 24 -17 concept 24 -7 Fuselage deflection due t o thermal stresses along BL 120 (intersection of fuselage and wing), monocoque waffle 24-18 c oncept 24 -8 Wing deflections due t o thermal stresses, monocoque waffle 24-18 concept 24 -9 Fuselage deflection net due t o l i m i t loads along BL 120 24-19 (intersection of fuselage and wing) honeycomb concept 24 -10 Wing deflections net due t o l i m i t loads, honeycomb 24 -20 concept 24-11 Fuselage deflections due t o thermal stresses along BL 120 24 -21 (intersection of fuselage and wing), honeycomb concept 24-12 Wing deflections due t o thermal stresses, honeycomb 24 -22 c oncept 24-vii Figure 24-13 Fuselage deflections net due t o l i m i t loads along BL 120 (intersection of fuselage and wing), semimonocoque (spanwise) concept 24 -23 24-14.

Wing deflections net due t o l i m i t loads, semimonocoque (spanwise) concept 24 -24 24 -1 5 Fuselage deflections due t o thermal stresses along BL 120 (intersection of fuselage and wing), semimonocoque (spanwise) concept 24-25 24 -16 Wing deflections due t o therm1 stresses 24-25 25 -17 Fuselage deflection net due t o l i m i t loads along BL 120 (intersection of fuselage and wing) semimonocoque ( chordw is e) c oncept 24 -26 "4 -18 Wing deflections net due t o l i m i t loads, seroimonocoque (chordwise) concept 24-27 24 -19 Fuselage deflection net due t o l i m i t loads along BL 120 (intersection of fuselage and end wing) semimonocoque (chordwise) concept 24-28 24 -x) Wing deflections due t o thermal stresses semimonocoque (chordwipc) concept 24-29 24 -2 1 Fuselage deflections net due t o l i m i t loads along BL 120 (intersection of fuselage and wing), s t a t i c a l l y determinate cone ept 24 -30 24 -22 Wing deflections net due t o l i m i t loads, s t a t i c a l l y determinate cone ept 214-31 24 -23 Fuselage deflections due t o thermal stresses along BL 120 (intersection of fuselage and wing), s t a t i c a l l y c ' rminate cone ept 24 -32 24 -24 Wing deflections due t o thermal stresses, s t a t i c a l l y determinate concept 24 -32 24-viii BL Butt l i n e Fs Fuselage s t a t i o n g Gravitational acceleration L Distance between end closeouts Paximum height o r depth of surface wave F h Wave length of surface perturbation CL Wave length of surface perturbation n o m 1 t o a i d l o w I 24 -ix

Section 24 PEIIFORMANCE ANALYSIS ?me performance analyses consisted of evaluating t h e primary structures, heat shields, and leading edges f o r performance degradation (aerodynamic dxag losses) due t o surface roughness and wing distortion.

NETHODS AKD p w m r c DESIGN DATA

Methods and parametric 3 e s i m data were established f o r evaluating per- formance degrahtion (aerodynamic drag l o s s ) i n terms of fuel increment due t o surface roughness and wing d i s t o r t i o n (due t o deflection). Performance degra- lution was investigated f o r the following types of roughness and distortion of t h e wing Unifomly Distributed or Equivalent Sand Grain Roughness This type of roughness results from t h e unpolished condition of t h e wing skin, coatings on the wing skin, sptwelds, o r anything e l s e t h a t mars t h e fin- ish of the wing skin. The uniformly distributed roughness increases t h e fric- t i o n drag throughout the e n t i r e f l i g h t regime.

The incremental drag contribution due t o t h e uniformly distributed (sand grain) roughness was assessed with the computer program described i n ref- ereme 24-1. This program, which was developed at NASA Langley Research Center, combines the Sommer and Short T' method (ref. 24-2) and Goddard's method (ref. 24-3) t o compute skin f r i c t i o n drag coefficients on a f l a t p l a t e with A l l portions of t h e wing were presumed t o variable and sand grain roughness.

have the same surface roughness. The drag increments due t o various degrees of surface roughness were assessed over the nominal f l i g h t profile. Perform- ance losses due t o partia.1 areas of rou-ghness a r e determined by reducing t h e f'uel increment using t h e r a t i o of t h e p a r t i a l area t o the t o t a l surface of the wing.

Sheet Metal Joints and Fasteners Surface protrusions and c a v i t i e s a r e produced by various sheet metal j o i n t s and fasteners. These surface imperfections produce pressure drag at all f l i g h t speeds.

A s suggested bg Hoerner i n reference 24-4 (Chapters 'j and l7), the e s t i - mation of the drag contributions due t o the sheet metal joints and f a s t e w r s w a s based on the local flow properties within that part of the boundary layer affecting the protuberance o r cavity. An appropriate form drag coefficient brised on the shape o f the surface imperfection and the l o c a l %ch number was determined. This drag coefficient, combined with the l o c a l dynamic pressure, was used t o estimate the drag contributions of the sheet metal j o i n t s and fasteners .

Two-dimensional Surface Waviness i n Which the Wave Crests a r e Perpendicular t o the Wing Chord This type of wing distortion may r e s u l t from fabrication tolerances and deflections i n the wing skin due t o a i r loads and thermal effects. The surface vaviness contributes pressure drag primarily during transonic and supersonic flight.

The pressure drag contributions of the surface waves.were estimated w i t h the linearized inviscid theory presented i n reference 24-5. The drag estimates txoduced by the inviscid theory a r e expected t o be s l i g h t l y conservative.

drag contributed by two- Test dzta reported i n reference 24-6 indicate t h a t the dimensional surface waviness on an ogive cylinder decreases from values pre- dicted by the inviscid linearized theory a s the r a t i o of the boundary layer height t o the wavelength i s increased. Unfortunately, there i s insufficient at t h i s t i m e t o quantitatively establish the effects of t e s t data available the boundary layer on the drag contrtbutTon of the surface waviness. However, the r a t i o of the average depth of the boundary layer t o the length of the surface waves (distance between spars) f o r the candidate wing concepts of t h i s program i s l e s s than the r a t i o t h a t existed for the t e s t s of reference 24-6.

Therefore, the effects of the bound.ary layer upon the drag produced by the surface waviness should be less than t h x e observed by the t e s t results.

Thus, the inviscid theory, although s l i g h t l y conservative, w i l l produce valid estimates of the drag due t o surface waviness f o r the candidate wing concepts.

The performance degradation due t o two-dimensional surface waves with wave crests perpendicular t o the wing chord was determined, The surface waves were taken t o be sinusoidal i n cross section shape. 1% t h e wave6 assumed the shape of a circular-arc, the performance degradation would be 8 percent The a d d i t i o m l f i e 3 greater than that produced by the sinusoidal waveform.

required t o perform the fixed range mission i s prametrFcaLiy i l l u s t r a t e d as a function of the height of the wave, E , and the wavel-engkh, A . Surface waves w i t h constant values of € / A were assumed t o exist over the e n t i r e wetked area of the wing. Performance degradation due t o partial areas of swface waviness a r e calculated by multiplying t'ne &el increment by the r a t i o of the area of the distorted portion of the wing t o the t o t a l wing surface area.

Three-Dimensional Surface Bumps or Depressions A i r loads, thermal effects, or fabrication tolerances may produce t h i s type of distortion i n surface panels whose outer edges are attached t o r i g i d structure, Pressure drag i n the transonic ana supersonic speed regime i s pro- duced by the surface bumps, Performance losses due t o three-dimensional depressions o r bumps were defined using the linearized inviscid theory o f reference 24-5.

The f u e l increment required t o compensate for the surface distortion i s presented as a flunction of € / A , where X i s the length of t h e depression measured parallel t o the wing chord (the chord-wise length of the wing panel) and E i s t h e dis- placement of the wing surface at t h e center of t h e panel.

Again, the bumps or depressions were assumed t o exist over the e n t i r e wing surface.

If only a portion of the wing surface area i s distorted by the depressions, t h e per- formance losses a r e calculat?d by applying the distorted area/total area r a t i o t o the f'uel increments.

Surface Corrugations P a r a l l e l t o t h e Wing Chord This source of roughness i s the r e s u l t of the beading or corrugations incorporated i n t o the design of the wing skin and heat shield. !These swface cormgations contribute pressure drag at transonic and supersonic speeds and increase f r i c t i o n drag at a l l speeds.

The performance degradation due to the corrugations incorporated i n t o ;he wing skin was determined f o r the applicable concepts, using t h e linearized inviscid theory of reference 24-5.

The corrugations a r e parallel t o the wing chord. The wave drag due t o t h i s form of roughness i s generated a t the front and rear face of the end closeout of the bead o r trough. Skin f r i c t i o n drag i s increased due t o the increase i n wetted area resulting from t h e corrugations.

The fuel increment required t o compensate for t h e drag caused by t h i s class of roughness i s presented as a f b c t i o n of € / A and L/h, where A is the width and E i s the height or depth of the bead o r trough and L is t h e distance between t h e end closeouts. When t k 2 end closeouts occur a t t h e Leading and t r a i l i n g edge of the wing, an effective length of 80 f t is used t o determine the L/A value. If there are no end closeouts, a value of 00 i s taken f o r L/A param- eter. Constant values of €/A and L/A were assumed t o e x i s t continuously over the e n t i r e wing surface. When a f l a t area e x i s t s between adjacent beads or troughs, the f u e l penalty i s determined by multiplying the f'uel increment by the cor.rmgated a r e a l t o t a l wing surface area ratio.

The corrugated surface area i s t h e sum of areas of the individual corrugations, where the area of a single bead or trough i s LA. In the case of t h e corrugated heat shield, the corrugations incorporated a single low drag end closeout on or near t h e lead- The performance degradation due t o t h i s type of closeout w a s deter- ing edge.

mined f o r each individual candidate wing concept; based on linearized inviscid theory.

Deformation of the Primary Wing Structure Themal and a i r loads produce spanwise and chordwise deflections of the This type of wing d i s t o r t i o n increases t h e z e r o - l i f t primary wing structure.

drag and a l t e r s the induced drag characteristics of the vehicle throughout t h e f l i g h t regime.

Distortions of the wing were dekemined a t the Mach 8 cruise condition.

These distortions were assumed t o exist thrc-ighout the e n t i r e mission. The increased drag due t o the various tyy.13 of ioughness and wing d i s t o r t i o n were assessed over the e n t i r e speed regime, using the computer program described i n reference 24-7. The performance penalties resulting from the increased drag were determined f o r the vehicle using the nominal acceleration schedule f o r t h e climb-acceleration f l i g h t mode and the nominal speed-altitude schedule f o r a l l phases of the ..tssion, The takeoff weight of the vehicle remained a t 550 000 lb.

Pare.. . - i i c Design Ihta * L a i G & i , r t . , L %!sign curves for various tries of roughness,, as d+ocussed eal.ller, are s:;om i n figures 24-1 through 24-4.

Etraluatinc Approach The incremental drag changes due %o the s i x types-of roughness and dis- t o r t i o n represent the drag iifference btztveen t h e rough, distorted wing and an ' i e a l l y smooth wing. The wing of the nominal vehicle was defizsd t o have an m o u n t of roughness and distortion t h a t would groduce a drag increase equal t o 10 percent of the smooth wing f r i c t i o n drag, The nominal wing roughness would be equivalent t o a f u e l penalty of Therefore, 1110 IF, and was compensated f o r the! nominal mission performance.

the PJel penalty used i n the concepts evaluation pmcadure i s the difference between the f u e l increment determined for t h e candidate wing concept and t h e f u e l increment of 1110 l b resulting from the roughness and d i s t o r t i o n t h a t was assumed for the nominal wing.

Performance degradation due t o surface roughness and wavinecs was evalu- ated for the heat shield and leading edee concepts, such that a f i n a l selec- t i o n of the heat shield and leading edge concepts could be accomplished.

A performance degradation evaluation was conducted f o r each of the s i x structural concepts, including the thermal protection system and leading edge for constant mission range. A performance comparison was conducted by com- paring a l l of the s t r u c t u r a l concepts i n terms of fuel/payload Increment.

The f u e l increment f o r each of the s i x concepts was input into the interaction factor evaluation.

The performance penalties resulting from the combined roughness and distortion of the wing a r e summarized i n table 24.-l f o r the candidate struc- t u r a l concepts.

74-4

Monocogue Waffle The surface finish of the wing skin of the concepts evaluated is smooth enough to result in no performance losses due to imiformly distributed (sand- grain) roughness. The waffle panels undergo three-dimensional surface dis- tortion, which results in a fuel increment of 31 lb.

The waffle panels are connected with a butt joint every 4 3 in., measured in the chordwise direction.

The corrugated heat shield has a lap joint every 4 3 in. Tbese sheet-metal

joints, plus those of the segmented leading edge, produce a f ’ u e l penalty of 19 lb. The corrugated heat shield and the end closeouts for the heat-shield

corrugations result in a fuel loss of 118 lb. The wLng deflections (figs. 24-5

through 24-8) for the cruise-limit loads were used to determine the fuel penalty due to wing deformation, which is 611 lb. The total fuel increment due to the combined roughness and distortion of the monocoque wing concept is 779 lb.

MonocoqJe Honeycomb Sandwich The f’uel penalty caused by three-dimensional distortion of the honeycomb panels is 282 lb. This value is larger than that for the monocoque waffle con- -ept because of larger thermal deflections (thermal gradients) imposed on the honeycomb sandwich. The joints, fasteners, and the segmented Leading edge cause a fuel penalty of 155 lb, The corrugated heat shield in the lower out- board surface results in a fuel increment of 1 1 8 lb. The f’uel penalty attri-

buted t o the wing distortion is 4 5 8 lb (figs. 24-9 through 2 4 - 1 2 ) . The total

fuel increment required to compensate for the roughness and deformation of this wing concept is 1013 lb.

Semimmocoque Spanwise mbular This concept has corrugated heat shields on all exposed surfaces and a segmented leading edge. The f’uel penalty caused by three-dimensional panel distortion is 73 lb. The lap joints of the heat shield, spaced every 9.0 i n . , and the sheetmetal joints of the leading-edge have a f’uel penalty of 23 lb.

The fuel penalty due to the corrugations on the upper ar-d lower heat shield is 427 lb. The fuel penalty attributed to the wing distortion is 314 lb

(figs. 24-13 through 2 4 - 1 6 ) . The total fuel penalty for the combined rough-

ness and wing distortion is 837 lb.

Semimonocoque Spanwise Beaded Skin This primary structure concept incorprates the corrugated heat shield and a segmented leading edge. The fuel penalties resulting, from the sheet- metal joints, corrugations, and primary-structure deformations are identical t o those of the previous primary-structure concept, ‘fie surface panels of these an wing concepts are subject to three-dimensional distortion, which introduces 8 1 . - i b fuel penalty. The total fuel penalty for the concept due to the rough- ness and distortion of the wing is 8 4 5 lb.

24-5 Semimonocoque Chordwi se Convex-BeadEd/Tuuular This candidate wing concept has convex-oeaded. panels on t h e upper sur"nce of the wing, which r" dire no heat shield, and tubular Fanels with a corrugated heat shield on the lower wlng surface. The f u e l penalty produced by three- dimensional distortion is 139 lb. %e Zap j o i n t s of the heat shield, spmed every 24 i n , , and the sheu5 metal j(Iirits of the segmented leading edge instal- lb. beads of the upper wing l a t i o n introduce a fuel l o s s of y, The convex skin have an end closeout every 21t in. Zllne f u e l penalty due t o ihe corruga- t i o n s of the upper wing skin and the corrugations of :he l9wer surface heat shield i s 1841 lb. Yhe f u e l increment a t t r i b u t e d t o the wlng ,listortion i s The t o t a l f u e l increment required. t o csmpen- 521 Ib (figs. 24-17 t o 24-20).

sate f o r the roughcass and deformation of t h i s wing concept i s 2553 lb.

S t a t i c a l l y I>e"urminate This ccncept has t h e leading edge and corrugated heat shield. empleyed by the spandise-stiffened semimonocoque cone3pts. T ? x l e p j 3 i n t s of the heat shield r e s u l t i n a f u e l penalty of 30 l b for t h e shee% metal j o i n t s and fas-.

teners. The surface panels d i s t o r t threa-dimensionally, producing a fuel genalty of 195 lb. The f u e l penaX",y f o r tke wing defonns.tion (figures 24-21 t o 24-24) is 383 p~unds, and t h e t o t a l f u e l increment required t o wmpensate f o r t h e roughness and distortion of the wing i s 1040 l b , Fuel Increment I>lumnary The performance penalties resulting from the Trasious ty-pes of roughness and. distortion of the wing are s m a r i z e d f o r the s i x candidate wing conceopts i n table 24-1. The t o t a l Puel increment f o r the combined. roughness and dis.,.

t o r t i o n of each of the candidate wing concepts i s nompmed t o t h e f u a l incre- ment of 1110 lb, allowed to compensate for t h e assumed roughness of the nomi- nalwing. The net; difference between the f u e l increment dekemined for a wing and the nominal 1110-1b f u e l increment i s also l i s t e d i n table 24-1, for each ,Jf the candidate wing concepts. A s shown Ln table 24-1, t h e concept fuel i n incremente a r e less than +.he nominal fuel increment except for the chord- wise concept The f u l l y heat-shielded surfa-es have no a-ppreciab1.e &rag increase over IXoweveT, a r e l a t i v e l y smooth ( p w t i a l l y shielded ) con-ept, such as t h e waffle.

unshielded upper surface panels wi%h bwds (chordwtse concept ) protruding i n t o the air stre:i?1 provide the niost drag, -.-en though *the beads are orient& In the direction of flow.

Using t h e net fuel increnmts f o r each concept, t h e fuel mss fractions for the baseline vehicle shown i n t a b l e 24-1 were determined f o r i n w t i n t o tk interaction evaluation fackor investigation.

The performance degrada-tion resultirig f m m t2e surface roughness, sheet-

metal joirrts and fasteners, suyfaca wm-iness , conugations, and deformation

of the prizary wing structure has bee= eveluated f o r the four heat shields used with the spumise tu$ulm structure. 3% evaluations are summarized i n table 24-2. Wing deflection drag ( d e f o m t i o n of pr3nax-y s.l;,wcture) is inci.idd t o indicate relative &rag of heat s h i e l h .

! b e corrugat.ed sheet me5aril heat shie2d en the upper and lower wing sur- races vas considered first. The surface f5nish on t h i s and all of t h e other heat shield concepts i s sufficiently mooth to cause no psrfomance penalties, but the surface of the cormgated heat shieSds suffers three-dimensional wave distortion, resuhtlng . i n a f a e l penabty of '13 Ib. I n addition, t h e skin of this heat sbie3.d has a rear-facfng l a p j o i n t every 90 b., which along with the joints and fasteners associated with the segmeBted leading edge, cause a 'The corngations of the beet sh5eld a d the end close- f u e l penalty of 23 lb.

outs OB the corrugatioizs near the Leading edge r e s u l t in a fuel penalty of k27 lb. Since a l l heat-shield concepts vere appl-ied t o the same primary struc- ture, the fuel inerrmen% of 314 A b h e t o the de5'omation of the primary struc- The t o t a l fuel increment due t o t h e roughness ture is cornon t o all. concepts.

anfi diskortlon of the wing for the cormgated heat, shield coacept. is 837 lb.

The second concept :o:isiderb.has a flat, dbple-stiffened skin en the upper ami lower surfaces, These panels are exbJc : t t o three-dimensioncl tave and the Riel increment h e t o t h i s surface waviness is 43 l b .

distortion, The g%aels also have a chordwise butt joint every 15.3 in.

The f u e l penslty due t o these sheet-metal joints aad those of the segmented leading e&e i s 31 lb, The to%al fuel increment for t h e combined roughness and distodlon of the wing w i t h the flat rskin, dimple-stiffened heat shleld i s 388 lb.

The third lieat shield concept consists of the simply suppart&, moduiar heat shiel& on t h e upper wing surface and the corrugated heat shield on the Lower swface of the wing. Age.-Ji-., the panels incur three-dimensional wave distortion. The fuel penalty resulting from the surface T.;aviness i.s 5 lb.

The skin of the modular concept has a rear-facing chordwise l a p j o i n t every 13.4 in., and the lower surface has a lap joint every 90 in. These sheet- metal joints, cambined wi%h the joints anzl fasteners of the segmented leading edge, resLlt i n 8 . fuel yemlty of 58 lb. !&e corrugations on the lower surface heat; shield cause a fuel p n a l t y of 231 lb. The total wing Fuel penalty for the simqz;Y supportxd mndular heat shield is 603 pounds.

The faurth arrangement, the cantilevered modular heat shS.elCi, is used on the u p p r surface. The surface waviness is identical t o t h a t of the third concept. The cantilevered modlu3ar heat shield has a rear-facing choi-dwise l a p j o i n t eveiy 2.61 in. The fuel penalty for the Lap joints and the sheet- metal joints of the leading edge is 1 4 . 9 lb. The t o t a l f u e l increment result- ing from the roughness aad dtstortSon of this wing concept i s 699 3.5.

24-7 The fuel increments required t o compensate f o r the combined roughness and distortion of the four wing concepts a r e a l l l e s s than the 1110-lb f u e l increments i n i t i a l l y allowed t o compensate for the assumed roughness and distortion of the nominal wing. As a result, the net f u e l increments c .

paylcad decrements used i n the evaluation pocedures have negative values for each of the c a d i d a t e heat-shield syskeas.

LEADING EDGE The performance degradation resulting from the sheet-metal j o i n t s and fastxners, and the corrugation and closeouts have been evaluated for the segpeilted and the continuous leading-edge concepts. The end closeouts f o r t t e corrugated heat shield are located i n the leadiQ3 edge for the continuous leading-edge concepts. “he segmented leading edge is cylindrical i n shape, requiring that the end closeouts of the corrugations be located i n the heat shield just behind the leading edge. The geometric characteristics of the end closeouts are the same for both of the leading-edge concepts and r e s u l t i n identical performance degradation. Because of 8 joggle joint at the attach- ment of the leading edge with the wing panel, there i s a fuel penalty of LO lb for either concept. In addition t o the joggle joint, the segmented leading edge has an e x p s i o n gap between each 20-in. segment. &,ch segment is fas- tened t o the wing structure with flush-mounted screws.

Because of the drag contributed by the expansion gaps and the flush-mounted screws as well as load deflection, the f u e l penalty associated v i t h the segmented leading edge adds another 10.2 I t . Therefore, the f‘uel/pa:,-load incremeEts f o r the continuous and segmented leading edges a r e 10 l b axid 20.2 l b , respectively.

24-1 Aker, L . J . : A Digital Computer Program f o r Calculating Airplane LR 19620, April 1966.

Turbulent Skin Friction Drag.

24-2 Peterson, J . B.: A Comparison of &pe&mental and Theoretical Results for the Compressible Turbulent-Boundary-Iayer Skin Friction w i t h Zero Pressure Gradient. NASA TN D-1795, March 1968.

Wfect of Uniformly Distributed Roughness on Turbulent 24-3 Goddard, F. E.: Journal of the Aero-Space Skip-Friction k a g at Supersonic Speeds.

Sciences. January 1959.

24-4 Roerner, S . F. : Fluid-Dynamic Drag. Published by t h e a t thor 1965.

24-53 Smith, K. G.: The Increase i n Wave Drag at Supersonic Speeds Due to R c y & 1 Aircraft Establishment Technical R e p o r t N o .

Surface Waviness.

No. 65173, August ls’j.

??ressur@ Drags Due t o Two-Dimensional 24-6 Czamecki, K. R.; and Monta, W. J.: FaSrication-Type Sllrface Roughness on an Ogive Cylinder at Transonic ‘Speeds. NASA T ! X E-3519, August 1966.

A Digital Computer Program f o r Calculating Pressure 24-7 Elliot, R. D.: Distributions and Aerodynamic Coefficients of Warped Wings of Arbitrary LR 18764, February 1.966. Planform i n Supersonic Flow.

I

00 0

m 3

0 k d Q, d I rl Q) f- t ?

I I 311-10

- --

In rl m rl cu rl rl rl rl rl _I d rl d d

-

.002 .m .006 .m Equlvqlent sand rwahncsr height, ft Fuel increment required t o compensate for wnifo-mly distributed Figure 24-1, roughness on wing surface f o r constant mission range 24-12 Airflow direction __I_z.

e + A - 4 T

10 x 103 - Wave crests ore psrpendicutc* A

to wing chord 0 .004 .012 .OM .020 = / A f i e 1 increment required t o compensate for mifomn waviness Figure 24-2.

over wing surface f o r constant miss5.on range 24-’ 3 Air flow direction 8 x 0 .004 .008 ,012 ,016 ,020 4 A

-------- I-

Figure 24-3. Fuel increment required t o compensate f o r uniform three dimensional waviness over wing surface for constant range 24-lh Figure 24-4. Fuel increment required to compensate f o r uniform corrugation in wing surface for constant mission range 24-15 E ...

- i

E O

3 i

-2 -4 .A Figure 24-5. Fuselage deflections net due t o ZLmiZ; load6 along BL 120 monocoqua waffle concept (intersection of fuselage and wing) 24-16 2 16 I - t ?

' 2 Y z d 12 -4 0 100 m 300 400 500 600 Wlng statton, I n .

Figure 24-6, Wing deflections net due t o litult loads, monocoque waffle concept 24-17 Fuselage station, in.

Figure 24-07, F'usclage deflection due t o thermal stresses along BL 120 (intersection of h s e l a g e and wing), monoc<.que warflr concept a l -4 Wlng cmtlon, In.

Wing deflections due t o %hemal str*esses, monocoque waffle concept Figure 24-8.

34-18 1 6 3.2 -4 Wing Station, in.

F.i@re 24-10. Wing deflections net due t o littiit lm.ds,honeycomb concept

I I

rl +2.0-g maneuver

- - - Cruise

1 6

2 i 12

rr @

g a

-4 0 100 200 300 400 500 600 wing Station, in.

Figure 244.2, Wing deflections due t o therm1 stresses,honeycomb concept .

I 2h-22

.- 2 4

C

._ .-

-

c $ 2 -2 0 400 800 I200 1600 moo 2400 2800 3200 Fuselaga station, in.

Figure 24-13.

Fuselage deflect5on.s net due to l i m i t Zoadc e h n g BL 120 (inter- section of f’ueelage and wing), semimonocoq~e (syanwise) concept 24-23 2 0 C .- s' 5 12

-

;f; -4 0 100 2M) 3M) 400 Wing station, in.

Figure 21~li.. - Wing deflections net due to limit loads, semimonocoque

(spanwise) concept 24-24 -2 -4 0 400 800 1200 1600 2000 2400 2800 3200 Fursloga station, in.

__ _. .

Figure 24-15. Fuselage deflections due to therms1 stresses along BL 120 (intersection of fuselage and wing), semimonocoque (spanwise) concept 2 4 ~ -4 100 200 390 400 500 600 Wino station, in.

Wing deflections due to thermal stresses, semimonocoque Figure 24-16.

(spanwise) concept E: %

P

a3 cu cu OJ a3 I I - 0 100 2oc 300 400 500 600 ' Wing Station, In.

Figure 24-18. Wing deflections net due t o l i m i t loads, sernimmac ogue (c hordw ise ) concept 24-27 24-28 +2 .O-g mmeuver -_.

- - -- Cruise

0 100 200 300 400 500 600 700 Wing Station, in.

Figure 24-20. Wing deflections duo to t h e m 1 s.tmsseq semimonoosque (chordwise) concapt -2 -6 Figure 24-21, Fuselage deflectione net due t o Iimir, loans along BL 120 (intersection of fieelage anu wing), statically deb ? m r L i m t e concept.

24-30 i

.-

.- t

t 8

-

-4 0 100 200 300 400 600 Wing station, in.

Figure 24-22. Wing deflections net due t o l i m i t loads, statical3;y determinate concept; 24-35.

-2 400 1200 1600 M(yJ 3200 F u ~ l c ~ e station, in.

- _ Figure 24-23. Fuselage deflections due to thermal stresses along BL 120 (intersection of fuselage and wing), statically determinate concept I !

i

-

I -4 0 100 200 300 400 500 600 Wing station, in.

Figure 24-24. Wing defLections due to thermal stresses, statically determinete concept 24-32 I

Section 25 RELIABILITY bY C . C . Richie, G . W . Davis, I . F . Sakata, B. C . Wollner 25-i PRECEDING PAGE BLANK NOT FILMED.

CONTENTS METHOD O F ' EVALUATION ?5-?

HEAT SHIELD RELIABILITY 75-3 L F M I N G EDGE FELIABILITY

75 - 3

PFUMARY STRUCTIRE RELIABILITY ?,s-L S L ? Y OF CONCEPT RELIABILITY EVALUATION 2s-iii PRECEDING PAGE BLANK MOT FILMED.

'I'atilc NO. Page No.

25-1 Summary of r e l i a b i l i t y parameters 35-6 P5-3 Heat Shield r e l i a b i l i t y evaluation 75-7 25-3 Leading edge r e l i a b i l i t y evaluation 25-8 Summary of component w i n g weights f o r low, 25 -h nominal, and high l e v e l s of r e l i a b i l i t y ,

csnter area (A) 25 - i

25-5 Summary of component w i n g weights f o r low, nominal, and high l e v e l s of r e l i a b i l i t y , inboard area (B) 25-19 Sumii;ary of component wing weights f o r low,

?s -6

nominal, and high l e v e l s of r e l i a b i l i t y , outboard area ( 6 ) 25 - S l Summary of r e l i a b i l i t y evaluation 25-12 25-7 25 -v FRECEBING PAGE BLANK NOT FILMED.

ILLUSTRATIONS Page No.

Figure No.

Safe overload boundaries and operating lwundarics 35 -1 f o r low, nominal, and high l c v c l increase i n structural rdi abi.lity 25 -13 Wing, invcsLigation area: average unit weights vs k c t o r of' safety ;'[;-I(; ;)$-1(, Total wing: average unit weight vs factors 01' safety PRECEDING PAGE BLANK NOT FILMED.

SYMBOLS X and y distance between simply supported edges of panel Ultimate t e n s i l e strength Ftu Fatigue allowable t e n s i l e s t r e s s Gravitational acceleration Fatigue quality index I n e r t i a load factor i n Xdirections Pressure P Limit pressure Ultimate prwsure Pult .

Flat thickness of leading edge ‘FLAT Nose thickness of leading edge ‘NOSE

t Equivalent thickness

Room temperature RT Seckion 25 RELIABILJTY The r e l i a b i l i t y analysis consisted of selecting a range f o r factor of safety and calculating s t r u c t u r a l weight f o r low, nominal, and high levels of factor of safety. The key factors, involving safety, creep, Tatigue, and maintainability were evaluated i n t h i s study.

METHOD OF EVALUATION The primary factors affecting s t r u c t u r a l r e l i a b i l i t y are : 1. The physical environment within the operating limits of the vehicle 3 .

Design accuracy, including accountabiiity f o r a l l possible contingencies 3. Consistency of the reproduced a r t i c l e s t o engineering reqiiirement s 4. Maintainability.

A numerical approach t o a s t a ' i s t i c a l probability evaluation is not possible because data do not exist t o substantiate t h i s approach. Instead, the basic approach must establish a consistent r e l i a b i l i t y standard, adequate f o r mission performance over t h e vehicle l i f e span, which a l l concepts must satisfy. Therefore, -bo satisfy the primary r e l i a b i l i t y factors discussed above, a s t r u c t u r a l r e l i a b i l i t y evaluation method was established which consists of parametric variation of tho key factors affecting the r e l a t i v e r e l i a b i l i t y ( s e n s i t i v i t y ) of t h e s t r u c t u r a l concepts, as measured by weight These key factors, involving factors of safety, creep, fatigue, and main- t a i n a b i l i t y , were used f o r three levels of s t r u c t u r a l r e l i a b i l i t y (low, nominal design, and high) and three f l i g h t load conditions (-0.5-g, +2.0-g, and cruise) as shown i n table 25-1. Also, figure 25-1 presents the overload and operative boundaries f o r the low, nominal, and high levels of factors of safety.

The design limit load factor of safety of lU30 was specified f o r the f l i g h t load conditions. Normal a i r c r a f t design practice s e t s t h i s factor at a value of 1.00. Normal a i r c r a f t faotors were considered the minimum (low) acceptable leveZj the required value of 1 . 3 0 was the nominal value; and an arbitrary design l i m i t load factor of safety of 1.67 was chosen for the high value, SSmilarly, factors of s a f e t y on thermal s t r a i n of 1.10, 1.30 (required), and 1 . 5 0 were used. Creep and fatigue factors of s a f e t y operating t h e were 25-1 selected at 1 (low), 1.5 (nominal), and 2 (high).

The fourth primary r e l i a b i l i t y factor, maintainability, concerned with long l i f e , damage tolerance, and slow crack growth ( f o r long i n s p c t i o n intervals), is provided f o r by the s e n s i t i v i t y measured by the design factors of safety variations discussed above.

I n addition, r e p a i r a b i l i t y was assessed by evalua+ing refurbishment requirements of leading edges and heat shields .

Accessibility f o r i n t e r i o r wing inspection and repair was s a t i s f i e d by using t o be removed.

mechanical fasteners t h a t permitted the wing panels Using the established r e l i a b i l i t y method, a parametric evaluation was conducted t o establish the s e n s i t i v i t y of each concept (weight) for the three levels of r e l i a b i l i t y (low, nominal design, and high). After evaluating one concept (waffle) f o r the key s e n s i t i v i t y factors l i s t e d i n table 25-1, it was determined t h a t the 2.0-g load condition was the most c r i t i c a l load Therefore, a l l condition, with creep and fatigue not governing the design.

i t h e concepts were evaluated f o r the 2.0-g load condition and the three levels of factors of safety. These concepts encompassed heat shields, leading edges, and primary structures .

HEAT SHIELD RELIABILITY Results of the heat shield r e l i a b i l i t y evaluations are shown i n table 25-2, with heat shields applicable t o a typical spanwise tubular panel (46 in. by 92 in.). For each load factor, the optimum heat shield consists of minimum-gage skin with the support spacing decreased t o allow f o r ipcreased pressure loading. Thus, variation i n the equivalent thickmess panel (t) is dc The multisupported corrugated heat only t o changes i n support spacing.

shield, f o r example, has support spacing of 15.3 in., 13.1 in., and 1.1.5 in.

f o r the three levels of r e l i a b i l i t y .

Panel siz,es f o r the flat-skin, dimple-stiffened concept are 23 in., 15.3 in., and 15.3 in. Became only heat shield sizes t h a t are multiples of t h e primary-structure paEel s i z e are considered i n the heat shield evaluation, t h e support spacing and t f o r noninal and high factors of safety are identical.

The next larger s i z e (23 i n . ) would have larger oending moments t h m allowed by minimum-gage design .

The weights of the two modular concepts are not affected by variations i n factor of safety, since they are not influenced by the support spacing of t h e primary s t r u c t u r a l panel.

The results indicate t h a t r e l i a b i l i t y ( s e n s i t i v i t y ) had l i t t l e influence upon f i n a l selection of the heat shield concept.

25-2 LEADING EDGE RELIABILITY The leading edge r e l i a b i l i t y eval.uation r e s u l t s are shown i n table 25-3.

As indicated, the segmented leading edge provides considerably more f l i g h t s than the continuous concept; and the nominal design f o r the segmented leading edge more than satisfies bhe vehicle desierl l i f e of 8110 f l i g h t s . The con- tinuous leading-edge concept does not meet the l i f e reqtlirements f o r any l e v e l of r e l i a b i l i t y studied.

PRIMARY STRUCTURE RELIABILITY Relative s t r u c t u r a l r e l i a b i l i t y ( s e n s i t i v i t y ) was based on average unit weights f o r the e n t i r e wing cross section, To determine average unit wing weights, a spanwise distributicn Sased on t o t a l wing critsjs section weights tn the center ( A ) , inboard ( B ) , and outboard ( C ) wing areas was used f o r the dng-.:.nvestigation area. Then t o t a l weights were obtained. The wing weights fncliide upper and lower swface panels, spar caps and webs, r i b caps and webs, heat shields, insulation, panel closeouts, oxidation penetration, corner ,posts, fasteners. A s an exmple, tables 25-4 through 25-6 present a summary of component weights f o r the monocoque waffle concept f o r the three levels of r e l i a b i l i t y .

The r e l i a b i l i t y evaluation results f o r the s i x p r h a q , structures are shown i n table 25-7 f o r the win,g--inirestigation area and the t o t a l wing. The monocoque w a f f l results show comtant variation i n average wing weight of about 1.3 l b / f t 5 between levels of r e l i a b i l i t y . For the monocoque honeycomb- sandw'ch concept, the constant variation i n average wing weight is about 0.20 lb/ft between levels of r e l i a b i l i t y . For the spanwise tubular concept, the results indiccLe variations i n wing weight of abou" 0.30 lb/gt2, For the beaded-skin concept, a constant v a i a t i o n of about 0,40 lb/f't was indicated.

The chordwkse co-ncapt results indicate variations i n wing weight of about 0.65 lb/ftc: between the low and nominal r e l i a b i l i t y levels and about The s t a t i c a l l y 0.45 lb/ft2 between the nominal and high r e l i a b i l i t y levels inate concept results indicate variations i n wing weight of about 0.40 dete lb/ft ??l For the fatigue r e l i a b i l i t y evaluation, discrete loading spectra were used t o arzive a t a loading distribution (actual n h e r of cycles applied at dlscrete load l e v e l s ) f o r cmulative damage analysis. A fatigue-life versus allowable stress plot, based on the Palmgren-Miner cumulative dmag6 theory, provided a direct-reading method of determining the potential penalty (reduced Results of the fatigue-reliability allowable stress) f o r increase i n lifetime.

Fatigue l i f e requirements f o r l ~ w , evaluation are shown in figure 25-2.

nominal, and high levels of reliabilit3; were based on s c a t t e r factors of 1.0, 1.5, and 2.0, respectively, applied t o the specified vehicle l i f e of 10 000 Between low and nominal le-mls of: r e l i a b i l i t y , t h e allowable hours at l4OOOF.

mean stress a t oruise decreased 6 ksi.

, 25-3 The effect of creep on primary s t r u c t u r a l panel dssign was determined f o r the cruise condition loads and temperatures, and s c a t t e r factors corres- ponding t o low and high levels of r e l i a b i l i t y were applied t o the t o t a l cruise time. The resulting structures, optimized f o r creep only, accounted for only 70 percent of the weight of structures designed f o r the maneuver conditions and checked f o r creep life. m u s t be Therefore, creep conditions evaluated, although they are not c r i t i c a l t o the design.

SUMMARY OF CONCEPT RELIABILITY EVALUATION Reliability-. valuation results f o r the selected monocoque , semimonocoque , and s t a t i c a l l y determinate primary structure concepts are smmarized i n figure 25-3 f o r the wing investigation area and i n figure 25-4 f o r t h e t o t a l xing. As shown, for low, nominal, and high levels of r e l i a b i l i t y , they represent ultimate factors of safety of 1.5, 2 . 0 , and 2.5, respec-Lively.

Average unit wing weights were based on loads f o r the 4-2.0-g maneuver condition.

As shown i n figure 25-3, the chordwise concept is lower i n weight than *he honeycombsandwich f o r t h e low, but not high, reliability. This is due t o the minimum-gage restraint of t h e honeycombsandwich, The t o t a l wing weight evaluation of f i 4 a 25-4 indicates t h a t t h e minjmum-gage honeycomb-sandwich is heavier than t h e statically determinate concept f o r t h e low reliabilitg. However, t h e honeycombsandwich is lower Zn weight than both t h e statically determinate and tubular concepts at high (2.5) factors of safety, which indicates greater honeycomb efficiency i n the higher load raxges.

25-1 II(*ldenl'cls, 12. R.: The El'i'c:ct of Nonuniform Tcmpcraturc 0i:;tributionn on tho Strcsscs and Distortions of Stiffcncd- Skiell Structures. NACA TN ?%4U, Nov. 195'3.

TABLE 25-1 SUMMARY OF RBLIABIIJTY PARAMETERS -0.5-g and +2.0-g load conditions Life criteria for primary structure (applied to operating limit loads) (fatigue and creep allowables) ~ I _ b Fatiguea Creep Reliability load Ultimate thermal scatter scatter strain factor level level factor factor factor LOW 1.5 1.1 Low 1.0 1 . 0 Nominal 2.0 1.3 Nominal 1.5 1.5 High 2.5 1 . 5 High 2.0 2.0 a Applied to fatigue spectra.

b Cruise limit loads; 0.5-percent total creep tensile atrain; creep buckling based on isochronous s tress-strain curves.

26-6 In Q) E5 rl cu rl d rl 0 0 0 0 d d -4 IS

-

m m V rl W 00 m c3 6, l-l cu rl cv l-l rl 0 0 0 0 0 d d PC rl 00 M ey 6 , d ey rl ey rl 4 0 0 0 Q d 0 d d In c j II 25 -7 TABLE 25-3 LEADING-EDGE RELIABIUTY EVALUATION Leading-edge l i f e (number of flights) Level of reliability (a) @) (c) S t r u c t u d arrangement LOW Nominal High scatter scatter scatter factor = 1 . 0 factor = 1 . 5 factor = 2.0 Segmented leading edge tNOSE = 0.125 in.

10.0 x lo6 11.9 lo5 2 . 5 ~ 10

tFLAT = 0.030in Length = 20.0 in. i d ) Continuous leading 74 12 2 edge tNOSE = 0.625 in.

t~~~~ = 0.060in.

a Scatter factor applied to low-cycle fatigue strain allowable.

bFatigue quaIity index, = 2, applied to limit elastic thermal strain.

‘Analysis of end effect based on reference 4 1 .

dFor cumulative fztigue damage analysis, -0.5- and +2.0-g conditions are assumed to occur for one of ten flights.

25 -8 TABLE 75-h WING WEIGHTS FOR LOW, NOMII?AL,AND HIGH I;EvELS O F SWMARY OF COMPONENT RELIABILITY, CENTER AREA ( A ) (Monocoque 'uJal'l'le cuncep L: p a r t i a l heat s h i e l d a t outboard area lower surface with insulation; a = 40 in.. b = 20 in.)

---

ness, i n .

I Equivalent Thic Low . Nomina? FL gh I -t; e 1 9 factor

I factor

factor 0.06173 'Jpper 0 . Orf925 0.07082 LOWCu.

0.07'i'Ll 0.063*!', 0.056211- '?*-.*,.[I 0.1.1737 0 . .I 91.j6 0.1.36yf 0.01981, O .O18; 0.01G)I:? dpper r i b direction 0.00821

0 .00333 0 . OOr)l.$ Upper spar directim

0.0:!w1- ' i ' t l i :1.1

0 . o:!(t'lR 0.O;L'plo

1':tp :I111

0.0.1 yo8 0.01 "ill.,, 0.01 .:)I> Lower r i b direction

0.0081,h 0 . 00'/'[; 0.006y;-: Lower sgar d i r e c t i a O .o?~,G;: 0.02315 0.02c1.7 'i'c? t a1 0.04481 Tot a1 0.05540 0 -05055 0.0363 0.0363 0.0363 R i b web Iiib and 0.0182

o . o m 0.0182 Spar web

spar webs 0.0545 0.0545 Total 0.0545 0.00225 0.00225 Total 0.00225

-

-

Insulation 1 Jyr1:Lfl c:x

- -

Yaclcag-i ng insulation

- -

Il'otal

-

Corrugation

I

Ccrrrugatect

-

Clip heat shield Tot a1 0,000498 Tot a1 0.000498 0.000498 Ox idat i o n 0.00541 Fastener Total 0.00541 0.00541 0.22544 0 27130 ~1.61; 9 e67 Total unit weight, 1b/ft2 I 75-9

T A B m 25-5

WING WEIGHTS FOR LOW, NOMINAL,AND HIGH I;EvELS OF SUMMARY OF COMI'ONXNT RELIABILITY, INBOARD I ? ) (%onocoque waffle conccpt: partial httat shift d at; outboard area lower surfacl: with insulation; a = 40 in., 1) = 20 i n . ) i n .

D I. Th I c l u u Nominal High It em fact or factor factor 0.06899 0.07904 0.08852 Upper P:mels L,cwcl- 0.08224

i 0.07035 0.09395

Total ? .13924 0.16128 0.18247 . - - Upper r i b direction 0.0177c1 0. ozllc7 0.01975 Upper spar direction 0.00888 0.00988 0.01074 Tc t a1 0.02664 0.02963 0.03221 Cap and

close out -

Lower r i b direction L, .01560 0.01791 c).01981 single shear Lower spar direction 0.00780 0.00895 o.oogg1 Total 0.02340 0.02686 9.02972 Total 0.05004 0.05649 0.06193 --_ R i b web 0.0363 0.0363 0.0363 .cib and 0.0182 0. 0182 0.0182 Spar web spar webs Total 0.051+5 0 .og'c5 0.051~5

- -

Tot a1 0.00227 0.00225 0.00225 W e b intersect ion Lnsulat ion Dynaf lex

-

Packaging insulation

-

Tot a1

-

Corrugation Corrugated ..

Clig heat shield

-

Total Oxfdat ion 0.000498 0.3001~98 Fastener Total 0.00541

Total equivalent, thickness 0 . 2 5 l.91-1

Total u n i t weight, l b / f t 1.0.63 SUMMARY OF COIVIPONENT WING WEIGHTS FOR LOW, NONINALqAND HIGH LEVELS O F RELIABILITY, OUTBOARD AREA ( c ) (Monocoque waffle corlcep L: p a r t i a l llca t s h i e l d aL uu Lboard area lowar surlacc w i t h insulation, a = 40 in., b = 20 i n . ) in.._- High I

I Item factor factor factor

0.05962 0.068110 0.07660 0.03871 0. 0Ji.y> 0. O l j l G g 0.1.l.:363 0 . 1 ; 1 8 ; 1 9 0.09833

0 .Ol>Z)J. 0 . OL'il!')

Upper rib direction 0.018yI 0 OU1'7', 0.0080; 1 0.00()30 Upper spar direction 0,0;!:3;'(1 Y!oL:il 0 O;!lj@( 0.0;.!#189 0 .0092'( 0.OIOG5 0 .O11.'[8 Lower r i b direction Lower spar direction 0.00)+64. 0.00532 0.00589 Total 0.01391 0.01767 0.01597 Total 0.04184 0.03717 0.04556 Hib web 0.0182 0.0182 0.0182 R i b and Spar web 0.0091 0.0091 0.0091 spar webs T o t a l 0.0273 0.027: 0. C273

Total o .001125 0.001125 0.001.l.2~

W e b intersection

-

Insulation 0.001h.6 0.001h6 0.001)16 Packaging 0.0020;1 0.00202 0.0020; 1 Total 0,003118 O.Ooyl8 0.00311.8 Corrugation 0.01660 0.01.660 0.01660 Corrugated

Clip 0 00485 o I ooh.85 0.00485

heat shield T o t a l 0 02145 0.02145 0,02iJ+5

- -

Oxidat i o n Total 0.005664 0.005664 0.00 5664 F a s t e n e r T o t a l .

0 00541 0.00541 0.00541 Y- Total eauivalent thickness 0,21992 0.23828

o 19993

-

Total unit weight, lb/f'tz 8.58 10.23 9.44

- . -

75-11 .

t-

t

25-12 Foilure region

3 . 0 -

Safe overload region A. Low-ultimate design

CN 2 . 0 -

. factor = 1.50

c b Operating boundary U

-

9 1.0 -u - 0 (3 . Speed V

1 6 6 - d Fegibni - - -9

-1.0 Failure region Failure region

4.0 -

Sefe.overload region

3 . 0 -

N C 8. Nominalultimate design factor 2.00 U Operating boundary

---------

Safe overload region Failure region Failure region

5.0 -

-

4 . 0 Safe overload region * cN 3.0 .

A- b C. Highultimate

_ - - - - - . ) - - - - -

design factor = 2.50 Operating boundary

-------

Safe overload region Failure region -2.0 c F i p e 25-1 Safe overload boundaries and operating boundaries for l a w , nomina1,and high level increase i n structural --- .

- - r e l i a b i l i t y

-. - 25-13 Material: Rend 41

(Ftu - - 170 000 psi min at R T

Fatigue quality index, Kt = 4.0 Reliability levels Scatter factor Low 1 .o Nominol 1,s High 2 .o Temperoturs = 140OoF 10 20 3 0 x loJ Lifetime, hr Figure 25-2 Allowable tensile stress for faLlgue, Rend 43- 25-14 u cn I!

> Q: 1 .o 1.5 2 .o 2.5 Factor of safety (ultimate) Figure 25-3 Wing investigation urea: avrtxge unit rates vs factor; of safety 25-45 S8otkolly determinoto 1 .o 1 . s 2 .o 2.5 Foctor of rofety (ultimote) P PiC:;re 25-4 T o t a l wing: aver .:; 25-16

Section 26 RATING FACTOR l X ” ~ C T I 0 N by I . F . Sakata, R . D . Mijares, D . E . Sherwood 26- i PRECEDING PAGE BLANK NOT FILMED.

CONTENTS a g e INTERACTION PROCEDUFtE 20-1 Vehicle Weight C o s t S i z i n g Method 26-1 Cost Model Summary 26-2 I n i t i a l Investment 26-3 Total Operation Cost 26-3 Cost Model Program 2 6 4 INTE3ACTION RE=sULTs 2 6 - 1 4 Baseline Vehicle 26-6 Minimum Total System Cost Vehicles 26-7 Constant Weight Vehicles 26-8 INTERACTION SUMMARY 2 6 9 RESEXIBCES 26-10 PRECEDING PAGE BLANK NOT FILMED.

Table -€5e 26 -1 26-11 Summary of i n i t i a l investment cost elements 26 -2 Indirect one?&Ang expense constants 26-12 26 -3 Cost model coluputer program 26- 14 26 -4 26-21 Nomenclature fcr cost model Cost b r e a k d m i n dollars f o r each p r i m r y structure 26 -5 26-29 at each levcl of r e l i a b i l i t y 26 -6 Cost breakdown i n cents per ton-mile f o r each primary structure at each level of r e l i a b i l i t y 26- 30 26 -7 Group weight statement of sized vehicles 26- 3 1 26 -8 Baseline airplane mass fractions 26-32 26 -9 Geometry and des’-@ parameters f o r baseline vehicles 26- 33 26 -10 Cost results f o r baseline vehicle 26- 34 26 -11 Group weijjht statement of optimum-sized vehicles 26- 35 26 -E Optimum-size airplane mass fractions 26- 36 26 -13 Geometry and design parameters f o r optimum-sized vehicles 26 -14 Cost results for optimum-sized vehicles

Summary - structure concept design and cost data

26 -15

26- 16 Summary - wing weights and percentages for increase

and t o t a l system cost i n wing weight 26- 40 26-v PRECEDING PAGE BLANK NOT FILMED.

Figure w3e 26 -1 lhteractior. f a c t o r evaluation computer program 26- 41 26 -2 Total system cost f o r optimized vehicles of various wing constructions 26- 42 26 -3 Total system cost for various gross weight vehicles f o r the candidate wing constructions 26- 43

26 -4 Total system cost variation with vehicle s i z e -

monoc oque waffle c oncept 26- 44

26 -5 Payload and f l e e t s i z e variation with vehicle s i z e -

monocoque waffle concept 26- 45

26 -6 Total system cost variations with vehicle size -

monocoque honeycomb sandwich concept 26- 4 6

26 -7 Paylaad and f l e e t s i z e variations with vehicle s i z e -

monocoque honeycomb sandwich concept 26- 47

26 -8 Total system cost variation with vehicle s i z e -

semimonocoque spanwise tubular concept 26- 48

26 -9 Fayload and f l e e t s i z e variation with vehicle s i z e -

semimonocoque spanwise tubular concept 26- 49

26 -10 Total system cost variation with vehicle s i z e -

semimonocoque spanwise beaded concept 26- 50

26 -11 Payload and f l e e t s i z e variation with vehicle s i z e -

semimonocoque spanwise beaded concept 26- 5 1

26 -E Total system cost variations with vehicle size -

semimonocoque chordwise tubular concept 26- 52

26 -a3 Payload and f l e e t s i z e variations with vehicle s i z e -

semimonocoque chordw i s e tubular concept 26-53

26 -14 Total system cost variation w&th vehicle size -

s t a t i c a l l y determinate beaded concept 26-54 26-vii ILLUSTRATIONS (Cont. ) Figure page

26 -15 Payload and f l e e t s i z e variation with vehicle size -

s t a t i c a l l y determinate beaded conwpt 26-55 26 -16 T o t a l system cost variation with vehicle gross weight 26- 5 6 26 -17 Total system cost variation with operational fleet eiae 26-57 Total system cost variation with nominal wing unit 26 -18 weight for constant gross weight vehicle 26- 5 8 SYMBOLS Direct operating cost GTOW Gross takeoff weight IOC Indirect operating cost Initial investment cost

ov Operational vedcle

RDT&E Research, developnent, test and evaluation cost S Actual wing area Reference wing area sREF TOC Total operation cost TSC Total system cost W Vehicle weight Weight of payload wPL W Unit wing weight expressed in lb/ft ATSC Difference i n total system cost , J.’ 26- IX

Section 26 RATING FACTOR .T3UTl3.!3A@rION A rating factor interaction evaluation was conducted by interrelating the t o t a l wing factors of weight, cost, performance, and r e l i a b i l i t y t o a t o t a l vehicle system cost f o r each wing structural concept.

A common denominator, minimum t o t a l system cost (TSC), was selected as the basis for evaluating and comparing the wing-structure concepts. The baseline mission range requirement of 4000 nautical miles and a f l e e t s i z e of 200 vehicles (550 000 lb each) w i t h a payload of 55 000 pounds satisf’yfng 10 000 hours of l i f e (8110 missions) f o r 10 years resulted i n a f l e e t payload- range requirement of 205 b i l l i o n ton-miles (statute) f o r each concept.

The t o t a l wing weights and costs for the three levels of r e l i a b i l i t y and f u e l mass fractions associated with roughness drag performance (resulting i n payload changes for the wing structure concept of the baseline 500 000-lb vehicle) were submitted f o r integration i n t o a whole vehicle system. Except for the s t a t i c a l l y determinate concept, which requires additional fuselage weight, identical weight and cost scaling relationships were used f o r the remaining portioii of the vehicle. The vehicle integration was simulated by an analytical m h i c l e weight-cost sizing evaluation model.

Vehicle Weight-Cost Sizing Method A vehicle weight-sizing analysis pyocedure (ref. 26-1) was coupled with a cruise t r m s p o r t economics model (ref. 26-2). Basic input data included weight and volume coefficients, propulsion-system data, specific geometrical characteristics, and cost coefficients.

For the vehicle weight-sizing analysis, the baseline vehicle gross weight (W), reference wing area (SRW), and. t o t a l r?leZ weight t o vehicle gross weight (fuel fraction) were used. These baseline vehicle data a r e presented i n section 22. The vehicle configuration was assumed t o be geo- metrically s i a i l a r and t o have a constant take-off wing-loading f o r a l l sizes of vehicles.

Airplane procurement costs were established through use of the economics model of reference 26-2 and an economics subroutine employing supersonic trarre- port cost model techniques t o deternine. %he direct and indirect operating costs.

26- 1 The established baseline-vehicle cumulative cost estimates per unit for The labor cost t;us then factored along a learning 100 vehicles was used.

curve t o obtain labor costs f o r any required number of a i r c r a f t . Material costs were simjlarly factored along a learning c m e . ( A learning curve is an expression of t h e rate at which product?’on cost per u n i t decreases as t h e The learning curves c i t e d here are number of units produced increases.)

based on airframe industry standards (ref. 26-3). Total tooling costs were amortized over the appropriate production quantity. A summation of airframe manufacturing labor and material, avionic, and propulsion costs provided t o t a l vehicle costs for the established production quantity. One-time investment costs, including spares, f a c i l i t i e s , and production tooling required t o bring t h e system t o operational status were then added t o obtain t h e i n i t i a l invest- ment cost f o r t h e established number of operational vehicles.

In addition t o these data, payload (Wm) extreme values were bounded, as presented i n figure 26-1. A l l these constraints were put i n t o the weight- scaling synthesis model loop, i n which wing reference area i$ the primary seal- As the vehicle gross weight garameter varied, variacJions i n ing parameter.

fuel requirements t o perform the 4000-mile nautical mission resulted i n pay- load capability variations. Once the weight and sizing conditions were satis- fied for the basic mission requirements, the data were put i n t o the economics (fig. 26-i), i n which each element cost wa6 varied linearly with vehicle weight change. Then, the vehicle procurement (including anticipated spares), direct operating cost, indirect operating cost, and t o t a l system cost were com- puted i n d e t a i l f o r the specified mission. Because of structural efficiency variations between the wing concepts, the output provided variable fleet sizes and vehicle gross weights t o s a t i s f y the 205 b i l l i o n ton-mile (statute) fleet payload range reqcirement, a s vel1 &s t o t a l system cost.

Cost Model Summary The three major categories which make up t h e cost model f o r the cruise airplane are: 1. Research, Developnent, Test and Evaluatior, e - (RM!&E)

2. I n i t i a l Investment - (IV)

Total Operation Cost - (TOC)

3.

For t h i s study, however, only the l a t t e r two categories were used and a r e con- sidered t o make up the cruise airplane t o t a l system cost (TSC). !Jbus, TSC = IV + TOC.

26- 2 I n i t i a l Investment This category consists of a l l one-time investment costs required t o bring The elements comprising t h i s category are the system t o an operational status.

noted i n taule 26-1.

The major elements a r e the operational vehicles, spares, and f a c i l t t i e s . A learning factor on materials and on labor for fabrication, as discussed e a r l i e r , is taken into consideration i n determining the f l i g h t vchiclc manufacturing cost (ref, 26-3).

Total Operation Cost The costs of operating the system (both direct and indirect operating) f o r a 10-year period are included i n t h i s ce.ccgory. Both the direct operating cost (DOC) and indirect operating cost ( I O C ) are based on the Air Transport Association (ATA) method.

The ATA method, developed from reference 26-4, i s a universally recognized This method has been revised, uplated method f o r estimating operating expensesI 2nd used as a part of t h e FAA's economic model ground value f o r the U.S. Super- sonic Transport Developnent Program. The costing factors required f o r the ATA method of determining direct and indirect operating costs for various s i z e vehicles are obtained from cost analysis work described i n reference 26-2.

Direct Operating Cost. -The direct operating expenses a r e ealculated i n accord- 26-5.

ance with reference Fuel Cost: The cost of hydrogen f u e l i s a c r i t i c a l factor i n the future eco- nomic f e a s i b i l i t y of the hy-personlc transport. Reference 26-6 presents the resulte of a study made of liquid hydrogen production cost based upon projec- tion of the increased demand associated with hydrogen-fueled. a i r c r a f t . Produc- tion costs were estimated at 10 important international locations. Var-ables investigated were piane capabity, production methods, probable technological advances, and the effect of the goographical location of rav materials and energy sources.

The results of t h i s study indicated t h a t future production cost of liquid hydrogen m F b y range from 8 t o 13 cents per pund, depending on the location and quantity produced. This price includes amortization of the L J 3 2 plant cost.

For t h i s study, 11 cents per pound was selected as the cost of the liquid hydrogen f'uel.

Tndirect Operating Cost. - The U . S, Scheduled Airlines Indirect Operating

Expense Constants have been updated from the 1966 expense reported on Fom 41 t o the C i v i l Aeronautics Board (ref. 26-7). These carstants a r e used i n con- junction with the formula outlined i n reference 26-8.

The operating ex-penses composition and indirect expense subjects, con- sidered. i n t h i s research program, for t h e U.S. International Airllnes are presented i n table 26-2.

26- 3 Cost Model Urogram The various elements of the cost model computer program are presented in table 26-3 and t h e nomenclature defining the model i a shown i n table 26-4.

INTERACTION RESULTS The various s t r u c t u m l concepts z t each l e v e l of r e l i a b i l i t y were evaluated and compared using the results of t h e interaction computer program.

The segmented leading edge and rrxltiple-support corrugated heat shield con- cepts were used with each structure. The r e s u l t s include cruise vehicle weight and geometry data, as w e l l as vehicle procurement, direct and indirect operating costs, and t o t a l system costs. Data were obtained for a range of v e h i c h , payload, and f l e e t sizes t o meet the basic mission-mylad-range requirements of 205 b i l l i o n ton-miles (statute) so t h a t the minimum t o t a l system cost f o r each concept could be defined.

Results are given i n tables 26-5 and 26-6 i n dollars and i n cents per

ton-mile, respectively, f o r t h e minimum t o t a l system cost vehicle6 . These

tables indicate t h a t t h e semimonocoque spanwise beaded-skin concept is the minimum TSC w i n g structure. The spanwise tubular concept is the next Zowest cost cmcept. These tables a l s o show t h a t t h e minimum TSC is about $74.7 b i l l i o n dollars (36.4 cents per ton-mile) f o r t h e fleet requimment specified and t h a t the f l e e t procurement cost a r e $5.7 billLon or $9.35 b i l l i o n w i t h The tables also show a s i g n i f i c a r t cost aifference of $6 b i l l i o n spares.

( 3 cents per ton-mile) between the minimum cost and next. lowest cost primary structure. In addition, improved r e l i a b i l i t y from low t o nominal o r nornrml t o high for any of the concexks adds approximately $5 billi.on t o t h e TSC, except f o r the honeycomb sandwich low- to-nominal r e l i a b i l i t y , which is about $3 b i l l i o n . The differences i n roughness drag and i n i t i a l cost between con- cepts have i n s u f f i c i t A e f f e c t on t o t a l system cost t o change the e f f e c t of weight differences. One exception i s t h a t at high levels of r e l i a b i l i t y , honeycomb, even though it i s more costly t o fabricate than the next heavier concept, o f f e r s lower TSC; consequently t h e i r ratings change with r e l i a b i l i t y level.

A plot of minimum TSC ( i n terms of cents per ton-mile) as it varies with wing unit weig?.? fox the optimum-bize vehicle and t h e corresponding baseline-size vehic L~ f o r t h e various structusal concepts ( a t nominal f a c t o r of safety) is given i n figure 26-2. The waffle concept costs are large be- cause at t h e waffle-concept weight, the vehicle has l i t t l e pyload. Conse- quently lo23 vehicles (see t a b l e 26-5A) instead of 129 f o r t h e minimum-weight beaded-skin concept axe required t o perform t h e f l e e t mission requirements.

Figure 26-2 shows t h e effect of increasing unit wing weight, which i f e x t r a - polated t o about 12.0 lb/Pb2, would show %he TfjC approaching infinity, since at t h i s weight the payload is zero.

26- 4 Baseline-vehicle-size wing weic9ts are shown in addition t o t h e optitrum-size vehicle &+.a because the unit wing weights f o r t h e baseline vehicle are comparable t o one another, whereas the optimtun-si.e vehicle unBL weights vary as a function of vchicle size. This consistency f o r baseline- estimates t o be made of how other ?on- size vehicle wing unit weights enables c e @ ~ cal.cuhted for the baseline-size vehicle, such as those dropped out by intermediate screening, compare witk t h e listed concepts. For instance, t h e semimonoco ue spanwise trapezoidal corrugation concept wing avey5ge weight is

7.45 l b / f t 8 (see section 13), which from figure 28-2 indicates a weigh% and 8

TSC t h a t a r e greater than all but the waffle concept.

A plot of T j C ( i n terms of cents/ton-mile) as it varies with vehicle s i z e (expressed as gross takeoff weight) is given i n figure 26-3 f o r t h e

d5.fferent s t r u c t u r a l ccncepts . The minimum-cost bead.ed panel concept permits

a vehicle length variation of 350 t o 488 f t o r j expressed as a v a r h t i o n 7f from 620 000 t o 1 200 000 poumds, at l e s s cost than the nexblowest cost. t t i ~ u l a r wing structure yehicle. Moreover, t h e order of structure selection remias unchanged regardless of vehicle s i z e for the range given i n figure 26-3.

Total system cost, payload, and f l e e t s i z e variation with vehicle size f o r low, nominal, and high levels of r e l i a b i l i t y (factor of safety) are pre- Because of sented i n figures 26-4 and 26-5 f o r the monocoque xtffle concept.

large wing weights and resulting small payload capability, t h e monocoque waffle concept requires large fleets t o accomplish t h e basic mission, as shown For t h e monocoque honeycomb concept shown i n i n figures 26-4 and 26-5.

figures 26-6 and 26-7, t h e variation is cost w i t h vehicle s i z e and f o r the USO, Tor three levels of r e l i a b i l i t y t h e variation is sml1 (I.ess t'ne k55).

the high l e v e l of r e l i a b i l i t y , the system cost is less than the cost of the vehicle with the semimonocoque tubular wing.

For t h e semimonocoque tubular concept, the cost mriance is approximately *8 percent for t h e minimum-cost vehicles for t h e m r i o u s levels of r e l i a b i l i t y , as indicated i n figure 26-8. Fleet s i z e varies from 132 t o 166 between the low For t h e minimum cost and high l e v e l of r e l i a b i l i t y , as shown i n figure 26-9.

low and high levels of r e l i a b i l i t y 18 approxi- system, cost variation between mately f 10 percent of the nominal level, 8 s indicttted i n figure 26-10 for t h e beaded concept. The fleet s i z e varies between 115 t o 149 f o r the low and high l e v e l of r e l i a b i l i t y , with t h e nomind being 129 f o r t h e nominal 882 621-p0~nd vehicle of the beaded concept (ftgure 26-11).

The data fo;. t h e semimonocoque chordwise concept are given in. figures A greater spread i n cost and f l e e t s i z e results, as shown.

26-12 aad 26-13.

Fleet s i z e varies from 168 t o 244, respectively, f o r t h e low and high level For t h e s t a t i c a l l y determinate concept, t h e cost of re!.iability designs .

variations between low and high l e v e l r e l i a b i l i t i e s vehicles are similar t o t h e minimum-cost vehicle, semitrtonoLoque spmwise beaded, resulting i n a -1;lO The spread i n percent variation from t h e nominal, as siiown In figure 26-14.

fleet s i z e for t h e minimum-cost vehicle is between 353 t o 1 9 9 with t h e nominal being 3.75 vehicles (figure 26 -15) 25- 5 &s eline Vehicle (Gross Takeoff Weight = 550 000 Pounds) A group weight statement for t h e 550 000-pound gross weight vehicle of each cowept is presented i n t h b l e 26-7. These vehicles s a t i s f y t h e spec ff led ntission-payload-range and f u e l fraction requirements f o r t h e nominal -level of r e l i a b i l i t y . The results indicate a tradeoff between wing weight and p y l o a d , which i n t u r n affects t h e nm-ber of operatzonal vehicles required t o perform the myload-range schedule. The structure and payload mass fractions vary from the init5a.all.y azsigned values, giver i n table 26-8. The increase In t h e structure mass fraction is attributed t o t h e increase i n wing unit weights f o r the various structure concepts evaluated. It is Qoted t h a t t h e semimono- beaded concept coque spanwise is t h e only concept witn a ??load mass fraction equal t o the assigned value of 0.10. Both senimonordgue tubular and monocoque honeycomb concepts have payload mass fractions of 0 -09, whereas monocoque waffle has only 0.02 payload mass fraction.

A sumnary of vehicle geometry data as w e l l as pertir;ent design p a r a m - eters a r e shown i n t a b l e 26-9. O f significance are the wing weights (table 26-7) which when divided by t h e t o t a l wing area results Zn t h e nominal w i n s unit weights used f o r concept comparison. For the s t a t k a l l y determinate concept, t h e fuselage weight increase i s included! w i t h the wing weight t o obtain an effective wing unit weight, so that t h e wing design concepts can be compred on a common basis. Table 26-9 shows t h a t the semimonocoque, spanwise beaded skin concept has t h e least w e i g h t , with the next least weight being the semimonocoque, spnwise tubular concept (5.4 percent heavier).

Cost results f o r t h e operational vehicles are presented i n t a b l e 26-10, including ini-iial investment costs for the specified number of vehicles re- qu5red t o perform t h e basic payload-range schedule. Total operational costs (includes direct and indirect operating costs), and t o t a l system costs f o r each concept are sham. The individual f l i g h t vehicle costs, regardless of concept, do not vary apprxiably ($30.9 million t o $32.2 million). The f l e e t cost (OV - o p r a t i o n a l vehicles) varies d i r e c t l y with t h e number of vehicles required t o perform the specified payload-range scbedirle. Since unit vehicle costs do not vary appreciably, t h e pr',mry i-lfluence on operational-vehicle and initial-investment costs is t h e f l e e t s i z e rc.q'Jhewzz Similarly, fleet size has t h e major irapact on t h e t o t a l operatronal cost (TOG), which is t h e primary factor influencing TSC. The t o t a l operational costs are Etpproximateljr 88 percent of t h e 'cotal system costs, as indicated. The importance of wei@;h% i s indicated, f o r the design of the vehicles, and lesser inflasnce of i n i t i a l - cost. For t h e given gross weight (550 000 l b ) , an increase i n structure weight This decrease directJy a f f e c t s t h e decreases t h e payload carrying c a p b i l i t y .

Since, in general, f l e e t size required t o perform the specified mission.

operating costs (DOC plus ICC) are nearly t h e same for all concepts (except monocoque wzffle) regardless of f l e e t size, t h e t o t a l systetn cost varies direct!g with wing weight.

26- 6 TEtble 26 -10 indicates t h a t t h e semimonocoque, spanwise-beaded skin concept is the lawest TSC wing structure.

The semimonoccque, spanwise tubular concept is t h e next lowest cost concept, with the monocoque honeycomb being the t h i r d lowest cost. The cost increase of t h e tubular and honeycomb con-

c ept s over the minimum-cost beaded concept , which is approximately $86.3

billion, is 6.9 percent and 9.0 percent, respectively.

For t h e tubular con- cept, t h i s increase almost equals the procurement cost f o r the beaded concept and t h e increase f o r t h e honeycomb concept exceeds it.

Minimum Total bzstem Cost Vehicles (Gross Takeoff Weight = Variable) A group weight statement for t h e vehicle sized t o achieve minimum The gross takeoff weights vary systein cost is presented i n t a b l e 2 6 - 1 . 1 .

between 562 904 lb t o 882 621 lb f o r t h e minimum cost systems. The trend f o r t.thicles w i t h larger payloads and consequently srnaller fleet s i z e s i s noted.

D e resulting mass fraction f o r the various components is given in table 26-12, which indicates a structure-payload variation. The heavier w i n g weights result i n large structure mss fraction w i t h the decrease i n p y l o a d fraction. The decrease i n propulsion as well as equzpment mss fractions - r e attributed t o constants used i n the computer program.

Although t h e m i n engine and pro- pellant distribution system a r e sized and weighted t o s a t i s f y t h e thrust requirements for change i n variable gross weight, t h e air induction system is as-,ned constant (44 689 lb). Thus, w i t h increase i n vehicle size, the pro- pu:sFon mss fraction tends t o decrease.

This assumption was made t o avoid an air induction system design exercise, which was considered unwarranted f o r t h i s study effort.

Pertient geometry and des@ parameters f o r the optimum-sized vehicles are shown i n t a b l e 26-13.

(minimum cost systms) The resulting wing unit weights show a 20 percent increase over t h e baseline vehicle f o r the semi- monocoque spanwise beaded concept.

The cost results f o r each vehicle are given in table 26-14. %e air- frame labor, mterial, and mnufacturing costs are presented, i n addition t o avionics and propulsion costs. The individual f l i g h t vehicle costs, regard- l e s s of concept, do not vary appreciably ($31.4 million t o $43.8 million), a trend also noted on t h e baseline vehicles. I t is n o k d that t h e vehicle unit costs f o r the honeycomb concept and s e monocoque, spanwise beaded concepts

F

an: approximtely the same ($43.8 x 10 ). However, t h e fleet size require- ments due t o t h e payload capability of each concept increases t o t a l cost. over Therefore, the primary the minimum cost system by approximately 11 percent.

factor influencing cost is t h e f l e e t s i z e requirement, which is dictated by The operational costs for the sized the wing-wei~tlpaylond-weight tradeoff.

vehicles a r e approximtely 88 percent of t h e t o t a l system cost.

26-7 Constant Weight Vehicles (Gross Takeoff Weight = 882 621 Pounds) The vehicle weight corresponding t o t h e vehicle sized for minimum t o t a l s y s t m cost ;semimonocoque spanwise beaded-skin concept) was used for.

final Compaiicsr, of t h e structure concepts. Figure 26-16 presents t h e t o t a l system cost. (dollars) variation with vehicle s i z e ( i n terms of gross takeoff weight ) f o r each concept.

Several approaches were taken i n comparing concepts, including consideration of t h e following: a. Baseline vehicle (gross takeoff weight = 550 000 lb) b. Optimum-size vehicle, minimum total system cost vehicle (GTOW = variable) c.

Constant gross weight vehicle (GTOU = 882 621 l b ) d. Constant payload-fleet s i z e vehicles (GTOF? = variable) Constant gross weight vehicles (GMW = 882 621 l b ) were selected f o r comparison of t h e concepts since the vehicles are of constant s i z e (as i n t h e case of t h e baseline 550 000-lb vehicle) but a l s o since t'ne t o t a l system costs a r e closer t o t h e minimum for each concept. Cross plots 02 available data, such as shown i n figure 26-17 of total system cost variation with fleet size, were used t o obtain t h e f l e e t s i z e required f o r each of the vehicles having a constant gross weight. The w i n g weight f o r each conczpt was obtained through use of t h e wing weight equation (ref. section 22)b Table 26-15 presents the r e s u l t i w wlng t o t a l weights and wing unit weights, as w e l l as f l e e t s i z e re- quirements and t o t a l system costs. The t o t a l system cost variation with the nominal wing unit weights f o r the constant gross weight vehicle (GTOW = 852 621 lb) as w e l l as t h e f l e e t s i z e requirements, are preserit.ed i n figure 26-18. The data indicate t h a t a difference i n cost between t h e w2.nimwn t o t a l system cost vehicle and t h e next least cost is $6.370 b i l l i o n . A l s o note- worthy is the trend of increasing fleet s i z e with t h e increase i n wing unit weight. As previously noted, t h e increase i n wing decreases t h e payload c a p - b i l i t y , requiring additional vehicles t o perform t h e basic mission. Since t h e major portion of t h e t o t a l system cost i s primarily due t o t h e fleet s i z e An approximate cost -weight comparison can be mde between t h e lowest increase.

weight (beaded) and the next lowest weight concept ( t u b u h ? ) - Assuming an average f l e e t size (135 vehicles) and using thc uiit w ; n & ~izL&;s m d corre- spcmding t o t a l system costs shown i n figures 26-18, t h e approximate cost- weight relationship can be determined from t h e following expression: 26- 8 where ATsC = Total system cost d i f f e r e n t i a l = $6.37 x 109 w = U n i t wing weight d i f f e r e n t i a l = 0.41 lb/f’t2 S = W i n g p3anform area = 16 206 ft2 Fleet = Average f l e e t s i z e = 3-35 Thus the dollar per pound of saving by selection of t h e beaded-skin concept over t h e tubular concept i s $7000/lb of wing structure per vehicle.’ INTERACTION SUMMARY A surmnary of wing u n i t weights and percentages f o r increase i n w2r-g weight and t o t a l system cost is presented 2t-1 table 26-16 f o r t h e baseline vehicle (550 000 l b ) , minimum system cost vehicles (variable gyoss weight), and f o r t h e constant weight vehicles (882 621 1b)- Sfnce only the baseline and constant weight vehicles are f o r a constant vehicle size, the weight comparison data a r e meaningfa. For the constant weight vehicles (882 621 l b ) , the tubular concept is approximately 5.5 percent heavier than t h e beaded-skin concept, but t h e total system cost is 8.5 percent greater. The t h i r d ranking p r i m r y structure i s the honeycomb-core sandwich. Thts concept 1s 6.2 per- cent heavier and 1 0 . 8 percent more costly than the minimum weight concept.

The s t a t i c a l l y determinate, chordwise-stiffened, and waffle are more costly first three concepts. It should be noted t h a t small w e i g h t in- than the c r m s e s cause l a r g e cost increases. The weight order of cancepts, which varies by as l i t t l e as 6 percent, controls the t o t a l system cost i n t h e same order, but t o a g r a t e r degree.

26- 9 26 -1 Jones, R . : Weight Synthesis and Sensitivity Programs, Lockheed-California Comparly; LR 25205, 1967.

26 -2 Morris, R. E.: A Study of Advanced Airbreath- bunch Vehicles with Cruise Capability, Vol. V: Economics Model LR 21042, LockheedGalif ornia Company, February 1968, NASA CR 73198.

Brewer, Glenn M.: The Learning Curve Fn the Airframe Industry, 2 f ; -3 A i r Force I n s t i t u t e of Technology, Mright -Patterson AFB, Ohio, August 1965 26 -4 Mentzer and Nourse: Some Economic Aspects of Transport Airplanes, United Airlines, Journal of Aeronautical Science, A p r i l - h y 1940.

Federal Aviation Agency: Supersonic Transport Economic Model 26 -5 Ground Rules, Washington D .C . , revised September 1965.

26-6 Wilcox, 1 9 . E.; and Smith, C . L.: Future Cost of Liquid Hydrogen For Use As An Aircraft Fuel, April 22, 1968. m-68-3 Working Paper N A S A OAF@, Mission Analysis Division, Moffett Field, California.

26 -7 Indirect Operating Expense Constants f o r 1966, CMR 1010 Commercial b r k e t Research, A3ril 1968. Commercial a r k e t mgineering and Research Division, Lockheed-California Company, 26 -8 Nethod of Estimating Airline Indirect Expenst . Boeing/Lockh.eed J o i n t Report, August 30, 1964.

26- 10 TABJX 26-1 SUbMARY O F INITIAL IWLESTMENT COST ELEMENTS Element Description

1 . Operational Vehicles - OV

Operational f l i g h t vehicles a. Spares - O S Replacetwnt during operational period

F a c i l i t i e s - F A C

Complete launch f a c i l i t y and H2 plant 3.

4. Production lhgineering - PE

Preliminary design conversion t o product ion

r . - F T

Production Tooling Hard tooling

6 . Sustaining Ehgineering - SE Engineering support of operations

Sustaining fooling - ST

7. Changes t o tooling due t o design

8. Aerospace Ground Equipment - AGED Additional equipment f o r operations

Technical Data - TDO Production vehicle data

9.

10. f i s c ellaneous Equipment - M E Stock items, including trucks and

off i c e equipment

I n i t i a l Stocks - E T 30-day supply of f u e l and misc items

11.

12. I n i t i a l Training - IT Opemtion, mintenance, and personnel

training equipment

I n i t i a l Transportation - TRI Personnel and hardware transportat ion

Initial Investment - I V = s u m of items (I) through (13)

26- 11 TABU 26-2 INDIRECT OPERATING EXPEXGE CONSTANTS r It e m Description No.

--

Ground Property and Equipment

Expense - System

0 kintenance 0 kintenance Burden 0 Depreciation 2 Ground Property and Equipment 0 Wintenance 0 Wintenance Burden 0 Depreciation Ianding Fees Aircraft Servicing 3erv i c e Adrninist rat ion Aircraft Control and Communication 4 Zabin Attendant Expense ~ Food and Beverage Expense TABLE 26-2. Concluded INDIRECT OF'ERATaJG EXPENSE CONSTANTS . .

It e m Description NO --I- 6 hssenger Handling Reservation and Sales 7 Baggage and Cargo Sandling

r-'i-

~~ ~ ~

8 Passenger Service - Other Expense

Passenger Agency Commission Passenger Advertising and Publicity - Freight Commission Frei&'c Advertising and Publicity General and Adtninistmtive Expense 26-13

m m 26-3

COST MODEL C O M P W R PROGRAM

m E A R C H , DEVEIX>PMENT, TEST, AND EVALUATION - FUYI'E

AIR??RAME DESIGN AND DEVEXOPMENT ENGJXEERIIJG - ADDE

-6

CONCEPT FORlvIlTIATION - C F = 2080 * EHR * PSEF * NYF * NCF * 10

-6

CONl'RCICT DEFINITION - CD = 2080 * M R * NED * myD * NCD * 10

AIRFRAME DESIGN - AFD = (3.82 * (100 * AC) 0 . 9 1 ) * l o ' *

-6

MISC SUBSYSTEM DESIGN - E D = CPPD * W U B * 10

SUF'POFtT EQUIPMENT DESIGN - SE 0.047 * WEMPT * * 0.59

SYSTEM INTEGRATION - S I = 0.084 * W E M P T * * 0.48

FLIGH!P TEST OPERATIONS -FTO = (985 * WG * * 0,8 * N'P * * 1 . 3 ) * 10 -6

ADDE = C F f CD + AD + S D + SE + S I f F'IQ

AVIONICS DEVEIOPMENT - m = 550 * (WGNAV + WGOMM) * * (-0.24)

PROPULSION DEVEMPMENT - PD = PCF * TSIE * * 0.744 * M E * * 0 . 1 7

DEVEI0PI"T SUPPORT - Ds

GROUND TEST m r c m - GTV = NG * AMFC

PROTOTYPE VEHICLE: - P V = N P * FV

PROTOTYPE SPARES - ps = 0.25 pv

TOOLING AND SPECIAL TEXT EQUIPMENT - TST = 0 . 1 0 * (W!3MFT) * * 06

-6

FLIGHT TEST FUEL - 5°F = CH2(NFT) WFTOT *. 10

FLIGHT Y m T MAINTENANCE - FTM = 1 . 5 * VM * NIiT * loa6

GENEEUXG SUPFORT - G S = 0.3 (PTO + PTF + F T M )

m A N C E TRAINEW - lvar = M ! (INpvzl)

T A B U 26-3. Contlnued COST MODEL COMPUTER PROGRAM

OF’EFUlTIONAL TRCIINEFE - OT = T I ! (INm)

AEROSPACE GROUND E Q U I P = - AGEP = 0 . 1 5 * PV

TECHNICAL DATA - TDP = 0.02 * W

DS = GTV + GTS + P V + F S + TST + FTF+ FTM+ GS + M T -I- OT + AGEP + TDP

lU@E = ADDE + AD + PD + I 3 3

IJ!JITzcIL IIYVESTME??T - IV

OPEF4ATIONAL VEHICLES - OV

FLIGHT KEHICIJS - FV

AIRFRAME MANUFACTUNIYG -AMFG

LClBOR IJURNDJG CURVE - LlC

L E = (Nv) * * d”322

AIRFRclME LABOR - AL

FUSELAGE - FUSL = (WBODY + WDR) * CFUSL * LIC * lod

FIE3 - F I N L = (W’l?AIL) * CFIXL * L E *

WING - WINGL

MAIN WING STRUCTUE3 - A - MWLA

-6

MWLA = K5 * (WING) * C M W L A * L E * 10

MAIN WING STRUCTURE - B - MWLB

-6

WIB = K6 * (WING) * CMWLB * LIE * 10

VAIX WING STRUCTURE - C M W U

-6

* (WING) * CWIC * L E * 10

M W L C =

W I N G EDGES - LEL

LFIL = K4 * (WING) * CIXL * LLC * lod

26- 15 TABU3 26-3. Continued COST M3DEL C O M P U T l 3 R P R O G R A M ELEVONS - E L

EL = K 3 * ( W I N G ) * CEL * LLC * SOg6

WINGL = mL + EL +

+ M W l B +

INLET INLL = WAIND x- CIXLL * LIX: * SOm6

-5

NOSE C A P - NCL = ICBSS * CMCL * Lu: * 10

I N S U T I O N - IXSLL = IC2 JC (WTFS) * CINSL -IC L E Jc

HEATSHIEIDS -HTSL

MATERTAIs JXARNING CURW -Mu: = (NV) * * -0,074

PUSEMGE - FWM = {WBODY + WDR) * CFUSM * MIC * LO4

FINS - FINM = (WAIL) * CFINM * MU: * io-6

WING - WIIVGM

MAIN WJXG STRUCTURE A - M W M A

-6

MWMA = K? *WING * CMW?M * Mu: + + 10

26- 16 TABLE 26-3. C o n t i n u e d COST MODEL COMPUTER PROGRAM

ELEVONS - W

-6 EM = K3 * WWTNG * C E M * Mu:

WINGM = L?3M -I- EM + !wMA -t- Mk:yiB I-

-6

II\J’LE!T - INLM = WAIND * CITWM*ICLC * 10

-6

N E E CAP - NCM KBSS * CNCM * Mu: * 10

I N S W I O N - INSLM = E * WTFS * CINSM .)t-

* 1 , - 6

-6

HEATSHIEIS -H’TSM = K1 * W T E * CHTSM * MIX * 10

AM = F l B M -I- FlpJM -I- WINGM + 3NLM -+ NCM + INSIM I- M 5 M

MISCELLANEOUS SUBSYSTEE - Ms

FLIGHT CONTROIS - RC = WFC * (CFCL * LE + CFCM * M E ) * LO+

-6

INSTRUMENTS - JJYS’IIC = WINST * (CINTL * L E + Sam * MIX:) * LC

FURNISHINGS AND EQUIP - FUEQ,C = ()EORCE + WEQTXP

WCOMM) *

(CFEQL * LE + CEW * fim) * 10“’

26-17 TABU3 26-3. Continued COST MODEL COMPUTER PROGRAM

AVIONICS - AV

W A V = WGNAV + ICOM

AVIONCS PROCURENEW 1 - AVP = CPAV + , W A V *% lod * blIC

AVIONICS INSTALLATIOJS - AVI = ICPAV W A V * lo4 * (NP + NV)

** - . 3 2 2

AV = AVP + AVI: n T = M G + AV + PROP

- N T '

N i m R OF OPERATIONAL VEHICLES N V = NV (IXF'UI!)

m r r ~ 4 ~ S P ~ - 10s = 0.25 * cv

R E ~ I S I D E ~ spARFs - ROS = 0.25 * 10s

OS = 10s + ROS

FACILITIES - YAC = FAC (INPUT)

PRODETION ENGINEERII~G - m = 0.25 * (CF + CD + m ~ )

PRODUCTION WOLLNG - FT = 0.05 * WEMFT -* * 0.75

SWTnJNmG EL\TGjJ!E8RING - S W = O.O'j05 * (ADDE - CF .... CD)

* KL * NV * $e C.848

SUSTAINDG TOOLING - ST =

AE3OSPACE GROUK' EQUIPMETJT - AGE0 = 0-15 * OV

26- 18 T A B m 26-3. Continued COST W D E L COMPtFL%R F ' R O G W

TECHNICAL DATA - TDO = 0 . 1 0 * TDP

-6

MISCELLANEOUS EQ,lJIPMWI' - M E C = 500 * NPER * 10

-6

lKl"IAL STOCKS - IST = 0.083 Je VM + 100 * NPER * 10

I!'?ITIAL !I!RAINING - I T = 0.10 * OT * NPL

INITIAL !JXAIVSPOIi!CATION - TRI = 0.005 (OV + OS + AGEO -+ EIEC + E T )

IV = OV + OS + FAC + TE + PT + SEC + ST + AGEO i - TDO + MEE + IST + I T + TRI

DIRECT OPERATING COST - DOC

FLIGID TlME - TF1 = DIST * TFU/(DJST + TFU * WIND)

TOTAL F L I G H T TlME - T7 = G R T ! X G T + 'IF1

.-&9UJlT OF FUEL -FcTEL = WETOT

PCET = [Cl * T7 + C 2 (FUE3;) + C 3 Je TF1 + C 4 ] * (1. + PDOE)

-6

PROT = C2 * WEZ Jc 1 0

INSUWNCE - &INS

--

PA = M@'G AV = PAW.)

QINS = (PA . + PROP + PAW + PROTI * RCON/(TVL/TFi) * (1. + PDOE) *

TOTAL F L I G H T TIME PER DAY - TFTD = TFl * NFD

NUMBER OF AVAIWIZ FLIGHT DAYS - NAFD = TVJ_/TETD

NUMBER O F F L I G m YEARS - NFY = N A F D / ~ ~ ~ . O

NUMBER O F F W G H S PER ?EXR - NFPY = JJFD * 365.0

DEPIIECWIOK PERIOD - TA = NFY

EllGZNE DEPHEZIATION PERIOD - TE = IJFY

AVIONICS D E m C W I O N PERIOD e - TAV E NFY

DEPmCUTZN - WEP

26-19 \ TDLE 26-3. Concluded COST MIDEL COMPU'JXR PROGRAM

(.. * (1. -RA + CSF) PROP * (1. -Re + CSEFT) PAVO * (1. -RAV + CAW))

+ +

QDEP = TA TE TAV

* 10+6

DOC = PCCEjT + BMAN * [C5 * T7 i- ~6 * (T7-GRNlX) + C7] + &INS + &DEP/NFPY

V M = BMAN * [CS * T7 + ~6 * (T7-GFmYr) + c7]

?WIIXECT OPERATING COST - E3llXlC

&'COST = E l l * I C 5 * T7 + ~6 * (T7 - GFtNDIj) + C 7 ] + El2 + 13 * WG + E14 *

* T7 + DLST * [E20 * S E P L T S + E21 * PIMpx/!RlN~ * (1. + PIOE)

FUBBER OF PASSENGER - PAS = S w * AL;F

PA;SSENGER BmCK HOUR - TPBH = PAS * T7

w m = PAS * C K l * cK2

L4RGO MII3 - TCMI = W A R T * DLST

r

Z'YX = [EL5 * ~8 * TPBH + E 15A * ~8 * PAS + ~ 1 6 * P A S + (E17 ++ CK1 * CK2 *

PAS -F E I ~ A * WPAEP)/TON + ~ 1 8 * T P M I + ~ i g * c9 * TCMI/TON *

(1. + PIOE)]

ENDOC = [ ( 1 . + E22) * (FCOST + ZYX) + E22A * (Q S + PCOST)] -

B M A N * [ C 5 * T 7 + C 6 * ( T 7 - - G R K E C ) + C 7

NFPY * r : , Y * (DOC + ENDOC)

TOTAL OPERATIONS COST - TOC

TOTAL SYSTEM COST - TSC = TOC -t IiDlCE + I V

26- 20 T A B m 25-4 NOMENCUTUI@ FOR COST MODEL AC inlet capture area AD avionics development cost airframe desiqn cost AFD

AGE0 aeroswct? . xud equipment cost

aerospace ground equipment cost AGEP AL airframe labor cost AIJ? average passenger load factor airframe material cost minimum cargo weight airframe manufacturing cost AMFG minimum payload weight APAY available seat mile ASM BLKF t o t a l amount of fuel maintenance burden fact or B M A N DOC block hour factor C1 c2 DOC fuel (lb) factor DOC f l i g h t hour factor c3 c4 DOC departure factor

DML - block hour factor

c5

DML - f l i g h t time factor

c6

DML - departure factor

c7 passenger block hour weighting r a t i o c8 r a t i o of freight t o t o t a l cargo c9 26- 21 TABLE 26-k. Continued

I1

NOMENCIATURE FOR COST KIDEL CAW? value for spare avionics factor CD contract definition cost C F concept formulation cost CH2 cost of hydrogen ($/lb) CKl volume of baggage per pssenger CK2 density of baggage and cargo C PAV cost per pound of avionics CPIG cost per pound of landing gear (labor) , PIGM cost per pound of landing gear (material)

F i

CPM3 cost per pound of miscellaneous subsystems cost per pound of development of miscellaneous subsystem CPPD C PROF value f o r spare propellants factor CPT cost per t i r e value for spare engine factor CSEFl CSF value f o r spare p r t s factor DIST f l i g h t distance E l l IOD - d i r e c t maintenance labor

E l 2 IOD - a i r c r a f t departures

IOD -departure times maximum landing weight E13

E l k IOC - cabin attendant block hours

IN - revenue passenger block hours (food)

S l 5 A IOC - revenue passenger carried (food)

~ 1 6 IOC - revenue passenger carried (servicing and sales)

IOC - passenger baggage caxxied

E17 26-22 TABLE 26-4. Conthmed NOMENCLATURE FUR C O S T KIDEL

I O C - cargo carried

IOC 'revenue passenger miles

IOC - revenue freight ton miles

E20 IOC - available seat miles

E 2 1 I O C - available ton m i l e s

E22 IOC - general and administrative - indirect

E22A I o ( : - geaeraland administrative -direct

M R engineering hourly rate

A m o c indirect operating cost

f i n a l assembly of structure c o s t FA FAC f a c i l i t i e s cost amount of fuel trapped i n the vehicle f l i g h t t e s t f u e l cost f l i g h t test mintaiance cost FTM f li&t t e s t operations FTO amount; of f u e l fli&t vehicle cost Fv ground t a x i time (hr)

s m

support cost Gs general ground t e s t spares cost GTS ground test vehicles cost GTV installation cost per pound of avionics ICPAV inl'iial spares cost inithl stocks cost IST 26-23 TABLE 26-4. Continued NOMENCLATURE FOR COST lvlODEL IT i n i t i a l training cost IL; landing gear cost *w engine maximum operational &ch number M f a miscellaneous equipment cost E miscellaneous subsystem cost

mo miscellaneous subsystem design coat

maintenance trainers cost

m

n m number of available f l i g h t days

i K D number of contractors doing contract definitions NCF n W e r of contractors doing concept formulation IJE number of modules I r n number of engineers on contract definition TIEF number of engineers on concept formulation Nl?D n-amber of f l ights per day WPY number of f l i g h t s per year NFT number of f l i g h t t e s t number of f l i g h t years NTY NG number of ground test vehicles number of landing gear units N I G U nwber of miscellaneous subsystems units N E U number of prototype vehicles N P number of propulsion engines NPE number of t o t a l personnel NPER number of' p i l o t s NPL 26- 24 26-4. Continued NOMENCLATURE FOR COST MIDEL N T number of tires per landing gear unit Y"3J number of operational vehicles NYD amber of years for engineering contract definition

rm number of yams f o r engineering concept formulation

WTST Plight distance 0 1 1 E P f u e l tankage fullness ratio OT opemtional t r a i n e r s cost OV operational vehicle cost JE opera-Ling weight empty PA airframe cost PAS number of passengers PAYO t o t a l avionics cost PCF propulsion development cost factor DOC less insurance and depreciation PCOST PD propulsion development cost PDOE percent change i n DOE F % product ion engineering cost I?Ec t o t a l engkne cost per aircraft PIOE percent change b IOE PLBM payload capacity weight PLMi mximum payload PROP1 propulsion i n s t a l k t i o n cost PROPP propulsion procurement cost t;&e,l cost of propellant PROT 26- 25 TABU3 26-4. Continued NOMENCLc1'JIITRE FOR COST IWDEL ps prototype spares cost F T production tooling cost

J ? v prototype vehicles cost

quality control cost QC insumnce cost

ams

airframe residual value RA RAV avionics residual value insurance rate RCON engine residual value ,E IiPRO propellant residual value SE support eqd pment design cost t o t a l number of seats SEW23 SEC sustaining engineering cost systems integration cost SI ST sustaining tooling cost sea-level s t a t i c thrust 'I!

t o t a l f l i g h t time (including t a x i s time) airframe depreciation period avionics depreciation period rate of t a x i fuel ( l b l h r ) cost per cargo mile number of operational hours per day production vehicle data cost supporting technical data cost 26- 26 TABLE 26-4. Continued N 0 M E N C I . A ~ FOR COST MODEL TE engines depreciation period number of cabin attendants

mm

t i m e t o f l y given distance with wind factor W1 t o t a l flight time per day n ? J ! D t i m e t o f l y given distance V U TL scheduling loss i n hours t o t a l operating cost WC t o t a l operating time i n hours TOPER , " O N pounds per ton cost per passenger block hour TPBH TPMI cost per passenger m i l e propellant depreciation period TP'RO E C t o t a l system cost sea -level t h r u s t per engine !ISLE tooling and special test equipment cost TST TT minimum turnaround time t o t a l vehicle l i f e TVL percentage of flight f u e l f o r reserve mop u t i l i z a t i o n f a c t o r U vehicle m i n t enance cost VM weight of avionics equipment WAV vehicle empty weight WE engine weight 'WEN gross fitage weight WG TABLE Z6-4. Concluded N O M E N C L A W FOR COST MODEL W G R c x j S mximum gross take-off weight WIND wind f a c t o r W L A N D a i r c r a f t weight for airport fees W U B weight of miscellaneous subsystems WPART weight of cargo W P A S S passenger weight WST structure weight 26- 28 TABLE ?6-5 COST BR'JKDOWN IN DOLLARS FOR EPEH PRIMARY STRUCTUTE AT EACH LEVEL OF REIIABILITY . - - --- \VI IIK Vrliic*lc Lrvcl S t n i r t u r r llllt t PaylowJ, I r n i:t h , o f cn I1 ('e pt wc.i [:!I t , lb I1 8 sellnhlllty lt,/ft2 --A. . - 7 . 1 m S r m l mcinoctquc. 923 970 94 942 I.0w 7.454 H5 06R .sp:itiniHe Nom i tin I nx2 621 I wfrtl A36 R24 4 0 7 I . 784 74 058 High 7.497 Srniitnnnocnqire I.0w A74 2R7 83 324 #1unwlae Nom 1 nnl 8.10 870 7.718 75 6 l H tu tni I:] r High 791 110 7.924 6 G 349 142 .14. t l N 7.598 77.47R Mnnncctque Ilow 842 818

Iwnr-yc.omb- 14R i 7.748 74. I79

Nominal 835 241 42.3fiH ' 7.041 rorr lllgh 161 799 753 GR 308 Statir:d ly 7. R97(h) Low 153 43. eon 036 318 407 71 933 . dc.ternilnnte 175 41.997 8. inri 62 non Nominal 797 493 397 spanwlse I!M 40.4!)8 8.462 55 322 High 782 021 I)c adr d ..

1 i;n 40. GI 5 7.89R rii, 2nn Scm 1m oiiocoque Low 799 788 39R 37. C l i 5 379 8.251 51 (iri!)

cho rtlwlse Nominnl 728 862 709 737 375 8.596 45 O H 5 tubular High 599 236 9. no9 20 903 33i 10.432 10 748 562 904 10.B88 3 323 - -_ -. .. -.- - ..- --..

- 7

Trhl Total R r l a t l v e

I

Coqt Initial S t r u c t u r e roc, operxtlon:\l system- bhl- vehi c l e a t Investment , concept COHt.

c*lltlt. clvatem- billions blllions bi II 1 on8 bi I 1 i o n s CORt -.- - - - - - - - --- - -_-- _c_-l- i 5.204 Srmimonocoque 8.689 en. 952 tifl. G4l 5.666 9.354 Rpanwise 65. 3H0 74,742 1.00 bradcd 6.244 10.186 82.455 72.270 Scmlmonocoque 6.720 9.430 8G. 268 75. (iP8 q m n w i s e 6.113 9.994 70.95'9 50.973 1.083 tullu lnr 6.648 10.757 77- 4c3 80.219 Monocoque 6.262 10.214 69. Reo 80.015 honcycomb- 6.497 10.558 72.434 82. 993 1,XlO 6.816 core 11.008 . 76.169 8 7 . 1 7 ~ Shtlc:tl l y 6.668 53.004 21.717 10.816 74.721 85.538 drtcrm Inate 7.344 57.984 24. GOO 11.796 02.744 1.263 94.3HO spanwise 8.051 63.194 27.731 12,822 90.925 103,747 Iwcadcd

I

6.843 Scm im onocaqur!

66.012 23.734 11.052 79.746 9 0 . 7 9 ~ chordwlee 64.748 8.019 29.454 12.747 94.202 106.949 1.431 8.985 tubular 14.167 72.536 33.612 106.148 120.315

i

18.142 134.453 70.450 27.699 204,903 ?35 501 Monomque 32.177 245.120 ' 135.739 48.247 380.859 429 106 6.74; waffle 89.973 737.476 435.1S6 133.278 1,172.630 1.305.908 Y n r l u d e a epres.

blnrludes welght of fudage M y penalty.

26- 29 TABLE 26-6 COST BRF%KEOhr "'XTS PER TON-MILE FOR EACH PRIMKl?Y STRUCTURE ' EKCH LWEZ O F ' REL'WILITY - - - - -1

- -TI,,,, -

Vvhicir Structu-r at wvidlt, conwpt rdlahillty Ib ----. -- _-_____ -- _-_-.

Srm imonnruque LOW 913 970 127 7.126 95 942 115 sInnwisr Nominal m 2 621 41% 7.454 n5 O W 129 Iwntled High R36 A24 407 7.784 74 056 149 Srinlmonomque Low 8 7 . 1 2R7 410 7.497 83 324 132 .;113llWi R e Nnnilnal n m 670 408 7.716 75 6 1 R 115 tulrr Inr Jligh 66 319 166 791 110 795 7.924 M o n o c o q u e Low 842 B i n 40R 7.598 77 47n 142 honrvcomb Nomlnnl 835 241 406 7.74~ $4 179 118 core Wgh 799 753 39? 7.841 68 3Wl 161 S h t i c d lv R97(b) Low 836 31R 407 7 . 71 933 1 53 3 97 8. lH6 Nominal 797 493 62 906 175 spinwise 55 32% 199 High 762 021 3414 8.462 twcndcd 398 7. ABR 799 766 6s 2n3 164 726 862 379 4.251 51 I;@) 213 375 H. 596 709 737 45 085 244 LOW 599 236 344 9.809 20 903 526

Monocoqw waffle 1 No~yn 562 904 334 10.432 10 740 1023

529 2 t ' 3 323 3310 323 10.986 -- i------- 9 - Rr l at( VF 7hfA Initial :me. IOC, S.ructurr opprntlonal

mnccpl I investnienta j cents/ 1 cents/

corit, cost Ryatcm- centdton-ml ton-mi top-mi rrntdtnn-mf centdtor. -ml Srmimonoc aque 4-22" 21 * 09 8 . 0 9 29.18 33.41 Rpanwlse 4.5:. 22.79 9.03 31. H2 36.3R 1-00 k a d e d 4.967 24.91 10.26 35.17 40.13 Svm i mono mque 4.589 23.04 9.21 32.25 36.84 spa nwi se 4.864 24-48 10.07 34.55 39.41 1.083 tubular 26.29 5.235 11.41 37.70 42.94 M o n o c o q u e 4.971 24.14 9.83 33.97 38.94 honrycomb- 5.138 25.00 10.25 35.26 40.39 1.110 core 5.386 '26.06 11.02 37.07 42.4: Statically 5.264 28.80 ZO. 57 36.37 41.63 drtcrminnte 5.741 28.22 11.97 40.19 45.93 1.263 apanwise 6.240 30.75 13.50 ' 1 . 2 5 30.49 h-arled Semirnonomque 5.379 27.26 11.55 38.81 44.19 6.204 14.34 ctxdwiee 31.51 45.85 52.05 1.431 tubular 6.895 35.30 16.36 51.66 58.56 13.432 68.44 34.28 99.72 113.16 . ~orlocoque 23.482 119.30 6 6 . 0 6 185.36 208. R4 5.741 waffle A66 62. 358.92 211.79 570.71 635.58

-

26- 31 TABLE 26-8 Initial Monocoque Semimonocoque Statically Component Determinate Valves I Waffle Honeycomb Tubular Beaded Chordwise

(a 1

Fuel

0 . &!I-- 0.40 0.40 0.40--0.40 -- 0.40 0.40

Structure 0.27 --O. 35 0.28 0.28-- 0.27 0.30 0.31 Landing Gear -

- 0.03--0.03 0.03 --0.03--0.03 0.03 0.03

0.15 0 . 1 5

Propulsion _ _ 0 . 1 5 ~ - 0.15 0.1-5 - - i - 0.15--0.154.-

Equiprnent (b)-

0 . 0 5 - - 0.05 0.05 i O . O ~ l o . o 5 + ~ : ~ ~ 0:); Payload 0 . 1 0 4 0.02+ 0.09 0.09 0.10 a Includes residuals, reserve, inflight losses, l o i t e r , taxi, run-up, and performance propellant bIncludes equipment, crew, and design reserve 26- 32

a

h

i

26-33 26-35 TABLE 26-12 OFTtMUM-SIm ATRPLANE i?BS FRACTIONS (GTOW = VARIABLF:) I n i t i a l Monoc oque Semimonocoque S t a t i c a l l y Component

Valves I Waffle I Honeyc6mb Tubular Beaded ChorCiwise Determinate

fie1 (9 0.40-- 0.40 0.40 0.40 0.40

.ructure 0*27--0.35 o 31

0.31 -- 0.32 0.32

Landing Gear

0.03 0.03 -- 0.03 0 . 0 3 0.03 --0.03 0.03

Propuls ion 0 . 1 3 0 . 1 3 0 . 1 3 0.15--0.15 0.13 I

Squipmen-t (b) 0.04 0.05 0.04 0.05 __ 0.05 0.04

0.04 Payload

0.10 -- 0.07 - 0 . 0 8 0.10,- 0.02 0.09 0.09 --

aIncludes residuals, reserve, i n f l i g h t losses, l o i t e r , t a x i , run-up, and perf orwnce propellant bIncludes equipment, crew, and design reserve 25- 37 .. .- h E U> G u TABLE 26-15 N I O SUMMARY -STRUCTURE CONCEFT DESIGN AND COST DATA GROSS WEIGHT: 882.621 lJ3, WING AREA: 16,206 SQ

. . . I - -

i Structure Cwing WSng Weight F l e e t Total System C onc ept (Id (PSf) S i z e Cost (Dollars) .-- ...

; I-miLonocoque 1

I 5:Jmw-ise 0.007752 120 807 7.454 129 74.742 x log :: eaded

- --

i Semimonocoque S panwi se 0.008181 127 492 7.867 140 81.112 ?,,,iar -- - - * .

Monoc oque 82.853

' 1 oneyc omb 0.008236 u a 350 7.920 143

bandv ich I S t a t i c 0.11y 95.020 Determinate 0.008468 131 965 8.143(a) 164

i

; Spanwise-Beaded Semimoncc oque Chordwis e 0.009269 144 523 8.918 189 109.505 Convex -b eaded/t ubular

- .-

Monocoque

Wagf l e 0.012981 202 295 12.483 - -

45 x 4 5 O determinate concept body penalty = 0.2245 (26,300) = 5,900 1b = 8.507 Psf ~ ' = ( 1 3 1 , 965) + (5,900)/(16,26) 0 LT rl Ln

. oc!

a3 0 r- I rl cu

i

3.

cu.

OI' \o t - .

a

\o t- cu a3 .

c?

t - co ~ QJ a , ri 0 m

. T ! 7!

a3 4 A- ri

3 b-

0 Ln CR

. %

m M c h m IA (u cu M ;t c- b- co

. a

t ? c- c- t - 0 0 0 CR 0 4

. . 9

.

c h w, \o k cu % CR rl I < .

! m o\ 26- 40 u t-1 0 0 M ?

. .

L. \

cn -E!

I I II

&

cn "1

2 1

I

I

I

I

I

c1

I

I

I

I 2m

-

Key 0 Semimonocoque sponwise b o d e d

- a Semimonocoque sponwisa tubular

A Monocoquc honeycombcore sandwich u

-

- 8 S totically determinote sponwisc b o d e d

g 160 0 Semimonocoque chordwire tubular

+ s

e

-

c im i H E 100 c U >. In

-

g 80 t - 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0 9.5 10.0 10.5 Wing unit weight, Ib/ft2 Figwe 26-2. Total-system-cost for baseline and optimum-size vehicles of various wing constructions 26-42 .4 .5 .6 .7 .a .9 1 .o 1.1 1.2 Y lo6 Gross takeoff weight, GTOW, Ib Figure 26-3. Total Bystem cost for various gross weight vehicles for the candidate wing construotions 26-43 t - .

I I

I

I

I I I I I n .

I I

t

I

-?

0 0 0

0 8

In s 8

\o O

8 8 8 8 8

R 3 a 3

8 .

a

!j

H

t

I I I

I

26-48 ' 0 0

8 0

4 a cu

0 0 0 0 0 a 3 \D c u 26- 50

5 A

26- 52 l-4 \ I I

I

\ \

\

(fleet s i z e = 189) det ermlnat e (fleet s i z e = 164) 1-80

I I

I I A

7 8 9 10 NominaL wing imYt weight, gsf FLgure 26-18., .Total syetem cost variation with nominal wing unit wlight . for amstanit; gmss weigh% !vehicle # - / r h . 1‘

Section 27 STRUCTUFtAL TESTING R. S o Jusko, R. Swartz, C . E, Stuhlman, K. A. Wilhelm, R . C. Dickason, J. J. Panik, L. D. Fog& A . B . Burns, G . W. Davis, I . F. Sakata, R . E. Hubka, F . T . Bevan PRECEDING PAGE BLANK NOT FILMED.

27-1 TEST P L A N 27-1 DESCRIPTION AND FABRICATION OF PANEL ELEMENTS 27-1 Tubular Panels 27-4 Beaded Panels 27-6 Trapezoidal Corrugation F'anels 27-8 Corrugation- Skif f ened Panels 27-10 Circular-Arc Corrugation Shear Panels 27-12 Spar Cap Crippling Panels 27-13 TEST SEIPUP 27-13 Room Temperature Compression Tests 27-13 Elevated Temperature Compression Tests 27- 14 Shear Panel Tests 27-14 INSTRUMENTAI)ION 27-15 DATA ACQUTSITXON 27-15 TEST PROCEDURES 27-15 Preliminary Tests 27-55

F'ailure Tests , Room Temperature Compression Panel8

27-15 Failure Tests, Elevated T a p e r a t w e Compression Panels 27-16 Failure Tests, Shear Panels 27-16 TEST RESULTS c : 27- 16 Panel Material Tests 27-16 Panel Tests

CONTENTS (Cont . )

-@;e COMPARISON OF ANALYSIS AND TEST REXULTS 27-22 Pi'-24 Tubular Configuration Beaded Configuration 27-28 Corxugat ion- St iffened Panel Configuration 27- 32 Trapezoidal Corrugation Panel Configuration 27- 36 27- 40 Circular-Arc Corrugation Sheafi Panel Configuration Spar Cap Configuration 27- 42 REFERWCES 27-45 TABLES Page 2 7 - h . 6 27- I . Structural element test schedules Summary of panel element fabrication 27-2 27- 47 Instrumentation schedule for structural element tests 27- 4 8 2 7 - 3 Mechanical properties data for some R e d 4 1 compression 27- 49 27-4 panel materials subjected to various thermal cycles Summary of panel element test results 27- 5 27-50 27-6 Thickness measurements of corrugation-stiffened 27-51 end-closeout panel Thickness measurements of beaded end-closeout panel 27-7 27-52 27-8 27-53 Thickness measurements of tubular end-closeout panel of corrugation-stiffened 27-54 Thickness measurements 27- 9 crippling panel (room temperature) Temperature distributions for corrugation-stiffened 27- 55 27-10 crippling panel

27- 5 6 27-11 Thichess measurements of corrugation-stiffened

crippling panel (‘elevatedternpepatwe) 27-57 27-12 Thickness measurements of trapezoidal corrugation crippling panel (room temperature) 27-56 Temperature diatributions for trapezoidal corrugation 27-13 crippling panel Thickness measwements 09 trapezoial corrhgation 7-59 27-14 crippling panel (elevated temperature) 2 I -60 Thickness measurements of beaded crippling pmel 27-15 (room temperature ) 27-61

27- 1 6 Temperature distributions for beaded crippling el

27-62 Thickness measurements of beaded crippling panel 27-17 ( elevated t emperatwe ) 2’7 -18 27-63 Thickness measurements of tubuhr crippling panel (room temperature) 27-v

TABLES (Cont . )

Temperature distributions for tubular cxiip:~2!.ng 27- 19 p a l e l 27-20 Thickness measurement E of t*u.L?.? cr;lp:;ling panel 27-65 (elevated temperature j 27-66 27-21 'kickne: s measurementz of spar cap crippLing specimc ' L (3/8-inch fh.nqe) (room tempwattme) 27- 22 ThTckn~,measurements for I:wrw&tl.on-Z;tj tfened skin 27-67 compression panel (room temperakum) 27-68 Temperature disti*Lbutions for cnimgat Lx.- titif f ened 57-23 skin compression panel 27-69 27-24 Thickness measurements of corrugation-stiffened skin compression panel (elevated temperatwe) 27-70 27-25 Thickness measurements of tmpezoidal corixgation compression panel (room temperature) 27-26 Temperature distributions f o r trapezoidal 27-71 corrugation compression panel Thickness measurements of trapezoidal corrugation 27-27 27-72 compression (elevated temperatux ' , 27- 28 Thickness measurements for beaded compression panel 27-73 (room temperature) 2 ' 7-74 21-29 Thickness measurements of hbular com.pression panel (room temperature) Temperature distributions for tubular compression 27-75 27- 30 panel 27-77 Thickness measurements of tubuiar compression panel 27-31 (elevated temperatwe) 27-78 Thickness measurements of circular-arc coiqrugat20n shear panel (Ren6 41, f i l l e r wire j Thickness measurements of circular-arc cormgation 27-79 27-33 vertical. web (Hastelloy W f i l l a r wire) 27&0 Summary ct-rrehtion of .; c r u c t u m l element tests 27-34 27-81 Comparison of tubular and beaded emf iguration i n i t i a l 27-35 b~~.clcling 2; sst results with predicthns 2 7 4 IUUSTRATIONS Figure Page

Test panel assembly - tubular concept

27-1 27-82 27-2 Fornblock f o r tubular panels 27-83 27-3 Tubular panel dstails pr:ior t o assembly 27-83 Tubular panel 5x1 weld fixture ready for resistance 2 7 4 spot weld assembly 27-84 welded 27-5 Panel details being resistance spot 27-84 27-6 Tubular panel after resistance spot welding 27-85 27-7 Finger doubler extensions for tubular panel 27-85 27-8 Circular arc stiffened tubular end closeout spechen prior t o end casting and grinding 27-86 27-9 Tubular crippling panel shown with ends cast i n densite 27-86

27-10 Test panel assembly - beaded panel 27-87

27-11 Beaded panel hy&aulic forming dieblock 27-88 27-12 Beaded panel trimmed prior t o assembly 27-88 27-13 Beaded panel details i n weld fixture prior t o resistance sp& weld assembly 27-89 27-14 Beaded panel after aging, heat oxidation and install- ation of end bars 27-89 27-15 Beaded panel showing finger doubler extensions 27-90 27-16 Beaded crippling specimen with ehds cast i n densite 27-90

27-17 Test panel assembly - trapezoidal corrugation 27-91

27-18 Trapezoidal corrugation panel forming die 27-92 27-19 Trapezoidal corrugation panel details showing central section corrugation, end corrugation, aee section and fingered splices 27-92 27-20 Trapezoidal corrugation panel details i n weld fixture ready for resistance spot weld assembly 27-93 27-21 Trapezoidal corrugation panel assembly a f t e r aging and heat oxidat 3 n 27-93 27-22 Trapezoidal corrugation panel. cut t o two 8-inch lengths f o r crippling panel tw” 27-94 Figure - Page

27-23 Test panel assembly - corrugation stiffened

27-9 5 27-2L Forming die f o r closed end corrugation-stj ffened panel 27-96 27-25 Corrugation-stiffened panel details including corruga- tion, skin, fingered doublers, and end doublers 27-96 27-26 Corrugation-stiffened ?anel a f t e r aging and heat oxidat ion 27-97 27-27 Corrugation-stiffened end closure specimen cast i n dens ite 27-97

27-28 Shear panel assembly - circular arc corrugation 27-98

27-29 Corrugated shear web forming block 27-99 Shear panel details including web, caps and edge 27-30 doublers prior t o assembly 27-99 27-31 Overall view of tracer template ready f o r fixturing parts f o r shear panel T I G welding 27-100 Two-shear web panel assemblies, one before and one 27-32 after aging and heat oxidation 27-100 27-101 Beam cap crippling specimen 27-33 27-102 27-34. Typical room temperature compression test set-ups Typical elevated temperature t e s t set up f o r 30-inch 27-35 27-103 compression panel test set-ups 27-104 27-36 Typical elevated temperature compression Test set-up f o r in-plane shear panel tests 27-105 27-37 27-38 Tensile stress-strain curves f o r .016 gage Ren6 ICJ.

compression panel sheet material, longitudinal grain direction 27-106 Tensile stress-strain curves f o r .019 gage %ne' 4- 27-39 compression panel sheet material, longitudinal grain direct ion 27-107 27-4Q Tensile stress-strain curve f o r .060 gage R e d 4 l beam cap shear and crippling specimen material, 27-107 longitudinal grain direction Strain gage locations f o r corrugation stiffened 27-4l 27-108 skin end-closeout panel 27-42 Axial strains f o r corrugation-stiffened skin end 27-109 closeout panel, room temperature Panel shortening curve A L/L f o r corrugat ion-st i f f ened 27-43 27-3.14 end closeout panel, room temperature Figure 27-44 Corrugation-stiffened end closeoQt panel after failure, room temperature 27-315 Strain gage locations f o r beaded end-closeout panel 27-116 27-45 27-46 Axial strains f o r beaded end-closeout panel, room temperature 27-117 Panel shortening curve AL/L f o r beaded end-closeout 27-47 panel, room temperature 27-121 2748 Beaded end-closeout panel after failure, room temperature 27-122 27-49 Strain gage locations f o r tubular end-closeout panel 27-123 27-124 27-50 Axial strains f o r tubular end-closeout panel, room temperature 27-51 Panel shortening curve AL/L f o r tubular end-clmeout 27-128 panel, room temperature Tubular end-closeout panel after failure, room 27-52 temperature test 27-13 21-53 Strain gage locations f o r corrugation-stiffened crippling panel 27-130 Axial strains f o r corrugation-stiffened skin crippling 27-54 27-131 panel, room temperature Panel shortening curve AL/L f o r corrugat ion-st i f f ened 27-55 crippling panel, room temperature 27-135 27-56 Corrugation st iff ened crippling panel after failure, room temperature t e s t 27-136 27-57 Thermocouple locations f o r the corrugation-stiffened crippling panel 27-137 27-58 Panel shortening curve f o r corrugation-stiffened bL/L 27-138 crippling panel, 14.00° F Corrugation-stiffened skin crippling panel a f t e r 27-59 failure, 1 4 0 0 ' F 27-139 27-60 Strain gage locations f o r trapezoidal corrugation 27-140 crippling panel 27-61 Axial strains f o r trapezoidal corrugation crippling panel, room temperature 27-141 27-62 Panel shortening curve AL/L f o r trapezoidal corrugation

crippling panel , TOO^ temperature 27-145

27-63 Trapezoidal corrugation crippling panel after failure, 2 ' 7 - 1 4 6 room temperature test 27-ix Figure 27-64. Thermocouple locations f o r the trapezoidal corrugation crippling panel 27-147 Panel shortening curve AL/L for trapezoidal corruga- 27-6 5 tion crippling panel 27-148 27-66 Trapee0id.d corrugation crippling panel after failure, 14.C3G F test 27-3-49 27-67 Strain gage locations f o r beaded crippling panel 27-150 27-68 Lxhl strains for beaded crippling panel, room t emperatme 27-151 27-69 Panel shortening curve AL/L for beaded crippling . .

panel, roan temperature. * 27-154 27-70 Beaded c r i p p l h g panel after f allure, room temperature test 27-155 27-71 Thermocouple locations f o r the baaded crippling panel 27-156 27-72 Panel shortening curve AL/L f o r beaded crippling panel, 14OO0F 27-157 Beaded crippling panel after failure, U000F test 27-158 27-73 27-74. Strain gage locations f o r tubular crippling panel 27-159 27-75 Axial strains f o r tubular crippling panel, room 27-160 temperature 27-'76 Panel shortening curve AL/L f o r t-lbula-r crippling panel, room temperature 27-163 27-77 Tubular crippling panel a f t e r failure, room temperature test 27-164.

27-78 Thermocoupling locations f o r the tubular crippling panel 27-16 5 Panel shortening curve A L/L f o r tubular crippling panel, 27-79 U O O O F 27-166 27-16'7 27-80 Tvbular crippling panel a f t e r failure 1400°F test 27-81 Strain gage locations f o r the spar cap crippling 27-168 s pe cimens

27-82 Axial strains f o r spar cap crippling specimen - 3/8 inch

27-169 flange, roam temperature 27-83 Spar cap shortcniug curveAL/L f o r 3/8 inch flange 27-170 specimen, room temperature 27-84. Spar cap crippling spechen (3/8 inch flange) a f t e r failure, room temperature t e s t 27-171

27-85 Strain gage locations f o r corrugation - stiffened

27-172 skin compression panel 27-x Figure Axial s t r a i n s f o r corrugation stiffened skin 27-86 compression panel, room temperature 27-173 27-87 Panel shortening curve bL/L f o r corrugation-stiffened skin compression panel, room temperature 27-181 27-88 Normal deflection curve f o r corrugation-stiffened skin compression panel, room temperature 27-181 27-89 Corrugation-stiffened skin compression panel after failure, room temperature test 27-182 37-90 Thermocouple locations f o r the corrugation-stiffened- skin compression panel 27-183 Panel shortening curve AL/L f o r corrugation-stiffened- 27-91 skin compression panel, 1400°F test 27-184 Corrugat ion-stiffened skin compression panel after 27-92 failure, l4OO0F test 27-185 S t r a i n gage locations f o r trapezoidal corrugation 27-9 3 compression panel 27-186 Axial s t r a i n s f o r trapezoidal corrugation compressj sli 27-94 panel, room temperature 27-187 Panel shortening curve AL/L f o r trapezoidal corrugation 27-9 5 compression panel, room temperature 27-196 27-96 Normal deflections f o r trapezoidal. corrugation 27-196 compression panel, room temperature Trapezoidal corrugation compression panel after failure, 27-97 room temperature test 27-197 Thermocouple locations f o r the trapezoidal corrugation 27-98 compression panel 27-198 Panel s h o r t e n h g curve AL/L f o r trapesoidal corrugation 27-99 compression panel, 1400°F 27-199 27-100 Trapezoidal corrugation compression panel after 27-200 failure, 1400° test 27-201 27-101 S t r a i n gage locations f o r beaded compression panel 27-102 Axial s t r a i n s fcr beaded compression panel, room 27-202 temperature 27-103 Panel shortening curve AL/L f o r beaded compression 27-211 panel, room temperature 27-104 Normal deflection$ f o r beaded Compression panel, 27-211 room temperature 27-10 5 Expansion of beaded compression panel due t o axial 27-212 compression loads room temperatawe Figure Page 27-106 Beaded compression panel a f t e r failure, room temperature 27-213 27-107 Strain gage locations f o r tubular compression panel 27-214 27-108 Axial strains f o r tubular compression panel, room temperature 27-215 27-109 Panel shortening curve &IL/L f o r tubular compression panel, room temperature 27-223 27-110 Tubular compression panel a f t e r failure, room tempera- ture test 27-224 27-111 Thermocouple location f o r the tubular compression panel, l4OO0F t e s t 27-225 27-112 Panel shortening curve AL/L for tubular compression panel, U O O O F t e s t 27-226 27-113 Tubular compression panel a f t e r failure, l4OOOF test 27-227 27-ll.4 Strain gage locations f o r the circular arc corrugation shear panel 27-228 27-115 Relationship of principal strains and applied vertical cantilever loading f o r circular arc corrugation shear panel (TIG weld with Re& 13 f i l l e r wire), room temperature 27-229 27-116 Relationship of shear s t r a i n and applied vertical cantilever loading f o r ciroular arc corrugation shear panel (TIG weld with Ren6 4 l filler wire) , room temperature - 27-23 27-117 Circular arc corrugation shear panel (TIG weld with R e d 4.l f i l l e r wire) after failure, room temperature t e s t 27-230 27-118 Relationship of principal strains and applied vertical cantilever loading for circular arc corrugation shear panel (TIG weld with Hastelloy W f i l l e r wire), room temperature 27-231 27-119 Relationship of shear strain and applied vertical cantilever loading f o r circular arc corrugation shear panel (TIG weld with Hastelloy W f i l l e r wire), room t em@ rat ure 27-231 27-120 Circular arc corrugation shear panel (TIG weld with Hastelloy W f i l l e r wire) a f t e r failure, room temperature 27-232 27-121 Degree of conservatism i n the wide-oolumn analysis as applied t o compression panels vs width-to-length ratio 27-233 Figure Page 27-122 Comparison of tubular and beaded configuration i n i t i a l buckling test results with predictions 27-231, 27-123 Plasticity factors for 0.016-in. Rend 4.l sheet at room temper at ure 27-235

27-124 Plasticity factors for 0.016-in. - 0.019-in. Re& 41

sheet at U O O O F 27-235 27-125 Plasticity factors for 0.060-in. Rent$ 4.2. nheet, & t room temperature 27-236 PRECEDING PAGE BLANK NOT FILMED.

LIST O F SYMBOIS A Area; mean of area enclosed by outer and inner boundary Area of element n of cross section ph X and Y distances between simply supported edges of plate a,b a/b Aspect r a t i o b Width of flat p l a t e f o r buckling analysis Ratio of c r e s t width t o diagonal width f o r trapezoidal b/d corrugak ion c &d f i x i t y Stiffness coefficients of governing differential equation DI, DII, D1, D2, D3 of plat e d Width of diagonal element of trapezoidal corrugation Eel Elastic modulus E, Compression modulus of e l a s t i c i t y Compression yield strength FCY Ftu Ultimate t e n s i l e strength Tensile yield strength F t Y Stress corresponding to modulus of 0.7 Eel FO .7 fc, C r i t i c a l crippling s t r e s s C r i t i c a l crippling stress of element n of cross section fccn C r i t i c a l compressive buckling s t r e s s f o r sides of trapezoidal fc,d,cr corrugation Shear s t r e s s f/s fs,m Crktical shear buckling stress G Modulus of rigidity Moment of i n c r t i a , in.

I Moment of inertia per unit length, in.

f

-

J Torsional stiffness per unit length K ' Spring constant Stress correction factor

Ka

k Diagonal tension factor Buckling coefficient i n analysis of local compressive buckling Buckling coefficient i n analysis of local compressive kc,d buckling f o r diagonal element of trapezoidal corrugation Buckling coefficient i n analysis of shear buckling kS L Length L L ' -Effective panel length

&

Material correction factor W F m Number of half-waves i n plate buckling equation n Shape parameter Pitch Radius Room t emperatwe

m

t Thickness

-

t Effective panel thickness

-

Area per unit of diagonal width f o r beaded concept tL U Length of median boundary Rectangular Cartesian coordinates &ne1 length i n orthotropic panel buckling analysis 27-xvi Panel effective width i n orthotropic panel buckling analysis *I1 P C Correction factor Y

AL Change i n length

Panel shortening r a t i o BL/L

6, Flas%ici.tu reduction fac

or for calculat n g i n t rl 31 buck&@; stress- of a flat p l a t e simply supported -, E , Secant p l a s t k i t y reduction f a c t o r % Eel E

st, S t a r e l l ' s p l a s t i c i t y reduction f a c t o r

%IT

Et

-

Eel' Tangent p l a s t i c i t y reduction f a c t o r rlT

-

,/--, P l a s t i c i t y reduction f a c t o r f o r calculating t h e buckling stress f o r the c i r c u l a r a r c see-tions v Poissons r a t i o Summation

Section 27 STRUCTURAL ELENENT TESTING TEST PIAN Standard element tests were conducted concurrent with t h e theoretical analyses and the latter portion of the mterial screening test program (section 5 ) t o evaluate primary structural concepts applicable t o wing struc- ture designs. The results of' these tests and subelement tests (section 5 ) were used t o refine the methods of analysis and concept design.

Twenty-two structural element panels were designed and fabricated f o r test and evaluation i n accordance with t h e s t r u c t u r a l element test shcedule outlined i n table 27-1.

End closeout, crippling, compression panel, and inplane shear tests were conducted at room temperature and at 1400OF f o r evaluation of the con- The information obtained from these tests included: struction concepts.

1. Evaluation of end-closure designs 2. Evaluation of joining mettods Combined effects of temperature and b a d 3 .

4. Substantiation of element and panel shear, crippling, and compression buckling stresses.

Details of the panel elements, fabrication and assembly schedules, test arrangements, instrwnentation, test procedures, test results, and comparison of analyses with test results are presented i n t h i s sectton.

DE3CRIPTION AND FABRICATION O F PANEL ELTINWE Twenty-two panels were constructed for t e s t and evaluation.

The WneL types, sizes, and. the number of each panel element fabricated aye given i n t a b l e 27-2.

A detailed description of each of these panel elements is given below and includes t h e fabrication and assembly schedules used i n t h e i r construction.

Panels !l?ubular The t e s t panel design (fig. 27-1) consists of two beaded skins, four Beaded face sheets fingered end doublers, and two end bars f o r testing, were formed i n a high-pressure Verson-Wheelon prees; doublers were blanked 27-3.

using s t e e l r u h dies. Fhd bars were intta?.led u s h g Hi-bk, high-strength fasteners. Ehds of the panel azaembly were mchine-ground t o a close tolerttncc (-FO,OOl inch) across the panel width. Cr5-t ,.:15ngand end closeout panels were saw cut from fhll-lnngth panels; ends of' c:r-L;>qling pane-s and one end of an end closeout panel were cast i n Densite o r kPyrofarm f o r testSng, depeading on the test environment.

Fabrication and assembly plan f o r 3L -%An. panel:

1. Formed skins - two required per ~ m 1 assembly

Shear 24.0-in. by 34.0-in. blanks, 0.016-in. gage Re& 41 Process clean - degrease, alkaline wash, pickle rinse, isf 1 dry &case i n preoxidized Type 321 qres s t e e l envelope

(26 .O-ir.. by 36 .O-in . ) , evacuated Fund seam-welded

F i r s t stage forming at 3500 p s i (17-20$ elongation) on form block FB-CL ll.25-1-9 (fig. 27-2) An;:&,1 pckage - a i r furnace 1950' fo 2000% f o r 15 min; air cool t o 1C33?3' within 3 sec

Pickle - n i t r i c -hydrofluoric (vapor blast t o remov? residual scale)

Second stage forming at 3500 p s i (8-1c$ elongation) on FB-CL 1125-1-9 -air furnace 1 9 5 0 ' t o 2000°F f o r 15 min; air cool

Pickl- - nitric-hydrofluoric (vapor blast t o remove residua?. scale)

Third stage forming at 3500 p s i (4-5$ elongation) on FB-CL 1125-1-9 Anneal package - a i r furnace 1.950' ' : o 2000°F f o r 1 5 min; air cool t o 1000°F within 3 sec Reinove package from part; hand s~F?T~,s* F i n a l stags forming at 8000 psi (2-35 elongation) on FB-CL 1125-1-9 b y out finish p n e l dimensions .ind shear D r i l l No. 30 vent holes i n ont; erid o : Leads, une panel only (fig. 27-3) Debuxr Clean f o r welding (chrmic-sulfuric per ref. 274) Prepare coupons Prm trim material, b required 2'7-2 2. Finger doublers - f o u r required per panel Shear 5.45411. by 17.37-in. blanks, 0.030-in. gage Reng 41

Blank - s t e e l rule die i n 200-ton punch press

Deburr Final clean prior t o assembly (chromic-sulfuric per ref. 27-1)

End bars - four required per panel 3 -

Saw .38-in. by 1.00-in. Inconel bar t o 17.38-in. length N o m l i z e at 1800 F f o r 30 min; air cool Check and straighten Eill one face and one edge square Clean prior t o assembly 4.

Assembly - record weight of each d e t a i l part

Locate panels and doublers i n universal weld fktcre (fig. 27.*'a); resistance weld (figs. 27-5 and 27-6); (ref. 27-21 Remove electrode deposit -hand swab using chromfc acid followed with almhol rinse Age and heat oxidize at S400% f o r 16 h r i n air furnace using ce7amic fixtures f o r heating and air cooling D r i l l and r e a m panel and end bars i n d r i l l fixture (no coolant or lubricant ) Deburr holes Record weight of assembly, l e s s end bars Install Eti-Lok fasteners H11 panel end bar6 normal t o axis of beade within Wo l5*, faat and parallel within Xl.002 in. (ref. 27-1).

Modified tubular panels. - The end closeout designs were modified by the

addition of-tapered doublers (0.016 in. thick by 5 . 0 0 in. long) t o each side of each f l a t of the finished panel assembly (figs. W-J. and 27-7). The area t o be was hand-sanded scraped, and wire-brushed t o remove covered by these doublers

oxide. Doublers were sheared t o size (0.b in. wide at one e.d; 0.26 in. wide

a t the other end), cleaned (alkaline wash, chromic/sulf'uric pickle, hot water and deionized water rinse, a i r dry), and located 2n posttion by probe tack welding. Structural welds followed schedule previousLy established for four thicknesses of 0.016 in. Rene 41. Due t o inability t o remove a l l surface con- tamination from the heat-oxidized surfaces, spot-weld strength per spot was reduced; average values obtained * o m t e s t strips indicated loss of approxi- mately 30 l b per spot. Average shear strength of spots was 517 lb (547 lb per spot on clean material), which exceeds M I L W-6858C specification requirements.

Fabrication of crippling and end closeout panels. - One panel assembly

was completed i n accordance w i t h the above plan except that end bars were The remaining panel w a s then sawed into re- omitted from one end of panel.

qGired end closeout and crippling sections (figs. 27-8 and q - g ) .

1. Crippling panel end casting for test - one crippling panel after being

sawed t o 8.0-in. length, w a s fixtured i n 1.0-in. deep mold and cast with Densite. After drying, panel w a s reversed and o p p s i t e end cast in Densite. a d s were then ground flat, parallel, and normal t o bead axis.

A second crippling panel was cast in Pyrofoxm (a high-temperature ceramic) i n a similar manner, except that shims were placed t o provide space for panel elongation during high-temperature testing.

2 . Ehd closeout panel was sawed t o 9 . 0 - i n . length with sawed edge cast i n Densite, then ground parallel t o end bars.

Beaded Panels The teFt panel assembly (fig. 27-10) consists of one beaded skin,' four Beaded panels were formed by hydrau- fingered ena doublers, and four end bars.

l i c forming in a Clearing l5OO-ton press, using auxiliary pump for fluid movements.

Fabrication and assembly plan f o r 3O.O-in. panel

1. Forme& skins - one required per panel assembly

Shear 32.O-in. by 38.0-in. blank, 0.020-in. gage Rene' 41 Deburr Clean -alkaline wash, pickle, rinse dry First stage forming at 2000 psi using EFB (hydraulic forming block) (fig. 27-11) Degrease

- a9r f m a c e 1950° t o 2000°F f o r 10 min; a i r b l a s t cool

h e a l 1000°F within 3 see t o

Descale - deoxidizer, nitric-hydrofluoric pickle, rinse, oven

&Y

Final stage forming a t 3000 psi using HFB - CL ll25-l-lO

lay out finish panel dimensions and shear (fig. 27-12} Flnal clean prior t o assembly (ref. 8-1).

Prepare coupons from t r i m material, 8 required

Finger doublers, 4 required - same as for tubular panel

2 .

Fhd bars, 4 required - same as for tubular panel

3.

4 . Assembly - same procedure as for tubular panel (figs. 27-13 asd

27-14); record weight of each d e t a i l and final assembly, less end bars Modified beaded panel. - m e end closeout designs were modified by exten- sion of the finger doublers.

This w a s achieved by use of 0.016-inch and 0.020-inch by 2.0-inch doublers laminated on each side of the original fingers and extending a t o t a l of 3 . 0 inches toward the panel center. Twenty 0.020-inch by 0.35-inch by 2.0-inch; Ldjznty O.Ol6-inch by 0.35-inch by 2.0-inch; and twenty 0.016-inch by 0.35-inch by 3.0-inch Re& 41 doublers were resistance spot-welded (fig. 27-25).

N e w weld schedules were developed f o r the laminated sections modzfied bv the finger doublers. %e locations and thicknesses for the laminated section are as follows: Location 2.0 in. beyond A t end of 1.0 in. beyond 3 . 0 in. beyond finger end of finger end of finger end of finger doubler doubler doubler doubler - Added ( -016 x 2.0) 0.016 0.016 Added ( .016 x 3.0) 0.016 0.016 0.016

Added (.020 x 2.0) 0 .ox) 0.020

Finger doubler 0.030 Corrugation 0.018 0 . 0 1 8 o . 0 1 8 0.018 Finger doubler o .030

0 .om

Added (.020 x 2.0) 0 .om

Added ( . 0 1 6 x 3.0) o . 0 1 6 o .016

0.016 Added ( . 0 1 6 x 2.0) 0 . 0 1 6 0 -016 - I

0.118 I 0.122 Total thickness o .082

o .050 I No. sheets Fabrication of crippling and end closeout panels: End closeout - similar t o tubular panel.

Crippling - similar t o tubular panel (fig. 27-16).

Trapezoidal Corrugation Panels The t e s t panel assembly (fig. 27-17) consists of a trapezoidal corrugation center, two trapezoidal corrugation ends, four finger splices, and two zee sections. The corrugations were formed on a corrugating die as a one-piece panel, then cut into c a t e r and end sections. Doublers were blanked from sheet using s t e e l rule die i n a punch press; zee sections were power-brake formed on standard tooling.

Fabrication and Assembly Plan f o r 30.0-in. panel:

1 . Trapezoidal corrugations - one 22.0-in. and two 4.0-in. sections

required per panel ads embQ .

Shear 32.O-in. by 36.0-in. blank from O.Ol6-in. gage Rene' 41 rl-6 Deburr Form i n corrugation d i e CD-CL 1125-1-12 (fig. 27-18) S i z e t o 0.578-in. height, standard tools, power brake Iay out f o r saw Saw mrts (center corrugation and end corrugations) Prepare coupons from t r i m material, 8 required Deburr

Joggle - cerrobend cast tooling, arbor press

Clew f o r welding

Finger splices - 4 required per panel assembly

2.

Shear 3.71-in. by X).O-in. blank from 0.040-in. gage Renee 41 Deburr Blank - s t e e l rule die B [ D CL ll.25-1-11-6 and -7.

Cut t o length - shear per -6 and -7 details

Deburr Clean for welding.

Zee sectfon - 2 required per panel assembly

3 - Shear 2.58-in. by 19.46-in. blanks from 0.020-in. gage Rene' 41 Deburr Power brake form, standard tooling.

Clean for welding 4. Assembly Record weights of each d e t a i l part (fig. 27-19) Locate corrugation sections and zee sect5ons i n weld fixture; resistance weld I n s t a l l splice plates; resistance weld (fig. 27-20) Remove electrode pickup, swab with chromic acid Alcohol rinse Age and heat-oxidize, 14OO0F, air furnace, f o r 16 h r (fig. 27-21) M i l l ends of panel square, parallel, and normal t o corrugated axis End cast in Densite and F’yroform (one panel each material)

Fabrication of crippling panels. - One center corrugation w a s saw-cut into

two 8.0-in. lengths, aged and heat-oxidized, 1400°F f o r 16 hr. Ends were cast (one panel i n Densite, the other i n Pyroform) then ground f l a t , parallel and normal t o corrugation axis. (fig. 27-22).

Corrugation-Stiffened Panels The panel assembly (fig. 27-23) consists of one corrugated sheet w i t h formed closeouts, one flat skin, two tapered fingered end doublers, two end spacer doublers, and two Tee end bars. Corrugation, skin, and doublers aye resistance spotwelded together; end Tees are attached w i t h high-temperature shear fasteners. T w o full length panels (30.0 in.), two crippling panels (8.0 in.) and one end closeout panel (9.0 in.) were fabricated.

Fabrication and assembly plane for 30.0-in. panel: 1 . Corrugation with formed closeouts -one required per panel ass embl y Shear 24.0-in. by 34.0-in. blank from 0.016-in. gage Reng 41 Encase i n preoxidized type 321 Cres s t e e l envelope F i r s t stage forming i n Verson-Wheelon a t 6000 psi (17% elongation) on CL 1125-1-13 form block (fig. 27-24) Anneal at 1950’ t o 2000% f o r 25 min; air quench Remove scale -pickle and vapor blast Second stage forming at 6000 psi using f i l l e r s t r i p s i n CL 1125-1-13 form block ( 1 8 elongation) Anneal Pickle Third stage forming at 6000 psi using f i l l e r s t r i p s i n CL 1125-2-13 form block (10% elongation) 27-8 Anneal Remove envelope; hand shear Final stage forming at 10 000 psi using filler s t r i p s i n 1125-1-13 form blcok (2% elongation) Lay out and trim to 19.00 in. by 30.75 in.

Prepare coupons from t r i m material, 8 required DrillNo. 30 holes i n one end of each bead Deburr Clean for welding 2 . Skin -one required per panel assembly Shear l9.00-in. by 30.75-in. finish skin from 0.026-Sn. gage Reng 41 Deburr Clean f o r welding -two required per panel Tapered fingered doublers 3.

Shear 4.00-in. by 1g.OO-in. blanks from 0.060-in. gage Rene' 41 Deburr

Blank fingers - BD CL 1125-1-13 steel. rule die i n 200-ton

punch press Deburr Mill taper fingers -mill fixture Deburr Clean f o r welding 4. End spacers - t w o required per panel Shear 0.75-in. by 19.00-in. blanks from 0 .Ob-in. gage R e n e ' 41 Deburr Clean f o r welding 5 . Tee bars Saw 19.00-in. blanks from 0.38-in. by 1.00-in. Inconel 600 all-oy bar Stress relieve at 2000% f o r 30 min

Check and straighten - hand arbor press

M i l l Tee configuration Clean f o r assembly 6. Assembly Record weight of each detail part (fig. 27-25) Resistance weld skin, corrugation, and doublers i n universal weld f i x t u r e (ref. 27-2) Remove electrode pickup; chromic acid swab Alcohol rinse Age and heat-oxidize at 1400% f o r 16 h r (fig. 2'7-26) D r i l l f o r Tee end attachment Deburr I n s t a l l end t e e s with Hi-Lok fasteners Grind ends of Tee members flat, pmallel, and ndrmal t o bead axis.

7. Fabrication of end closeout and crippling specimens. One f u l l length p n e l was cut i n t o smller specimens which, i n turn, were end cast either i n Densite or i n Pyroform similar t o t h e circular- a r c stiffened end closeout and crippling specimens (fig. 2r(-27).

Circular-Arc Corrugation Shear Panels The panel assembly (fig. 27-28) consists of circular-arc corrugated web design, with channel caps and edge doub"ers. The cap is TIG welded t o the corngation using Rene'41 and Hastelloy PI f i l l e r wires and doublers are resis- tance spot -welded t o t h e corrugation.

27- 10 Fabrication and Assembly Plan: 1. Corrugation -one required per panel assembly Shear 20.0-in. by 26.0-in6 blank from 0.016-in. gage Rent5 41 Deburr

Form on FB/CL 1125-1-12 - Verson-Wheelon a t 5000 psi ( f i g . 27-29)

Lrly out and saw/shear t o 15.50 in. by 17.00 in.

Prepare t e n s i l e coupons from t r i m material, 4 required Grind ends flat, parallel, and normal t o axis of corrugation using CL 1125-1-12 TIG weld fixture Deburr Clean for welding Side doublers - f o u r required per panel assembly 2.

Shear L.42-in. by 15 .&-in. blanks from 0.016-in gage Rene’ 41 Deburr Clean for welding

Cap Channel - two required per panel assembly

3.

Shear 3.25-in. by 17.00-in. blanks from 0.060-in. gage Rene’ 41 Debuxr D r i l l 10 V-size holes (0.376-0.383-in. diam) using drill j i g Deburr Form flanges -power brake using end holes f o r location of bends Clean f o r welding 4. Assembly Record treight of all d e t a i l p a r t s (fig. 27-30) Locate corrugation i n weld fixture Trace contous: and Ink template 27- 11 Locate cap i n position and seal Weld cap t o corrugation, tracing from template ( f i g . 27-31) Reposition and repeat sequence f o r other cap Weld schedule: Vicers DC arc welder Model M T 4K40, 400 amp

Weld amperage - 85

Voltege - 10

Tm.vnX speed - 9 in. per min

Electrode - thoriated tungsten (8 ThOg), 0.093-,n.

diameter

Torch nozzle - O.3l-in. diameter

Torch shield gas -Argon at 12 fd/min Backup gas -Argon a t 2 5 ft3/min

Trailing shield - 3.0-in. by 6.0-in. glass cloth attached

t o torch F i l l e r wire .- 0.045-in. diam Hastelloy W for one other p n e l , 0.060-in. diam Rene' 4.2. for other panel; both automatic feed edge doublers, hand clamp and resistance spot weld I n s t a l l

Iay out and d r i l l ten 6.4 mm holes (0.251 - 0.258-in. diam) each

edge of panel Deburr

heat-oxidize - 1400% for 16 h r ( f i g . 27-32)

Age and Record weight of finished assembly Spar Cap Crippling a n e l The panel assembly ( f i g . 27-33) consists of a circular-arc corrugation The caps a r e TIG welded (melt through) t o corrugation.

web and two channel caps.

27-12 Fabrication and assembly plan: Detailed fabrication and assembly plan f o r t h e beam cap crippling panels i s identical t o t h e circular-arc corrugation shear p n e l , except f o r the follcwing: l. Hastelloy W f t l l e r wire was use6 t o join t h e cap t o a r c (welding schedule same as f o r t h e in-plane shear panel t e s t ) The height of t h e cap flanges is 3/8 i n . 2.

The ends of t h e panel were milled flat, square, and p a r a l l e l 3 .

TEST SIBUP Room Temperature Compression Tests The test setup fc:. t h e room temperature compression tests of t h e end- closeout, crippling, compression panels, and t h e beam cap crippling specimens was essentially t h e same.

!&pica1 test arrangements are shown i n figure 27-34 f o r the crippling and compression panels.

The compression panels are shown positioned in the compression bay of a, suitable capacity t e s t i n g machine and are located between a base plate and a compression head test f i x t u r e . A l l bearing surfaces of these fixtures were Blanchard ground flat and p a r a l l e l .

Two cylindrical plates are shown sandwiched between the compression head and t h e positioning (or movable) head of t h e test machine. These plates are tapered i n thickness (0.001 in./in.) t o allow f o r i n i t i a l p a r a l l e l alignment of the ground surfaces of t h e base p l a t e and compression head test fixture prior t o i n s t a l l a t i o n of the test panel. The i n i t i a l alignment of t h e compression sur- faces was held t o within 0.0005 inches across t h e t o t a l bearing surfaces of t h e loading fixture.

Prior t o i n s t a l l a t i o n of t h e test panel, s l i t tubes were attached t o t h e f r e e edges of t h e panel. t o provide simple edge support. Sufficient clear- ance (0.059 i n . at each end o f t h e tube) was provided a t t h e tube ends t o avoid t h e introduction of a x i a l tube loading due t o specimen contraction when test loads were applied.

Elevated Tempemture Compression Tests The t e s t setup f o r t h e elevated temperature compression tests of t h e crippling and column panels was essentjally the same. A t y p i c a l test armnge- ment is shown i n figure 27-35, I n addition t o t h e room temperature test f i x - tures previoue7.y described (Including t h e tubular edge supports), figure 27-35 shows two Pyroform (cast ceramic) blocks, 1/2 in. t h i c k by 6 i n . wide by 24 in.

long, and a 3/16-in. thick Inconel bearing plate attached t o t h e loading: and The Pyroform blocks adjacent t o t h e reaction heads of t h e t e s t machine.

Inconel plates contained nichrome heating elements which were threaded through pre-cast holes in .the blocks. This arrangement reduced heat losses from t h e ends of t h e test p n e l and provided insulation a t t h e test machine loading and reaction heads.

27-13 The F'yroform block heaters were electriciii1.y connected irA p z a l l e l and were energized by an Lnductrol type 60.cycle por:r supply. ?\e Inductrol unit is essentially a two-winding power transforr .8r t h a t incorporates a t r % - - able secondary c o i l permitting a variable electr*.'.::al output from t h e t r a n s - former. This unit was used t o proviaa e l e c t r i c a : isolation between the block heaters and t h e @O-volt, 60-cycle power supply i z e d f o r t h e radiant heat lumps.

An overall view of the elevated temperatum test setups t h e c r i p - pling and column panels i s shown i n figure 27-36, Two radiant heat lamp assemblies were used, one assembly on either side of t h e test, p n e l . Rerrasil batting (a high-temperature spun gLaas insulatioE blanket) was used t o encap- sulate t h e p n e l t e s t setup. The lamp S..,rsembiies consisted of lOOOT3/C1~/HT

q u a r t z lamps and t h e Research Incorporated Aii8-SI.2 lampholders . Two Thermac

power units were used t o energize t h e heat Lam ? ssemblies. Chromel--0 a n e l thermocouples spotwelded centrally on each side .IT t h e panels provided t h e feedback signals t o regulate t h e power c o n t r o l k r $ .

Shear Panel Tests The general arrangement f o r the in-plane shear test is shown i n The t e s t panel i s mcmted i n 3, cac.;llever type l w d i n g test figure 27-37.

fixture. Flexure pivots a r e incorpurated i n t h e test fixture design a t each of t h e four corners t o eliminate t h s f r i c t i o n associated with pin conn ,tions.

A hydraulic jack was used t o apply v e r t * z a l loadin$ t o t h e cantilevered test fixture. Hydraulic pressure was supplied t o t h e gack by means of an Edison load maintainer. Test loads were inonitored by aeans of a load transducer mounted i n s e r i e s with t h e hydraulic jack. Iateral su2ports were pin- connected t o the cantilever f i x t u r e t o p r e v a t racking during load application.

The instrumentation schedule f o r t h e s t r u c t w a l element tests is out- l i n e d i n t a b l e 27-3 and indicates t k c number of s t r a i n @;ages and thermocouples used f o r each m n e l test. The strain gages bsrj3 included Elaldwin Lima HamLlton ( B I Z ) f o i l gages, type FAE-25-12 ~ 6 , and h d d f o i l gages, type Cb-12%.

A n epoxy d h e s i v e system was used t o bond the BLH gzges t o t h e specimens using accepted standard s t r a i n gage bonding techniques. The Budd gages were bonded

using the water-activated epoxy acchesive incorparii-:ed with each gage Speci -

men a x i a l deformtions (panel shcrtening) were p3az fled by means of e l e c t r t c a l at +: = :tour corners of t h e compression head deflectiou: transducers mounted f i x t u r e pi :viouuly described. Th9- :+ ,ection transducers, normally &signed i .

as LvlMls (liiwsr variable differer:'.: : : I transducers) a r e Model SS-105 (6-voic excitation), G. L, Collins Cwporation. t Specimen temperatures were measured using 30-gage chromel-alumel thermo- couples having glass -over-glass type insulation. The thermocouples were I I attached t o the 5est specimens by means of t h e c a p c r t a n c e discharce spot- !

weld methm!.

A 150% Pace reference junction was used for t h e thermocouple &ta reference point. The strain gage and thermocouple locations and i d e n t i - fication nutlhers f o r each panel specimen a r e presented i n the paragraphs describinc:; test r e s u l t s 27-14

I

DATA ACQUISITION A modified Sadic, 200-channel medium speed data acquisition system was used for the panel tests.

The systein has an ir-herent maximum speed of approxi- mately 250 msec per data point w i t h five dig% resolution t o 230 000 counts and 0.03 p e r c x t linearity. "his represents a system accuracy of &lo micro- inches/in. strain for all strain levels up t o 230 000 &croinches/in. strain, or +0.2% when using chromelalumel thermocouples from -300% t o +TOOOF. The syecem converts the millivolt signais from s t r a i n gages, deflection transducers, and thermocouples into d i g i t a l data and stores them on perforated tape. This information is then transferred t o I B M cards for further processing. For t h i s program, tab runs were the s i d product f o r data disp2ay. The s t r a i n gage and deflection transducer data have been plotted i n curvilinear form; t h e thermocouple data are presented i n tabular form. A l l . of these data are included i n the test results paragraphs of t h i s section.

TEST PROCEDURES Preliminary T e s t s Prior t o conductingthe compression f a i l u r e tests at either room or elevated temperature, a preliminary t e s t run was conducted t o assure proper specimen alignment i n the test lnachine so that a uniform loading would be achieved across the entire specimen width. Test loading during t h i s align- mg procedure was hell below 5; percent of the predicted i n i t i a l buckling load f o r the p r t i c u l a r specimen configuretic tested. Uniformity of load distribution was dztermined by the L W T readings that measured tect; head Jisplacement and by panel s t r a i n gage readings. After satlsfactory alignment of the test panel was achieved, t h e failure t e s t was conducted.

Failure Tests, Room Temperature Compression Panels The fallure t e s t consisted of the application of compresi5on loads i n suitable steps while pafie1 d e f o m t i o n s and strains were recorded at each loadl g step. The maxi- T e s t loading i9 this mnner was continued t o failure.

mum load sustained by the Wnel was obtained from the reading of the load indi- cating follower located on the face of the test machine console.

Failure Tests, Elevated. Temperature Compression Panels After satisfacto alignment of the test panel, the t e s t specimen was

then heated t o the 1400 3 t e s t temperature. This m s accomplished by first

energizing the heating elements i n the Pyroform blocks which i n turn heated the Inconel bearing plates located between the specimen ends and the Pyroform heating blocks. The radiant heat Lamps were then energized by means of the !l?hermac power regulators.

27-15 a In the actual operation of the Thermac units, the set-point control was adjusted for the desired temperature as determined from the calibration curves The l i m i t e r control was then advanced slowly t o l i m i t the r a t e of provided.

t panel temperature rise. The advantages accrued from t h i s procedure were.

J

1. Limitation of the temperature rise rate 2 . LimZtation of maximum power t o the lamp asEembliea, which is a safety feature i n the event of a c i r c u i t failure Minimum fluctuation of lamp intensity, which provides for a 3.

better steady-state temperature condition 4 . Increased l i f e of the radiant heat lamps.

I

Throughout the entire heating phase of the test p n e l t o the 1bO0F test temperature, a 2000-lb compression load w a s maintained on the specimen by the test machine operator. The test panel w a s soaked a t the test temperature f o r

I

a minimum of one-half hour before loading was commenced t o failure. The pro- cedure used for the failure test at the elevatedtemperature w a s identical t o the procedure previously described f o r the room temperature failure test.

Failure Tests, Shear Panels The procedure used t o conduct the shear panel failure t e s t s consisted of applying cantilever loads at a r a t e of approximately 100 l b per minxte by means of the Edison load maintainer. Test loading w a s interrupted t o permit s t r a i n gage data recording from both the back-to-back rosette gages on the t e s t panel and the load transducer mounted i n serfes w i t h the hydraulic jack.

A readout time of approximately three seconds was required. The load levels a t which data were recorl:ed are indicated by the t e s t points of the s t r a i n gage plots for each panel. An electrically operated dump valve was energized by hand t o dump the t e s t load a t panel failure.

TEST RESULTS Panel Material Tests The manufacturing processes used for the fabrication of the test panels M e - included interstage annealing for several of the panel cmfigurations.

chanical property t e s t s were conducted t o establish the material characteris- Stress- t i c s resulting From these processes, and ape summarized i n table ‘2(-4.

s t r a i n curves f o r each of the material conditions a r e presented i n figures 27-38, 27-39, and 27-40.

Panel Test6 A summary of the panel element t e s t s conducted i n t h i s program is pre- sented i n table 27-5 and includes panel descriptions, t e s t temperatures, panel areas (computed from the panel weight measurements), m e 1 ultimate loads, and 27- 16 ultimate stresses for each of the end closeout, crippling, comyession, and shear panel configurations tested. A detailed description of t h e test results for each of these configurations is given below.

Ihd closeout tests. - !Ibe end closeout panel. configurations were tested

at room temperature and includedthe following p e l s : Corrugation-stiffened panel- The strain gage locations for t h i s panel configuration are given in figure 27-41. Curve plots of the s t r a i n gage data are presented i n figure 27-42. A curve plot of the panel shortening due t o the applied compression loads is given i n figure 27-43. Photo- Thickness graphs of the panel after failure are shown i n figure 2 ' 7 - 4 4 .

measurements of the panel cross-section are given i n table 27-6.

Beaded panel- %e strain gage locations for t h i s panel configuration are given i n figure 27-45. Curve plots of the s t r a i n gage data are presented in figure 27-46. A curve plot of the panel shortening due t o compression loads i s given i n figure 27-47. Photographs of the panel after failure are shown i n figure q-48, Thickness measurements of the panel cross- section are given in table 27-7.

Tubular panel- The strain gage locations for t h i s panel configuration are given i n figure 27-49. Curve plots of the s t r a i n gage data are pre- sented i n figure q-50. A curve plot of the panel shortening due t o Rntographs of the panel compression loads is given i n figure q - 5 1 .

a f t e r failure are shown i n f;ieu,re 27-52. 'Ikickness measurements of the panel cross section are given i n table 27-8.

- Crippling panel tests were conducted a t room tempera-

each of the following panels.

Corrugation- stiffened skin panel Trapezoidal corrugation panel Beaded panel Tubular panel The spar cap crippling specimeii was tested at room temperature.

1 . Corrugation-stiffenea skin crippling panel room temperature test -

The strain gage locations f o r t h i s panel configwation are given i n figure 27-53. Curve plots of the s t r a i n gage data are presented i n figure 27-54. A cume plot of the panel shortening due t o compression loads i s given in figure 27-55. Photographs of the .panel a f t e r Tail- Thickness measurements of the panel ure are shown i n figure 27-76.

cross section are given i n table 27-9.

27-17

2 . Corrugation-stiffened skin crippling panel elevated temperature t e s t -

The themocouple locations for t h i s panel are given i n figure 27-57.

Tab runs of the thermocouple data showing the temperature distribution are presented i n table 27-10. A curve plot of the panel shortening due t o the applied compression loads i s given i n figure 27-58. Photo- graphs of the panel a f t e r failure are shown i n figure 27-59. Thickness measurements of the panel cross section are given in table 27-11.

Trapezoidal corrugation crippling pinel room temperature test -

3.

The *+?sin gege locations for t h i s p n e l configuration are given i n 27-60. Curve plots of the strain gage data w e presented i n f P = figdr: 27-61, A time plot of the panel shortening due t o compres- sion loads i s given i n figure 27-62. Fhotogra2hs of the panel a f t e r failure are ~h0w.n i n figure 27-63. Thickness measurements of the panel cross section are given i n table 27-12.

4. Trapezoidal corrugation crippling panel elevated temperature t e s t -

The thermocouple locations for t h i s panel are given i n figure 27-64.

Tab runs of the thermocouple data showing the temperature distribu- tion are presented i n table 27-13. A curve plot of the panel short- ening due t o the applied compression loads is given i n figure 27-65.

Photographs of the panel a f t e r failure are shown i n figure 27-66.

Thickness measurements of the panel cross section are given i n table 27-14.

Beaded crippling panel room temperature test - The strain gage loca-

5.

tions for this panel configuration are givw i n figure 27-67. Curve of the s t r a i n gage data are presented i n figure 27-68. A curve plots plot of the panel shortening due t o the applied compression loads is given i n figure 27-69. Photographs of the panel a f t e r failure are shown i n figure 27-70. Thickness measurements of the panel cross section are given i n table 27-15.

6. Beaded crippling panel elevated temperature test - The thermocouple

locations for this panel are given i n figure 27-27. Tab runs of the thermocouple data showing the temperature distribution are presented i n table 27-16. A curve plot of the *ne1 shortening due t o the applied compression loads i s given i n figure 27-72. Photographs of the panel a f t e r failure a r e shown i n f i g w e 27-73. Thickness measwe- ments of the panel cross section are given i n table 27-17.

Tubular crippling panel room temperature t e s t - The strain gage

7.

locations f o r t h i s panel are given i n figure 27-74. Curve plots of I?. curve plot of the strain gage data are gresented i n figu;-c 27-75, the panel shortening due t o the applied compression load i s given i n figure 27-76. Fhotographs of the panel a f t e r failure are shown i n figure 27-77. !&ickness measurements of the panel cross section are given i n table 27-18.

27-18 Tubular crippling panel elevated temperature t e s t -The thermocouple 8.

Tab runs of the locations for t h i s panel are given i n figure 27-78.

thermocouple data showing the tempei'ature distribution are presented i n table 27-19. A curve plot of the panel shortening due t o the applied compression loads is given i n figure 27-79. Photographs of the panel after failure a r e shown i n figure 27-80.

Thickness meas- urements of the panel cross section are given i n table 27-20.

Spar cap crippling specimen room temperature t e s t - The spar cap

9.

crippling specimen configuration presented i n figure 27-81 w a s tested a t room temperature.

The upturned flanges of the cap specimen were 3/8-in. The strain gage locations for this specimen are given i n figure 27-81. Curve plots of the s t r a i n gage data are presented i n figure 27-82. A curve plot of the spechen shortening due t o the applied compression loads is given i n figure 27-83. Photographs of the cap specimen after failure are shown i n figure 27-84.

Thickness measurements of the cap specimen are given i n table 27-21.

Compression panel tests. - Compression panel tests are scheduled at room

temperature and at 140O0F for each of the following panel configurations: Corrugation-stiffened skin panel Trapezoidal corrugation panel Beaded panel Tubular panel bdifications t o the finger doubler design were incorporated i n the beaded panel and the tubular panel as described i n the panel fabrication discussion.

After reviewing the room temperature t e s t data f o r the beaded compression paiiel, the elevated temperature t e s t f o r t h i s panel configuration was deleted from the test schedule. The results of the room and elevatedtemperature compression panel t e s t s are given below.

1 . Corrugation-stiffened skin compression panel room temperature test -

The s t r a i n gage locations for the corrugation-stiffened skin compres- sion panel are given i n figure 27-85. Curve plots of the strain gage data are presented i n figure 27-86. A curve plot of the panel short- ening due t o the applied compression loads is given i n figure 27-87.

Panel deflections, perpendicular t o the plane of the skin, were obtained from three d i a l gages mounted across the width of the panel.

These gages vere symmetrically positioned about the center of the panel, with the two outboard gages located approximately 5 inches from'the center gage. The normal deflections obtained from these gages are presented in figure 27-88. Photographs of the panel a f t e r failure are shown i n figure 27-89. Thickness measurements of the panel cross section are given i n table 27-22, 27- 19 I

2. Corrugation-stiffened skin compression panel elevated temperah : s t -

, J The thermocouple location for this panel are given in figure 27-90.

Tab runs of the thermocouple data showing the temperature distribution for this panel are presented in table 27-23. A curve plot of the , J panel shortening due to the applied compression loads is given in figure 27-91. Photographs of the failed panel are shown in figure

27-92. Thickness measurements of the panel cross section are given 1

in table 27-24. -1

Trapezoidal corrugation compression panel room temperature test -

3.

The strain gage locations for this panel are given in figure 27-93.

Curve plots of the strain gage data are presented in figure 27-94.

A curve plot of the panel shortening due to the applied compression loads is given in figure 27-95. Panel deflections perpendicular to

i

the corrugations were obtained from three dial gages mounted across the width of the pat& These gages were symmetrically positioned about the center of the panel, with the two outboard gages located approximately 5 inches from the center page. The normal deflections ;i Photographs obtained from these gages are presented in figure 2 7 - 9 6 .

of the panel after failure are shown in figure q-97. Thickness meas-

urements of the panel cross section are given in table 27-25. F

Trapezoidal corrugation compression panel elevated temperature test - 4.

The thermocouple locations for this panel are given in figure 8-98.

Tab runs of the thermocouple data showing the temperature distribu- tions for this panel are presented in table 27-26. A curve plot of the panel shortening due to the applied compression loads is given in figure g-99. Photographs of the failed panel are shown in fig- ure 27-100. Thickness measurements of the panel cross section are given in table 27-27.

5. Beaded compression panel room temperature test - The strain gage loca-

tions for this panel are given in figwe 27-101. Curve plots of the strain gage data are presented in figure 27-102. A curve plot of the panel shortening due to the applied compression loads is given in

-

figure 27-103. Panel deflection normal to the corrugation was meas- ured using a dial gage located at the centerline of the panel length and width. These data are presented in figure 27-1.04. Panel expan- sion (or widening) resulting from the applied compression loads was measured by attaching a scale to the panel and recording the change in position of fiducial lines. The expansion over two corrugation J pitches and four corrugation pitches is shown in figure 27-10?.

Photographs of the panel after failure are shown in figure 27-106.

Thickness measurements of the panel cross section are given in I

table 27-28. 1

6.

Tubular compression panel room temperature test - The strain gage

locations for this panel are given in figure 27-107.

Curve plots of the strain gage data are presented in figure ~?i'-108. A curve plot of the panel shortening due to the applied compression loads is given in 27-20 figure 27-109. Photographs of the panel a f t e r failure ere shown i n f i g w e 27-110. Thickness measurements of the panel cross section are given i n table g-29.

Tubular compression panel elevated temperature t e s t - The thermocouple

7 .

locations for t h i s panel e r e given i n figure 27-111. Tab runs of the thermocouple data showing the temperature distribution for this panel are presentg-3 i n table 2 ' 7 - 3 0 . A curve plot of the panel shortening due t o the applied compression loads i s given i n figure 27-112. Photo- graphs of the failed panel are shown i n figure 8-113.

Thickness meas- urements of the panel cross section are given i n table 27-31.

Shear tests. - In-plane shear t e s t s were conducted a t room temperature t o

of the corrugated web design. T w o evaluate the actual and predicted strength specimens were prepared: one TIG welded with R e d 41 f i l l e r wire, the other TIG welded with Hastelloy W filler wire, The results of the shear t e s t s are given below.

1. Sheer specimen TIG welded with Rem' kl FilJer wire - The s t r a i n gage

locations for t h i s sheer panel are given iu figure 27-114 which in- cluded back-to-back rectangular rosette gages. The rosette gage data were reduced by means of a computer and curve plots of the principal strains and maximum shear strain versus applied cantilever loading are presented i n figures 27-115 and 8 - ~ 6 . Photographs of the failed panel are shown in figure 27-l47. N o cracks were evidenced i n the Thickness measurements of the web cross section f o r t h i s panel weld.

are given i n table 27-32.

2. Shear specimen TIGwelded with Hastelloy W f i l l e r wire - The s t r a i n

Sage locations for t h i s panel are shown i n figure 27-114. Curve plots of the principal strains and maximum shear s t r a i n versus applied canti- lever loading are presented in figures 27-118 and q-119. Photographs of the failed panel are shown i n figure 27-120. No cracks were evi- denced i n the weld. Thickness measurements of the web cross section for t h i s panel are given i n table 27-33.

27-21 COWARISON OF ANALYSIS AND TEST FU3SULTS A summary of t h e correlation between the analysis and test results of the structural element test specimens is presented i n table 27-34. The following obscrvations are pertinent: The i n i t i a l compression buckling stress test results correlated rea- 1 .

sonably w e n w i t h i n i t i a l buckling stress predictions whenever it was possible t o positively identify i n i t i a l buckling i n e i t h e r t h e room- o r elevated-temperature tests. This correlation was noted f o r about The correlation with theory f o r the remaining tests half the tests.

indicated variations of approximately 50 percent. Table 52 gives reasons for t h e disagreements when possible. Tests i n which the vari- ation i s not explainable indicate a need for further tests.

The tubular and beaded-skin configurations exhiblt the same s e n s i t i v i t y 2 .

t o i n i t i a l imperfections and other disturbances as found i n axially compressed large t h i n cylinclrical shells. Consequently, a conservative method of predicting compression buckling was employed. Even with t h i s conservative m ethod, large variations between test and theory were noted, as described above.

A l l of the configurations exhibit about a ftl0 percent variation i n thick- 3 .

nesses across t h e i r widths, resulting from the forming process. T h i s is within the normal tolerance of the sheet material. The analytical methods show significant fluctuations with these thickness variations; however, fair agreement e x i s t s between test and theory when t h e thickness used i n calculations is based on the lower l i m i t of t h e tolerance.

4. The corrugation-stiffened concept demonstrated substantial post-buckling strength. Therefore, t h i s configuration has a higher potential than the i n i t i a l buckling analysis allows, providing permanent set due t o inelastic deformation after i n i t i a l buckling is acceptable. The test results indi- cated a variation of more than 20 percent over the predicted values for four of t h e tests performed. O f these tests, three were comparisons of the failure stresses.

Panel i n s t a b i l i t y was obxerved i n several of the tests of 30-in. speci- 5.

mens, and the test loads agreed favorably w i t h the analysis based on orthotropic theory f o r plates simply supported on a l l four sides. It i s shown t h a t the wide-column analysis used i n t h e optimizat9on of these configurations i s a simplified form of the orthotropic p.Late theory (n = 0). T h i s theory is v a l i d f o r panel width-to-length r a t i o s of 2 o r more when the unloaded edges are supported but it i s conservative f o r r a t i o s less than 2. However, the wide-column analysis is valid f o r any It i s con- width-to-length r a t i o when tested with unsupported edges.

c l u d e d t h a t the test panels demonstrated i n part t h e v a l i d i t y of t h e 27-22 theory. However, no tests were performed f o r unsupported edges, f o r buckling due t o inplane shear, or f o r bending due t o l a t e r a l pressure.

Since the optimum r a t i o for the hypersonic-vehicle wing structure i s greater than 2, the use of wide-column analysiP i n the optimization program i s also valid.

6. The configuration composed of a single beaded skin is susceptible t o a local i n s t a b i l i t y mode w i t h a very short transverse half-wavelength, which can bc predicted w i t h reasonable accuracy. T h i s triode of f a i l u r e Wac accounted f o r i n t h e analysis.

The shear-panel test specimens correlated with 7 percent of t h e calcu- 7.

lated i n i t i a l buckling stresses.

8. The measured i n i t i a l buckling stress on t h e spar cap was within 5 per- cent of t h e calculated i n i t i a l buckling stress.

A comparison i s presented i n t h i s section between analyses and test re- sults f o r the four semimonocoque wing-cover configurations, and f o r the circular-arc corrugated web and beam cap configurations f o r spars or ribs.

Because of the nonconventional nature of t h e wing-cover configurations, three types of t e s t s were performed: namely, (1) end closeout, (2), cripgling, and ( 3 ) compression panel t e s t s . The lengths of these test panels were nominally 9, 8, and 30 in., respectively; t h e end closeout panels and t h e crippling panels were expected t o yield s i m i l a r test loads f o r a given configuration pro- vided no premature failure developed i n the closeout area. Although t h e crip- pling panel t e s t s were conducted t o f a i l u r e , primary i n t e r e s t centered on t h e t e s t load at which l o c a l iuckling developed, since l o c a l buckling rather than crippling was the mode considered i n the optimization analyses f o r sizing Crippling ( f a i l u r e ) r e s u l t s are a l s o hypersonic cruise vehicle structures.

shown t o supplement t h e i n i t i a l buckling data.

A l l of the compression panels were supported along their unloaded edges w i t h s l o t t e d tubes. Because of the panel dimensions, the wide column analysis yields conservative predictions, and for t h i n reason t h e general i n s t a b i l i t y It should be noted, analysis of equation 10-34, section 10, was employed.

wide column analysis, as used i n t h e optimization analyses however, t h a t the f o r sizing hypersonic cruise vehicle structures i s an appropriate means for analyzing compression panels, when the width-to-length r a t i o i s equal t o o r This i s shown i n figure 27-12, which has been developed greater than about 2.

from t h e geometry f o r the tubular compression panel, discussed i n the following paragraphs. A curve representing an unstiffened p l a t e i s also shown for com- The l a t t e r , of course, could represent a p l a t e equally stiffened i n yarison.

the x and y directions, and shows t h a t a predominance of stiffening i n t h e x direction, as i n the tubular c o n f i s r a t i o n , causes the difference i n analy-ti- c a l methods t o decrease much more rapidly with increasing b/a. The optimum b/a developed by t h e optimization analyses f o r the semimonocoque wing-cover Of further i n t e r e s t i s the f a c t t h a t the wide column configurations i s 2.25.

27-23 analysis, and the general i n s t a b i l i t y analysis for compression panels as rep- resented by equation 10-34, section 10, may both be derived from the Same s e t of equations, where m, the number of half-waves i n the y-direction, i s taken t o zero for the wide column, and t o unity for the compression panel. Thus, the theory may be tested for any panel dimensions, but for b / a > 2 the simpler wide column analysis may be u t i l i z e d with small. conservatism.

Tubular Configuration - The t e s t panel drawing i s shown i n figure 27-1.

Analysis. After forming, the nominal sheet thickness of 0.010 in. varied across the panel width.

Traverses of the t e s t specimens are given i n tables 8-8, 27-18, 27-20, 27-29, and n-31 f o r the end closeout, room and elevated crippling, and room and elevated temperature compression panels, respectively. The cross- sectional areas presented i n table 27-34 are based on the actual weights of the specimens.

A correlation between the t e s t results f o r c r i t i c a l buckling of the circular-arcs i n compression and predictions based on equation 12-14 of section 12 are presented i n figure 27-122 and table 27-35. Rt=Crm figure 27-122, it i s evident t h a t the average stresses i n the t e s t panels at buckling f o r the beaded configuration were well below the predicted stresses. The tubular elevated panel t e s t failed at an average s t r e s s greater than the predicted based on least measured thickness. With t h i s exception, a l l of the panels buckled a t an e l a s t i c average stress and thus p l a s t i c i t y reduction factors based on the average stress do not come into play.

The c r i t i c a l buckling stress i n the arc of the tubular configuration is, therefore : where, E = 29 x lo6 p s i a t 75OF = 21.2 x 106 at 1 4 . 0 0 ~ ~ R = 1.05 in.

t = 0.011 in.

then, fC,Cr = lo5 300 psi a t RT = 78 500 psi a t 1400°F i 27-24 The i n i t i a l buckling stress of the f l a t i s based on a simple supported flat plate : where t = 2 x single flat thickness, in. = 0.030 in.

b = 0.556 in.

t) = (figures 2'1-123 and 27-124) then fC,Cr = 130 000 psi a t RT = 111 000 psi a t 1400°F Note that the supports along the unloaded edges of the panel are arranged t o simulate the next tube; that is, the visible f l a t at each edge is 0.556 in.

However, since the t o t a l edge width i s 1.10 inches, a width of f l a t equal t o 0.544 in. i s hidden from view inside the edge support. Because t h i s f l a t has a free unloaded edge, the buckling coefficient f o r t h i s f l a t i s 0.7, rather 4.0, and the buckling stresses are 39 930 p s i at room temperature &id than 29 200 psi a t 1 & O 0 " . These are the lowest local buckling stresses i n the panels and they ma) have precipitated buckling of the panel. Pertabations i n the tubes closest t o the unloaded edges of the panels due t o buckling of the panel edges inside the support tubes may have occurred, indicating t k l t a smaller f l a t with a small flange may be required at the panel edges.

Local compression buckling i n the f i e l d of the end closeout and crippling t e s t specimens i s expecied t o occur i n i t i a l l y i n the circular arcs a t the stresses shown. Because the circular arcs a r e not expected t o have any post-buckling strength, and they represent over 80 percent of the panel cross section. the onset of buckling also constitutes failure.

Referring t o equation 10-34 of section 10, panel i n s t a b i l i t y f o r a 30- inch panel length may be calculated when J, D3, Dl and kc are formulated a s follows : a 4A2t = - 1 . 0.00922 in? .

PU

where T = effective torsional stiffness

A = enclosed area of tube = 2.488 in?

t = thickness = 0.011 in.

p = pitch between tubes = 2.614 in.

U = circumferential length =: 5.648 in.

(X E correctlon factor = 0.50 27-25

m - 7VsE

- = 0.00177 qSE

2.

D 3 = - = 2 5.2 3 .

4. k, = [ 2 ( + ) ' ( %, + 11 ($)2

Note t h a t the correction factor (yefs. 27-5 and 27-6) i s based on lru#sted t e s t s of corrugation-stiffened p!ix.els performed at bckheed. In effect, it accounts for distortions of the tahes a s a torsional moment varying with the amplitude of the axial wave pattem i s applied t o the tube.

The c r i t i c a l stress for panel instability i s now:

- kc = Dl

f c , w - - 2 t b = 134 600 V S + 93 400 qT, a t 75OF = 98 400 qS + 68 300 qT, a t lk0O0F = 146 TOO psi a t RT = 108 500 psi a t 1400°F A comparison of these stresses with the local buckling stresses calculated e a r l i e r shows the localbucklinz stresses t o be c r i t i c a l , The proportions for the panel configuration were not necessarily optimum since the forming dies were fabricated for panels of a different material.

Test Results. - The room temperature end closeout t e s t specimen failed a t

Because t2.5s t e s t load was be- 44 000 lb, at an average stress of 85 000 psi, low that foi- the crippling t e s t specimen, additional doublers were added t o the compression panel t e s t specimen. mamination of the strain gage data (see figs. 27-49 through 27-52) indicates some local buckl'n.5 along the unloaded 27-26 edges of . a e l a t loads belo: failure, ete exyect2d from the panel b u c k l i q t e s t r e a i t s . Backling o f a tub= a.rc occurred at 43 000 lb, at average Failure Sollove' quickly.

stress of 83 000 psf. The r a t i o of test-to-predicted i n i t i a l buckling stress i s 0.7'3.

The room temperatwe crippling t e s t specimen fai-led at 47 850 lb, at a i average stress of 90 LOO psi. The specimen behaved very much l i k e the end closeout specimen. The strain gage data are presented i n figures 27-74 through kcklfng of a %ube arc occurre& between 46 000 and 47 000 J L b (approxi- 27-77.

mately 88 000 psi). The r a t i o o f test-to-predicted i n i t i a l buckling stress is 0.8k " h e elevated temperature cripglitig test specimen failed without prior local buckling at 31+ 100 Ib, at 8x1 aveltage styess of 66 700 psi. The test data are presented in figures 27-78 through 27-80 and table 27-19. The thermocouples on t h e specimen Irxktcatzd a small thermal g r a a e n t , which when accounted f o r would reduce the predicted stress by a small 9mount. In addition, some detached spotwelds between kubes were observed after the test. It maybe shorn that the buckling stress of the f l a t between tubes, based an one sheet thickness, is 56 000 psi se 1QO0F. This stress i s essentially the same as the local buckling stress Cor the arc of the tube (53 500 pi). Thusj the presence of some de- t a c h d aptwelds .?as probably not s significant influence on the strength of the t e s t specimen. The r a t i o of test-to-predicted initial buckling s t r e s s i s 0.85, neglecting m y t h e m a 1 stress effects.

The room-t%;;:*. - dxre cornpression p e l test failed without prior h c a l buckling a t 40 300 Ib, at an average stress of 73 800 psi. The s t r a i n gage data are presented i n figures 27-107 throyqh 2;-lO9. A photograph of the failed panel .Failure was due t o local buckling of the tube walls; is shown In figure 5-110.

the failu-.*e vis not significantly different from the failures in the previous fests. (See, for example, t h e room-temperatwe crippling specimen a f t e r test, fig. 27-77.) Tie m t i o of test-to-predicted i n i t i a l b u c k 3 stress is 0.70.

% l i s m t i o is belov -those for the previous tests and probably reflects the fack thtiit t h i s test p e l had a somev3at p o r e r quality than the other t e s t panels.

Tkte t h a t the em3 c1osaou.t. specimen, and the tvo crippling specimens were a11 cut from t h e sane 30-511. long panel. Thus, ths quality of these three specimens i s reasombly consistent, and one would exkect their r a t i c s of test-to-predicted initia1bucklir;g stress t o be rather close, wh+,ch i s saen .to be the case.

The elr!vated-temperature compression panel test specimen failed without prior local buc!iling at k? 800 UJ, at an average stress of 80 200 psi. The t e s t data are presented i n figures 27-111 through 27-113 and table 27-30.

Failure was due t o local buckling of the tube walls. The r a t i o of test-to- predicted i n i t i a l buckling stress is 1.02. The relatively large amount of conservatism i n the predicted stress ' . n t h i s case may be due t o the quality of th.e specimen, as d&scussed earlier, or it may he sue t o some variance i n the compressive elastic modulus a t the t e s t temperature. The tendency of Kie material t o thin out i n highly formed areas such as the tub arc requires m e of the least x~zt,. r.al thickness i n the analysis, which i s qitite sensitive t o small chmges i n sheet thickness. Although panel i n s t a b i l i t y was not experienced i n

the compression panel tests, the cicvated temperature specimen reached 74 per-

cent of the predicted p e l instability stress before failing in a localbuck- ling mode. Calculations show that in order for the present cross section to bccome critical in the panel instability mde, the length of the panel would have to exceed 4.9 inches.

Beaded Configuration Analysis. - The test panel drawing is skrown in figure 2 7 - 1 0 .

Again, the

--

test panel. cross-sectional areas presented in table 27-34 are based on the Traverses of the specimens are presented in actual weights of the specimens.

table 27-7, 27-15, 27-17, and 27-28 for the end closeout, room and elevated temperature crippling and room-temperature compression panel test specimens, respectively. The arcs of the beads for analysis purposes are 0 . 0 1 3 i n .

(least measured value) in thickness; the flats between beads are 0 . 0 1 7 in. in thickness for all panels.

Referring to the discussZon of the tubular configuration, the initial bucMing stress of the crippling specimen of the beaded configuration is:

fc , cr = 1 . 7 5 q & ) 1 * 3 5

where = 2 9 x 1 0 6 psi at RT R = 1 . 0 5 i n .

t = 0 . 0 1 3 in. (least measure value) then fC,Cr = 130 0 0 0 psi at RT = 92 500 psi at 2400°F The initial buckling stress of the flat is based on a simply supprted flat plat e : where t = flat thickness = C . 0 1 7 i n . (least measured vaLue} b = 0 . 5 5 6 ina

tl = $ (figs. 27-123 and 27-124)

then

fc,cr - - (3( 500 pei at 75OF

= 7 1 200 psi at lbO°F

27-28 Y As discussed for the tubular configuration, the unloaded edges of the panel are supported by tubes which grip the specimens i n about the center of the available edge width. Therefore, an element w i t h one edge free lies inside This element buckles at room and elevated temperatures at stresses the tube.

substantially below the stressss noted above, and t h i s w i l l very l i k e l y influ- ence t h e strain gage data on tne nearest beads. Local compression buckling i n the f i e l d of the end closeout and crippling test specimens, therefore, is expected to occur i n i t i a l l y i n the arcs o f t h e beads a t the stresses shown.

The arcs are not expected t o have any post-buckling strength, end buckling w i l l also constitute failure.

If equation 10-34 of section 10 is utilized t o predict local buckling which occurred during panel instability t e s t f o r the beaded configuration, predictions for a 30-in. panel length and a 16.37-in. panel width are obtained, which exceed the calculated local buckling stresses reported for the crippling specimens. These predictions, however, are based on the assumption of isotropic cylin-ler type buckling i n the panel, and the beaded panel does buckle i n t h i s meaner when specimens are longer than the c r i p p l k g specimens. Instead, the beads tend t o behave under axial load l i k e plates, w i t h e l a s t i c support provided along their unloaded edges at the crests of adjacent beads. It is apparent that local buckling occurs between adjacent beads l i k e small individual panels, and that these small panels may be analyzed by the proper application of equation 10-34, This is, indeed, the development leading t o equation 12-13. Utilizing 12-13, the folhwing predirtf f o r panel i n s t a b i l i t y is obtained (where the term "panel" refers t o a s i " . ? . . . repeatable element of the beaded configuration): f c,cr n -

X;r tL

where the buckling coefficient is defined by the following equation: where the bending stiffnesses DI, DII, and 9 7 , which a r e defined i n section 12 by equation 12-35, have the following e l a s t i c room temperature values: DI = 5666 lb/in.

DII = 8.13 Lb/in.

D3 = 6.59 Ib/in.

and the effective panel aimensions are: XI = 3 0 . 0 in., panel length XII = 3.085 in. effective panel width measured diagonally f i w n crest t o crest The termPC is defined by: 1/4 Therefore, the minumum buckllng coefficient (kc) is attained f o r a half wave length of 1 5 inches (m = 2)

k , = [ % F 2 (g) + 22 + 4 1 1.8 6

k, = 0 . 0 7 8 5 The area per unit of diagonal width, XI^, of the effective panel betveen crests i s t2 = 0 . 0 1 6 5 3 in.

The c r i t i c a l panel instability stress is then:

- n2 (0.0785) (5666)

-

(3.085 I2 (0.01653)

= 27,900 psi a t room temperature The values shown above were computed by a computer program f o r an arc thlckness of 0 . 0 1 5 in. and a column length of 30 in. A comparison of the above stress with the local buckling stresses calculated e a r l i e r shows the panel instability stress t o be considerably lower. The room-temperature is expected t o f a i l i n the panel compression panel t e s t specimen, therefore, i n s t a b i l i t y mode with a half-wave lengkh of about lj in. Further computations were performed t o determine if t h i s mode might also be c r i t i c a l for the other, shorter t e s t panels. For these calculations, the length of the crippling panels, which were examined first, was t e e n as 7 in. i n order t o allow for the east material a t both ends of the F7ecimens. The panel i n s t a b i l i t y stresses obtained were 72 800 psi a t room temperature and 53 300 psi a t lhOO"F, with the panels buckling into a single half-wave i n the axial diree- tion.

These stresses are based on the assumption of simply supported edges which is obviously conservative f o r a panel buckling into t h i s particular pnttern; a more reasonable approach would be t o set the length of the panels 27-30 equal to the effective column length. Thus for the crippling t e s t specimens, assuming clamped edges, . ! = 3.5 in. For the end closeou-t specimen, = 0.7 (8.5) = 5.95 in., taking the cast edge clamped and the other The room-temperature panel i n s t a b i l i t y stress edty simply supported.

obtained for the end closeout panel is 100 000 psi.

The prediction for the room-temperature crippling test specimen obviously w i l l be higher, and by examination one may see that panel instability for the elevated-temperature crippling t e s t specimen w i l l not be c r i t i c a l .

The end closeout and crippling t e s t panels, therefore, may be expected t o buckle locally and not i n the panel instability mode.

Test results. -The room temperature end closeout t e s t specimen failed a t 24 950 lb, a t an average s t r e s s of 84 600 psi.

Because t h i s t e s t load was below that of the room temperature crippling specimen, additional doublers were added t o the compression panel t e s t specimen. The s t r a i n gage data (see figs. 27-45 through 27-47) indicate that buckling occurred at about 000 l b at an average stress of 74 500 psi. &amination of the failed specimen, figure 27-48, shows failure by crippling at the end of the edge doubler. The back-to-back s t r a i n gages 5 end 6 show a fair amount of local bending across the sheet thickness, probably because of the proximity of an imperfection. A comparison of the data for these two gages with data from gages 15 and 16 shows significantly greater strains for the former pair. It would appear that t h i s is caused by stress concentrations a t t h e end of the doubler between the locations f o r these two l a i r s of gages. The r a t i o of test-to-predicted i n i t i a l buckling stress i s 0.58. This low value is prob- ably due chiefly t o the stress pileup at the end of the doubler.

The room temperature crippling test specimen failed at 32 500 lb, at an average stress of 105 000 psi.

The t e s t data are presented i n figures 27-67 through 27-69. I n i t i a l buckling occurred a t 30 000 lb, a t an average stress of 96 TOO psi. The failed specimen, figure 27-70, exhibits a crip- pling mode of failure. The r a t i o of test-to-predicted i n i t i a l buckling stress i s 0.75, which would imply the panel w a s of reasonably good quality.

The elevated-temperature crippling test speciment failed at 22 100 lb, a t an average stress of 72 200 psi. The t e s t data f o r t h i s panel a r e given i n figures 27-71 and 27-72, and table 72-16. I n i t i a l buckling occurred at about 20 000 lb, a t an average stress of 65 psi. The thermocouples indicated a small thermal gradient which would induce some thermal stress i n the specimen. The photograph of the failed specimen, figure 27-73, shows a crippling mode of failure. The r a t i o o f test-to-predicted i n i t i a l buckling stress, neglecting any thermal stress, is 0.71. If the estimated thermal stress of 4300 psi is incluaed, the r a t i o increases t o 0.75.

The room-temperature compression-panel t e s t specimen failed at 1 3 000 lb, a t an average stress of 42 600 psi, The t e s t data are given i n figures 27-101 through 27-10?. The failed specimen, figure 27-106, shows an obvious panel instability mode of failure, with an axial half-wave of about 10 in, Buckling occurred at about 10 000 lb, which corresponds t o an aTderage stress of 32 600 psi. The r a t i o of test-to-predicted panel instability s t r e s s i s 1 . 1 7 . It i s probable that the actual edge conditions for the t e s t were somewhat b e t t e r than simply support, which, of course would add Sligh+,idr t o the capability of a panel buckling into two t o three axial half-waves.

27-33.

The same summary comments presented f o r the tubular configuration also apply t o the beaded configuration, with the exception that panel instability is much more c r i t i c a l for the beaded configuration than for the tubular con- figuration. This of course i s t o be expected on the basis of the relative stiffnesses of open versus closed sections.

Corrugation-Stiffened Panel Cmfiguration

Analysis. - The t e s t panel drawing i s show i n figure 27-73. The t e s t

panel cross-sectional areas shown i n table 27-34 are based on the actual forming caused weights of the specimens. A s i n the previous conligurations, These variations are thickness variations across the width of the panel-;.

shown i n the transverses presented i n tables 27-6, 27-9, 27-11, 27-22, and 27-24 for the end closeout, room and elevated tempe--ature crippling, and respectively.

room and elevated tenperature compression panel tests specimens, thick, In the following analyses, the sides of the corrugations are 0.011 in.

the crests of the corrugations a r e 0.010-in. thick, t i e attach widths for the corrugations are 0.015-in. thick, and the skin t o which the corrugation i s attached i s O.027-in. thick. The panels are 19.00 in. wide and have 1.0-in.

wide flats a t either unloaded edge. The lengths of the panels a r e the same a s i n the previous configurations.

It i s well known that flat sheet develops varying amounts of post- buckling strength depending upon the configuration i n which it i s used.

Although the determination of i n i t i a l buckling stresses vas the primary purpose of the t e s t s , the panels were taken t o failure, which occurred i n a l l the specimens a t significantly higher loads. Predictions f o r crippling and panel i s t a b i l i t y a r e provided here a s supplemental information t o cor- relate with these failure stresses.

The i n i t i a l buckling stress f o r the sides of the corrugations may be obtained f Tom : where b/d = 0.65/0.82 = 0.793 k = 4.7 (refer t o section 12) c,d = 1.0 Stowell's plasticity correction factor '7*ST d/t = 0,82/0.011 = 74.5 then t 22 200 psi a t RT r' n - ' , t r

- 16 200 psi a t 1400°F

2-7- 12 The i n i t i a l buckling stress for the crests of the corrugations i s based on:

f c,cr = 3.62 Eel ($r

where b s 0.656 in.

t = 0.010 in.

then = 24 4.00 psi at RT f c,cr = 17 800 psi a t 1 4 0 0 ' F The i n i t i a l buckling strength of the skin i s base3 on the equation above: where

b = 2.125 - C . 38 = 1.745 in. (between spotwelds)

t = 0.027 in.

then = 25 000 p s i a t RT f c,cr = 18 300 p s i at 1400"F Thus l o c a l buckling i n the corrugation-stiffened configuration may be expected to occur i n i t i a l l y i n the sides o f t h e corrugation. However, the buckling stresses f o r all of the elements of the cross section, except the flats between corrugations, a r e close enough together that buckling may w e l l occur i n i t i a l l y i n any one of them.

Panel i n s t a b i l i t y f o r a 30-in. panel length may be calculated with The quantities J, D3, DL and k, are defined as follows: equation 10-34.

v-33 .

where A = t h e enclosed area of the corrugation = 0.7906 in.

p = t h e pitch = 2.125 in.

a = correction factor = 0.50 -

D = - - ' - 0.000382 7, E

3 2

= E T = 0.002521VT E

Dl I- 1 2 2

k C = 12 (-$ (2) + I ] ( : I

= 0.649 The basis f o r taking a = 0.50 i s the same as discussed f o r t h e tubular configuration. The c r i t i c a l stress f o r panel i n s t a b i l i t y is now: f

kcV2 9

c,cr = b2 32 900 p s i a t RT 24 000 p s i a t ?!bO°F A comparison of these stresses with the l o c a l buckling stresses calculated e a r l i e r shows t h a t l o c a l buckling should precede panel i n s t a b i l i t y i n t h e com- pression panel tests. However, since l o c a l buckling does not constitute f a i l u r e i n t h i a configuration, the panels are expectec t o sustain additional load and f a i l i n t h e panel i n s t a b i l i t y mode. The effect of l o c a l buckling on panel i n s t a b i l i t y is t o decrease t h e effective s t i f f n e s s of the cross sdction. Studies conducted a t Lockheed (ref. 27-8) on t h i s configuration i n aluyinum indicate, I however, t h a t the effect has s l i g h t influence on. panel ultima-t- capability, k even when Local bucklhg occurs a t one-half of t h e expected ul.Amate load.

27-34 Crippling of the composjte cross section may be predicted using t h e method presented i n LAC Stress M e m o 80c (see also ref. 27-7 for a description of Lnis method). Recause of thz thinness of t h e corrugated sheet at t h e attach- ment j o i n t , and the width o?.' f l a t required i n order t o place two rows of spotc between corrugations, the joint was checked f o r wrinkling i n s t a b i l i t y (using the methods of r e f , 27-9) and found t o be not c r i t i c a l . Therefore, one may expect the configxation t o carry t h e average crippling s t r e s s computed from the above reference. Using the material properties from figure 27-122 and 27-123, and the thicknesses cited prrviously, t h e followil?g average crippling stresses a r e obtained: f = 55 000 psi a t RT CC = $1 500 p s i at l h O ° F CC Note that panel i n s t a b i l i t y i s expected t o occur i n t h e compression panel t e s t s prior t o the onset of crippling.

Test results. - The room-temperature end closeoat test specimen f a i l e d

a t 35 950 l b , a t an average s t r e s s of 47 300 psi. The t e s t data a r e presented Failure occurred at t h e top edge of t h e panel i n figures 27-41 through 27-43.

as pictured i n figure 27-44. Bending due t o t h e eccentricity of t h e end load i s apparent i n t h e s t r a i n gage data a t an early stage of t h e test. The gages show nonuniformities at about 20 000 l b which presumably signaled t h e onset of l o c a l buckling. The The average s t r e s s a t this load l e v e l i s 26 300 psi.

r a t i o of test-to-predicted i n i t i a l buckling s t r e s s i s 1 . 1 9 .

The room-temperature crippling t e s t specimen f a i l e d at 53 TOO l b , a t an average s t r e s s of 69 200 psi. The t e s t data are presented i n figures 27-53 through 27-55. fro^ t h i s data, it may be determined t h a t i n i t i a l buckling occurred at about 26 000 psi. The r a t i o of test-to-predicted i n i t i a l buckling stress, therefore, is 1 . 1 7 . The specimen after f a i l u r e i s shown i n figure A crippling mode of f a i l u r e i s apparent. The r a t i o of test-to-predicted 27-56.

f a i l u r e s t r e s s i s 1.26. Note t h a t t h e specimen a t f a i l u r e carried twice t h e i n i t i a l buckling stress because of t h e post-buckling capability of the corners i n the cross section of the specimen.

a t The elevated-temperature crippling " , s t specimen f a i l e d at 35 000 lb, The t.est dRta f o r t h i s specimen a r e given i n an average s t r e s s of 43 700 psi.

figures 27-57 and 27-58, and t a b l e 27-10. '?bese data indicate i n i t i a 2 buckling took place a t about 30 000 psi, which i s rather high compared t o t h e predicted i n i t i a l buckling stress of 16 200 psi. !!%is d i s p a r i t y is due t o t h e absence of s t r a i n gages i n the elev8ted temperature tests and d i f f i c u l t i e s i n making visual observations in these ~ a m e t e s t s . There can be l i t t l e doubt t h a t some i n i t i a l buckling did take place at a s t r e s s l e v e l which i s more compatible Figure 27-59 shows t h e specimen a f t e r t e s t ; a with the predicted stress.

The ra't?.o of test-to-predicted f a i l u r e stress crippling f a i l u r e i s apparent.

i s 1.05. Again, t h e specimen supported R large load increment above t h e i n i t i a l buckling load before f a i l u r e occurred. A small thermal gradient i n t h e pAnel car, be noted from t h e t e s t data, but it has been neglected i n t h e above comparisons.

27- 35 The room-temperature compression panel t e s t specimen failed a t 32 000 l b , a t an average stress of 39 600 psi. The t e s t data are presented i n figures 27-85 through 27-88. The failed specimen i a shown i n figure 27-89. This specimen had a blow a f t e r fabrication measuring approximately 0.1 in. a t the center of the panel. This, combined with the fact that the end load is attached eccentric t o the centroid of the cross section of the panel, resulted i n substantial bending i n the panel as indicated by the s t r a i n gage data.

Because none of the gages were back-to-back pairs, the onset of i n i t i a l b u c k - l i n g under these conditions was not clearly defined. It is estimated that i n i t i a l buckling occurred a t 20 000 lb, or at an average s t r e s s of 24 TOO psi.

The ratio of t c * ;-tc-predicted i n i t i a l buckling stress, therefore, is 1.11.

As indicated i,- figure 27-89, the specimen failed i n the panel instability mode. The r a t i o of test-to-predicted failure stress is 1.20.

The elevated-temperature compression panel t e s t speciment failed at 25 900 l b , a t an average stress of 32 000 psi. The t e s t data a r e given i n The specimen a f t e r t e s t i s p!.ctured figures 27-90 and 27-91, and table 27-23.

i n figure 27-92. Again, the onset of i n i t i a l buckling was d i f f i c u l t t o deter- mine exactly; from the load shortening curve, figure 27-91, it was estimated.

The t o have occurred a t 14 900 lb, or a t an average s t r e s s of 17 300 psi.

ratio of test-to-predicted i n i t i a l buckling s t r e s s is then 1.07. The t e s t data indicate a small thermal gradient i n the panel, but t h i s w a s considered insignificant i n view of the approximate nature of the t e s t i n i t i a l buckling stress. The long axial half-wave buckle pattern associated with panel instabi- l i t y results i n a specimen a f t e r t e s t which does not show definite indications of the mode of failure a s one would find, for example, i n a crippling failure.

The ratio of test-to-predicted failure stress i s 1.33.

In sumnary, the trapezoidal corrugakjon-stiffened configuration t e s t s and analytical pred.ictions correlate reasonally well, both for i n i t i a l buckling and failure. Conservatism i n the predicted i n i t i a l buckling stresses is due i n some degree Lo the fact that the widths of the corrugation elements ignore the presence of bend radii. The importance of a capability f o r predicting i n i t i a l buckling i s here somewhat reduced, compared t o the two previous configurations, because of the post buckling strength of the flat elements i n t h e cross section of the configuration.

Trapezoidal Corrugation Panel Configuration As i n Analysis. - The t e s t panel drawing i s shown i n figure 27-17.

previous configurations, the panel cross-sectional areas presented i n table 27-34 are based on the actual weights of the panels because of nonuniformities across the panel widths due .f forming. Traverses of the specimens are pre- 27-25, and 27-27 f o r the room and elevated sented i n tables 27-12, 27-lii crippling, and room and elevs. ed temperature compression panel t e s t specimens, ' respectively. The panels we: 19.46 in. wide with a 0.715-in. flat along each vertical edge. Ehd closeout L2lices were simulated In the cornpression panel t e s t specimens by cutting the 30-in. long panel a t a distance of 3.90 in.

from each end, inserting a zee section of 0.020-in. sheet with 0.95-in. flanges, Each end of snd spotwelding an 0.040-in. sheet finger doubler t o each side.

all of the specimens was embedded i n Densite or Pyroform (for elevated tem- perature ta?sts) t o a depth of one inch.

27- 34 This configuration, l i k e t h e previous contlgureztion, is expected t o develop some post-buckling strength because the croEjs section of the config- uration consists of a number of corners. Therefore, both i n i t i a l buckling and failure stresses w i l l be calculated.

The i n i t i a l buckling stress of the wrrugation is: where k = 4.4, the buckling coefficient for b/d = 0.9 c,d (refer t o section 12)

RsT = LO, Stowell's p l a s t i c i t y correction factor

(refer t o section 12) Eel = 29 x 10 psi a t room temperature = 21.2 x 10 psi a t l h O ° F t = 0.016 in.

d = 0.65 in. = t h e widest; element i n the cross section then = 69 600 psi at RT fc,d,cr = 50 800 psi a t lkOO°F Crippling of the trapezoidal corrugation may be calculated using the methods of LAC Stress M e m o 80C. Based on ?he etress s t r a i n data of figures the average crippling ~ w e s : : a t roorn t e q e r a t u r e i s 27-121 and 27-122, G6 600 psi; a t 1kO0F, the average crippling stress i s 64 300 psi. Note that the differences here between initfel. buckling and crippling (failure) are much smaller than i n the corrugation-stlffcned skin configuration.

Panel instability was calculated both f o r the rull panel width, and for a single corrugation (the same as for the circular beaded configuration).

Because of the close spacing of the trapezoidal corrugations and the lack of J. f l a t l i n k f o r hinge between corrugations, the calculated panel instability stress for buckling of a single cmrmgation i s i n excess of 100 000 psi at This stress is substantially large? than the calculated room temperature.

crippling stress at room temperature; thus, t h i s modo is not c r i t i c a l and details are not presented here, The panel instability s t r e s s for the fi.11 panel width may be calculated from equktion 10-34. In performing thepe calculations, it is necessary t o note that the edge conditions along the loaded edges of the compression panels for a l l of the previous confiwrations conformed closely with the assumption of simply supported edge conditions, 27-37 The edge conditjons in the present com- which is inherent in equation 10-34.

pression panels are significantly different; the ends of the panels are cast to a depth of one inch in a matrix, and, in addfbion, a transverse splice is built into the panel at a distance of 3.90 in. from each end. It will be assumed that the transverse members provide the panel with an elastic support.

From an analysis of the stiffhese of this support, an effective panel length L ' may be determined. Thus, examining the splice geometry: where I = 0 . 0 0 7 6 5 i n . 4 (apprctx3mately' for the zee and splice plates Ll = 1 8 . 0 3 i n . , length of the zee and

q I - - " L3 - 126

8EI and x/L = 0.733 vhere I = 0.0215 in.4 for the panel L = 30 in. for the panel x = 22 in., the distance between the zees then c s 4 . 3 from figure c2.26 0% reference 27-3 (fw q = 1 2 6 and x/L = 0.733) and L 30

L ' =: a z - = - . = 14.45

fi 2.075

Now, referring to equations 10-34: where b = 19.g in.

then 75 200 psi at RT fc,cr !Be p e l i n s t a b i l 5 t y stresses are lower than the crippljng stresses calculated earlier; thus, the compression parel test specimens a r e expected t o fail i n the CalculatZons for panel i n s t a b i l i t y in the 8-in. long panel instability mode.

crippling panels yield predictions much higher than the prebicted crippling stresses. These panels, therefore, are expected t o fail i n crippling.

Test results, -The room-temperature c r i p p l b g test specimen faiXed a t 37 600 lb, a i ; an average stress of 92 400 psi. Die test data me p3esented i n figures 2"-60 through 27-62. The failed panel, shown - f n figure 27-63, shows a Initial buckXing occurred at an average stress of crippling mode of failure.

The r a t i o of ?est-to-predicted i n i t i a l buckling stress, approximately 69 600 psi.

therefore, is 1.0. The r a t i o of test-to-predicted failure stress is 1.07.

The elevated temperature crippling t e s t specimen failed a t 26 900 Ib, at %e test data are given as figures 27-64 and an average stress of 66 800 psi.

27-65, and table 2'7-13. The specimen after t e s t , shown i n figure 27-66, ex- h i b i t s a crippling mode of: failure, From the load-shortening curve, it appears that i n i t i a l buckling occurred at about 22 000 lb, at an average stress of 54 500 psi. A small thermal gradient was observed i n the p i e l , but its affect was neglected because of the approximate manner i n trhich i n i t i a l buckling was The r a t i o of test-to-predicted i n i t i a l buckling s t r e s s is 1-07.

determined.

The ratio of test-to-predic+,ed failure stress is 1 . 0 4 .

The room-temperature compression panel t e s t specimen f a i l e d at 29 500 lb, at an average stress of 75 600 psi. The t e s t data are presented i n figures 27-93 through 27-96. The strain gage data indicate i n i t i a l buckling occurred a t an average stress of about 69 300 psi. %e r a t i o of test-to-predicted i n i t i a l buckling stress is 1.0. The specimen, shown a f t e r test i n figure 27-97, failed i n panel instability mode w i t h one half-wave i n the axial direction. '&e onset of i n i t i a l buckling prior t o failure by panel instability maybe expected t o re- duce the stiffness of the panel t o some degree, which has not been taken into account in t h e prediction for -el instability.

Iia this test, t'ne ratio of test-to-predicted failure stress is 1.01.

This ratio is somewhat less than the ratios obkaineh in other tests failing i n panel instability, and is probably due t o the interaction of the i n i t i a l buckling and panel instability modes.

The elevated-temperature compression panel test specimen failed ti;t 19 95C lb, at an average stress of 49 800 psi, The test data are given in figures 27-98 and 27-99, and table 27-26. The loa& shortening curve is reasonably linear up t o t h e failure load, and on t h i s basis, i n i t i a l buckling and failure are considered coincident. The maximum themal gradient in the panel is 2 6 O F ' , which does not appear to be large enougb to be a signzficant factor i n the The specimen after failure is pictured in figure 2'7-lOU; behavior of the panel.

numerous l o c a l (initial) buckles can be seen, On the basis of the crippling %est results, and the loom-temperature compression pmel test Pesult, it is apparent that the configuration has some post budcling strength which may b e limited by panel instability. Since this test specimen d5d not develop any post buckling strength, it is concluded that loss of stiffkess caused by i n i t i a l buckling (and/or the geometric abnormalities) triggel ed premature ffiilure of the speci- men in the panel instability mode. This interaction between modes results i n a ratio of test-to-predicted failure stress of 0.90; the ratio of test-to- predicted i n i t i a l buckling stress is 0.9.

The same summary remarkS can be nade here as were made previously for the corrugation-stiffened skin configuration. It is tspparent i-n comparing the two configurations that the corrugation has less p s % buckling strength.

In adation, the corrugation compressim p e l s a r e nearer t o being optimum than their corrugation-stiffened skin counterparts, since i n i t l a l buckling and panel instability occurred nearly simultaneously in the comugatim compression panels. It is important t o note that there is apprently some %is interaction between these modes when they are close to each other.

interaction r e s d t s in a somewhat lower panel capabilitythan w3en either of these modes i s cri%icalalone.

Circular-Arc CorruF;..ation Shear Panel Configuration eaverses Analysis. -The test panel drawing is sbwn i n figure 27-28.

of the two room-temperature test specimens are presented in tables 27-32 and 27-33, which indicate tlwt the specimens may be considered t o be of uniform The analysis for these thickness, namely, O . C l 5 ; and 0.0145 in., respectively.

bitialbuckling, specimens also covers both i n i t i a l buckling and failure.

which may be expected to occur i n the circular arcs, does not necessarilymesn Therefore, analyses for panel that the panel cannot carry additional load.

instability and web rupture are also presented.

The shear stress for initial. buckling may be calculated fi*om equation 11-6 of section U.

27-40 where

= 29 x lo6 psi

= O . O P j l in. and 0.0145 in. for t e s t specimens 1 and 2, t respectively R = 0 . 8 0 in.

then = 41 200 psi for- 1 : = 0.0151 in.

fs , cr

= 38 700 psi for t = 0 . 0 1 4 5 in.

Note that equation ll-6 is based on extenpive tests and is applicable for cor- rugation half-angles between 20 and 90 deg. I t is assumed that t h i s equation a applies both t o initialbuckl5ny of th- arcs of the corrugation and t o buckling of the corrugation between aaacent arc crests, should t h i s latter mode occur within the range of corrugation half-angles cited.

The shear stress for panel instability may be caiculated from equations 10- 36 through 10- 37b : = fs,cr i . b2 t where Ks = 3.3 (from fig. 10-8) a = 17.00 in.

b = 15.62 in.

D1 = 6.86 lb/in.

D2 = 46 000 lb/in. for t = 0.0151 in.

D 3 = 12.14 lb/$n.

= 6.08 lb/in.

@ = 44 200 lbfin. for G = 0.0145 in.

D3 =: 10.73 1b/in.

then f S , C r = 44 200 psi f o r t = 0.0151 in.

= 43 300 psi for t = 0.0145 in.

Data are presented i n NACA TN-2661 (ref. 3-10) for t h e allowable web gross area shear stress for two aluminum alloys as a function of the diagonal tension factor k. It may be shown that approximate values foT other materials may be obtained by multiplying fs max f o r 2024-Tw aluminum by the r a t i v of the ultimate tensile stress of the ne4 material t o the ultimpe t e n s i l e stress of Taking Ftu = 165 000 p s i f o r Rene kl and k = 0.1, fs,m&x 2024-T3 (62 000 psi).

for the shear panels is 68 000 psi.

The anzlysis shows i n i t i a l buckling and panel i n s t a b i l i t y occurring rather close together; one might expect, therefore, some interaction between these modes modes.

Test results. - The room-temperature shear panel test specimens failed at

9500 l b (t = 0 . 0 1 5 1 in.) and 8700 l b (t = 0 . 0 1 4 5 in.).

These loads represent average shear stresses of 40 500 p s i and 38 400 psi, respectively. The test data are presented i n figures 27-114 through 27-120. From these data, it appears that the thicker specimen buckled locally at an average shear stress o r about 38 500 psi. The specimen carried only a small additional increment of load before failure. The thinner specimen showed no signs of i n i t i a l buckling prior t o failure. Both specimens developed the panel i n s t a b i l i t y mode of fail- ure, followed by rupture of the web (see figs. 2"-117 and 2 ' 7 - 1 2 0 ) . The r a t i o of test-to-predicted i n i t i a l buckling stress f o r the t w o specimens are 0.93 (t = 0.0151 in.) and 0.99 (t = 0.0145 in.). !Fhe r a t i o s of test-to-predicted It is apparent t h a t the nearness failure stress are likewise 0.92 a d 0.89.

of t h e i n i t i a l buckling and panel instability modes i n these specimens re- sulted i n some interaction between the modes, which lowered the capability of the panels. The rupture of the webs is considered t o be an aftereffect of primary failure i n the panel i n s t a b i l i t y mode.

Spar Cap Configuration - 1

- The t e s t specimen drawing is presented i n figure 27-33; thick-

Analysis.

ness measurements are recorded i n table 27-21. mese measurements indicate a Analyses for i n i t i a l buck- cap thickness of 0.058 in. i n the region of failure.

ling and crippling of the cap follow.

I1

I n i t i a l buckling i n compression of the cap may be calculated from the equation : 27-42 where

vT = 0.99 (see f i g . 27-125)

= 29 x 10 6 psi E e l t = 0.058 in.

b = 1 . 7 9 in. = the maximum unsupported distance i n the cap t h e edge bend radius between the corrugated web and then = 110 000 psi f c,cr The crippling stress as determined from LAC Stress M e m o 126* is: fcc An, lb An, in.

EXement

b/t or ( R / t ) ' psi n n

(0.192 x 0.056)'c 0.02227 3.31 5 6 0 0 0 " 1247

(0.125~ x 0.058)2 0.02806 (2.65 1 55 200 1549

(1.607 x 0.058) 0.09321 2 7 . 7 29 100 2712 (0.777 X 0.058) 0.04507 1 3 . 4 5 7 000

-

c 0.1886 8077

* The LAC Stress M e m o Manual recommends Stress M e m o 120 for the crippling analysis

of single sections, and Stress M e m o 80c for the crippling analysis of stiffeners attached t o panels. The use of Stress M e m o 80c here would yield a lower average stress, namely, 114 000 psi. Strc3s M e m o 1 . 2 6 u t i l i z e s the unit material approach; M C F i s the material correction factor.

* One edge free; other f l a t elements have no edge free, Assume F,y/Ec = FtJE - - = 0.00503 146 000/29 000 000 0.0214 (fig. 1 5 of LAC Stress M e m o 126)

3 I l =

MCF = 0.0214 x 146 = 3.12 then - - 8077 3*12 = 133 500 psi 0.1886

Test results. - The spar cap crippling specimen failed a t 48 000 lb, a t

an average stress (for two beam caps) of 127 200 psi. The t e s t data are given i n figures 27-81 through V-84. These data show i n i t i a l buckling occurriq; at an average stress of about 104 000 psi. The r a t i o of test-to-predicted i n i t i a l buckling stress is 0.95; the r a t i o of %e&-to-predicted failure stress i s also 0.95. Using the more conservative crippling analysis of Stress M e m o 80c, rather than that of Stress M e m o 1265 results i n a test-to-predicted failure stress r a t i o of 1.12. In t h i s analysis, the crippling stress for the element f o r which i n i t i a l buckling i s calcidated above is 83 500 psi. This value is probably con- servative; on t h e other hand it may be optimistic t o consider t h i s element t o be simply supported along both unloaded edges. In summary, the predicted stresses are somewhat high and consideration should be given t o the use of more conservrz- tive methods, 27- 44 Wilhelm, K . A.: Preparation of Sub-Elements for Materials and 27-1 Process Developnent Tests. CMRI-1465, bckheed-California Co., 1968.

27- 2 Panik, J. J.: Spot Welding of Structural Elements for Hy-personic Cruise Vehicle Wing Structure. CMIiI-1465-2, bckheed California Co. , 25 July 1968.

Bruhn, E. F.: Analysis and Design of Flight Vehicle Structures.

27- 3 Tristate Offset Co., 1965.

Donnell, L. H. ; and Wan, C. C. : Effect of Imperfections on Buckling 8- 4 of Thin Cylinders and Columns Under Axial Compression. Journal of Applied Mechanics, March 1950, pp. 73-83.

Lboley, J. F.: On the Torsional Stiffness of Closed-Section W e b Int. J. Mech. Sci., 1965, Vol. 7, pp 183-196.

Stiffners, Bushnell, D. ; Almroth, B. 0. ; and Sobel, L. H. : Buckling of Shells of Revolution with Various Wall Constructions. Volume 11, Basic Equations and Method of Solution, NASA-CR-1050, bky 2 . 9 6 8 .

Cozzone, P. P. ; and Melcon, M. A. : %ndi.mensional Buckling Curves -

27-7 Journal of Aeronautical Their Developnent and Applications.

Sciences, October 1946, pp. 511-517.

27-8 Holmes, A. M. C.: Compression Tests on Large Panels Having Corrugated Reinforcement. W C F-70-69-2, June 1969.

Semonian, J. W.; and Peterson, J . P. : An Analysis of the Stability 27- 9 and Ultimate Compressive Strength of Short Sheet Stringer Panels with Special Reference t o the Influence of Riveted Connection Between Sheet and Stringer. N A S A TN-3431, March 1955 3-10 Kuhn, Paul; Peterson, J. P.; and Levin, L. R. : A Summary of Diagonal Part I - Methods of Analysis, NASA n-2661, M a y 1952.

Tension.

27-45 TABLE 27-1 STRUCTURA L-E LEMENT TEST SCHEDULE Number of panels tested

I

Compress, End Inplane crippling Type Of test closeout pane1 shear Te st-panel configuration RT - Tubular I Beaded

I1

- Corrugation-stiffened

--

Trapezoidal-corrugation Shear web Channel cap

?zT

3 5 3 Total number of panels

:I

Grand total

-

27- 46 TABLE 27-2 SUMMARY O F PANEL ELEMENT FABRICATION Pane 1 Pane 1 Panel s i z e , NO. of panels d e s c r i p t i o n tY pe i n . f a b r i c a t e d End closeout 9.0 x 17-37 1 8.8 x 17.37 2 Tubular C r i p p l h g 30.0 x 17.37 2 Compression -ne1 End closeout 9.0 x 17.37 1 Beaded 8.0 x 17.37 2 Crippling Compression w n e l 30.0 x 17.37 2 Corrugation End closeout 9.0 x 19.00 1 8.0 x 19.00 2 s t i f f e n e d skin Crippling Courpression panel 30.0 x 19.00 2 I Trapezoidal Crippling 8.0 x 19.46 2 corrugation Compression panel 30.0 x 19.46 2 C i r c u l a r a r c Shear 1 5 . 6 2 ~ 17.00 2 corrugation Channel cap Crippling 5 . 5 0 ~ 2.75 x 1 T o t a l No. of panels 22 27-47 L! C I 0 0 0 0 0 0 0 0 0 0

I

4 O L n 400 400 0 0 0 0 0 0 rl (u (u 0 0 0 0 0 0 d.3 0 rli-Q .3 0 d O l n

+ 0 0 0 0 0 0 4 O v \

rl rl dC00 4 x 0 0 rl30 0001 0 0 0 - "'fl 0 r( d i

I '

0 0 0 d N 0 Fin0 o o c cu cu . . .

. . . . . . . . .

. * .

0 0 0 0 0 0 i 2 g g W B B G i 2 G G i i G 2 2 M

E

k m d I m TABU 27-4 ?

MECHANICAL PROPERTIES DATA FOR S O M E RENE $1 COMPRESSION PANEL MATERIALS SlTBJECTED TO VARTOW TIIEWAL CYCLES

-

Test panel Tiibuhr and Beam cap Beaded psnel e m f iguration corrugation-st iffened c r i p p l i n g and skin panels shear panels Eleme2t of panel Corrugations for both Bead caps configurations Material gage, in. .016 .01g .a50 Grain d i r e c t i o n Longitudirial Heat No. IIT-24%- 7-8 513 .

Thermal cyclea Exposed t o three anneal Exposed t o 2 Aged cycles and a.ged anneal cycles and aked Coupon t e s t 14OO0F

I temperature

I RT

Frope rt ies

I Mechmical

I

1 8 3 .

Anneal cycle: HeatEd t o 195OoF f o r 15 minutes, air cocled; then aged 14-00 F for 16 hours and a i r cooled Aging cycle: Heated t o I.40O0F f o r 16 hours and a i r cooled 27-49 P P v w a , rl

ii

.rl m m

E

PI

E

a a, rb a _.__

I 27-50

U - 1 D N N

l -

- 4 - !-. Y ~

b c

27-51 V m 4 +---I-+-- !h m - 0,

- - - ? - -i

1 1 27-53 i

If

-

- - ? -I 0.

i I m < !

I 1

27-54 P TABLE 27-10 TEMPERATURE DISTRIBUTIONS FOR CORFUGATION B T I F F E N E D CRIPPLING PANEL C k 152 cn . j 4 CU 156 CW 157 C n 155 7-4 1-6 1-7 1-2 7-5 OFG F OCt F 3EG F O t G F DEG F DEG f 1348.7f 1322.6 1455.9F 1*16*0F 1368.2F 1416b5F 137103F l747.5f 1326.4 1455 4F 1415.6F 1366.5F 1407.3F 136502f 1346.6F 1 ~ 2 7 0 2 145RelF 1417.3F 1371.3F 1 4 1 2 6F 1370.0F 1 w n . i l459.5F 1480 *F 1370.OF 1418.2F 137C 8F i349.1C 1779.3 1‘63.0F 1423.4F 1373 *9F 1417-8F 137703F 135600f 1330.1 146P.6F 14P?*OF 1370 . O F lb1309F 1367 3F 1353 9F 1460 *F 136806F 1 4 1 3 ~ 4 F 1369.5F a 349.5f 1330.9 1421 *7F 1461 3F 1615.2F 1331 * 7 14Pl.7F 136806F 1367 Ff 1351.3f 1732.6 137t.31 1466.5F 1425.6F 1369.5F 1418 6F 1356eOf 1333.5 1467 s 3F 142R.PF *.375*6F 142103F i 3 7 n . b ~ 1355.6F 17’17.9 1462.6F 1423 4F 1371 07F 1 4 l * r 7F 13f0.4F 1352.6F 1334.5 1469 I F 1427.3F 1376 -9f l*21=3F 1372.lC 1348.7f 1335.3 1469 1F 1427.3r :377 r3F 141806F 1374 7F 135609f 1736.0 1466.OF 1373.3F 1*16*9F 137304f 1356.7F 1426 -5F 1736.9 1465 6F 1373.0F 1371 *7F 1351e3f 1426.OF 1412.6F 1977.9 134803f 1466.9F lb26*% 1373.0F 14 15.61 1375aPF 1338.7 l46R.6F 14?9.lF 1376*0F 1 416.9: 1371.7F 1349m S F 1339.7 i 4 7 o . n ~ 1431.7F 1375.6F 1*12*1F 1374 3F 135201f 1740.7 1471.3F 1431.3F 1379.lF 1412.1F 1376.5f 1352.1P CU 159 cn 160 Cu 161 CM 162 CH 163 1-8 1-9 1-10 1-11 7-12 7-13 DCG F DEG F OEG F DEQ f DEG f DEB C 00’ 1325rb 1370.0F 1357.RF 1350 2F 14Oh3F 1376oOf 1347.1f 13Cb.4 i370.w 140309f 137107f 004 1354*7F 13b5.6F 135309t 111P7.2 1370.3f 1354 3F 1407*3F 006 135600F 1378-2F 1360.8l 000 1361.7F 1377.8f 132801 1372 1F 135708F 1608*2F 1352. b I 1376.0F 010 139903 1361afF 1360.8F 1*08*6F 1303.7f 1350-$C 012 1330.1 137304f 135600F 1357.8F 1406eOF 1376eOf 1313. W 014 1330.9 1372 6F 135506F 1319. I F 1403.9F 1376.5F 1349e5P 016 1531 e 7 1372.1F 135600F 135708F 1402 6F 137403f S3Slo7C 1332.6 137506f 1360 . O F n t n 1362.6F 160703F 13780Of 136Oe8F 020 133305 1375 .hF 1363.0F i 3 n . n ~ 13S80Zf 1362 o6F 1405*Pf 020 11133.9 1371.3f 1760eOr 1360.4F 1406- OF 1379oSf 1356 e 9 f 1334 05 1378. 6F 02’ 1357 8 3F 1356.0F 1410*OF 138504F 1342091 1735.3 1373.0F 1361.7r 024 1364.3F 1406.5F i 3 n 3 . 3 ~ 1 319 o l C 026 1336 e 0 ‘. 374 I F 136Oe4F 1361.7F 1406.5F i 3115.4~ 131309f O P R 133609 1373 *9F 1359-1F 1357.8F 1406.OF 130102f 1343e7f 0 3 0 1737.9 1374.7F 1356.5F 13R4.5f 1339.W 1362.1F l*O*.7F 0 32 1338.7 1371.3F 1359s 1f 1360.OF 1 378 6F 13S0.41 1403 OF 034 1739.7 1374 7F 1360 *OF 136008F 1379. 1F 1354eJf 1406eOF 0 35 1740.7 1374e7F 1355e6t 1360 *Of 136506F 1409.1F 1380 4 f L A I D T I M E CH 164 CM 179 CM 165 CM 180 CH 101 CH 182 7 - 1 4 1-15 ~ v 0 T - 1 LvOt-2 LVI)T-3 LvDT-4 OEG F DEG F INCHCS I NCWES I N C H E S !NCMES 002 172506 1410.4F 1393*0F 0.OOOJ 0.oooJ 0 0 0 0 0 J O r 0 0 0 J 004 132604 1411.7F 1393.4F 0.004 3.002 0.302 0.002 006 1327.2 14OR.PF 1401 r7F 00004 01004 0 006 0.005 oon 1 3 ~ 8 . 1 1400.8~ 1399elF OoCO7 0.007 c. 307 0.007 013 1329.3 1419.1F lkORe6f 0 005 0.008 0 006 0.007 01% 1330.1 14lO.4F 1399 .SF 0.009 0.009 0.006 00009 014 1130.9 1418.2F O.0OP 1400*8F 0.009 0.011 O e O l l 0 t h 173117 1416eOF 1401 s3F o*o1a 0.011 0 0 009 01010 014 1332.6 141RobF 140609F 0*011 u.011 0.011 0.010 023 1337#5 142704F 1406 *5F 0.010 c * o 1 ? 0.010 0.013 0 . 2 9 133309 142O*OF 1408.2F 0.009 Oe012 0.010 3.011 02% 1334.5 1412.6F 14130CF O O C l l 0.013 0.012 0.011 324 $335.3 1427.3F 1410 *OF 0.013 0.313 0.010 01013 026 1336.0 1422.6F 1409 *sF 0.013 0.015 0.315 0.013 O2R 1776.9 1425.6F 1411*3F 0*@16 d.015 0.017 01015 1413.4F 079 1337.9 1415 r2F 0.013 00019 O * O l 8 0.017 032 133n.7 1 4 ~ 1 . 5 ~ lSO9elF 0.018 Oe020 01024 6.018 034 133q.7 1426e5F 1410 o8F 0.020 C.023 0 025 0.023 93s 1?40.7 142117F 141 2 * 6 ~ 0.037 00039 0.337 Os038 27-55

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i 27- 57 TABU3 27-13 TEMPERATm DISTRIBUTIONS FOR TRAPEZOIDAL CORRUGATION CRIPPLING PANEL LBAD CW 181 cn 1SZ cw 155 CH 156 TC- 1 T C - t TC-5 TC-6 oca c ora F DEG F OEO F i*eo.rr 1382.W 1400s8F 1402.6C 1481.71 1319eW 1396 8 ZF 139905F 1431 *7C 1376*OI 1392.w 1395 4 1 1431 a 3 1 1 YY 9 . s 1 1403. O F 1397S3P 143Y*51 1388 031 1388*7F 1395e2F 140PeOC 1441 1401 r3F 1400aBF 1391 *31 1387.SF 1418.2r 1393.OC 1388. I F 14OOeW 1391 OF 1431.3: 1 bOi? 61 1392. 1F 1 3 9 8 0 2 t 1403.OC 1433.4F 1403.41 1393.OF 1403.4C 1401 3C 1 4tO.SC 1399.1C 139&lF 1355.6F 14OOe8F 1400.OF 1430 *4F 1404.3F 1396 *OF 1362.6F 1 bo7 8F 1401 I F 1433.41 1408.6F 1395.2F 1359.1F 1408. PF 14OO.bF 1427 8 : 1401 e3F 1392.6F 1358.2F 1405 6F 1404.7F 1433.9: 1401.7t 139O.Of 1353.4F 1400.bF 140OeOF 1 b33 *9F 1397 3F 1391.7F 1359.5F 14OJ.OC 1403.OF 1436e9F 1410.0F 1393.9F 136te6F 1406r9F 1408.2F cu 158 CH 159 C H 160 C W 161 CH 162 TC-7 Tc-n I C - 9 T C - 1 0 TC-11 TC-12 D€G F DEG F DEG F OEb F OECI F OEG F 002 1430.3 i 3 7 o . n ~ 1394.7F 1389elF 1401 7F 1415e2F 1350. a$ 904 14310 2 136309F 1394.7F 1392.1F 1403.OF 1416rOF 1357.8F 1431.8 1 369 5F 006 1392. lt 138705f 1400.4F 1419.lC 135304F 143P.5 1 3611 6F 1393.OF 008 1393.9F 1600 8F 142502C 1365r2F 010 1433.3 1375.21 1795 6 r 1YPIo5F 1382.9F 1402.1F 1361 7F OI? 1434.0 137004F 1395 *bF 1391 *7F 1406.5F 14 1905F 1356r9F 143S.1 1366*9F 1393 *OF 1392 r6F 1354r3F 014 1402.1F 1*10*8F 016 1436.0 136703F 1395.2F 1396.0f 1402.1F 141304F 135309F I b 3 7 . 0 o1n 1370 *OF 1391.9F 1389.5F 1403 9F 141905F 1 3 5 9 0 l F 1437.8 1369 -5F 1393eOF 020 139304F 1 400.4F 14l206F 1350.4F i * 3 n 0 5 1394.7; 1417e8F 1356.OF 021 1370 n~ 1390 *8F 1400 4F 022 1439.3 1393 *9F 1407.3F 1366.?~F 1392 6F 1423.9F 1365r2F O t 3 144609 1373.4F 1395 -?F 139609F 1423.0F 1.756.9F 1404 7F 0 P4 1441 06 137) 07F 1394e7F 1390 8F 1401.7F 1413.4F 1356m5F 1442.7 025 1373 4F 1396 .OF 1391.3F 1402 * I F 141107F 1353.01 1443.4 026 1373.9F 1395 . 6 t 139708f 1 4 ~ 7 . 3 ~ 1416.9F 1362.6F CU 164 cn 165 Cn 179 cn 180 cu 181 LVOT. LVDTo LVDT.

PT-2 PT-3 TC-14 P f - 1 TC-13 TC-15 OEG F INCHES INCHES INCHES OEG F OCrJ F lC6008F 1461 r7F 0.000J 002 1395.61 0.0004 0.OOOJ Or007 0.008 0 006 004 14OO.IIF 146008F 1461.3F 0.014 0.OlP 0.010 006 1402elF 14S8 I F 1*63*4F 1470**F 01013 0.014 0.012 oon 14 I 3 -9F 146708F 1 4 0 n . 6 ~ 1464.3F 1470. 4F 01016 0e016 0 . 0 1 3 1463.0C 1467.3F 0.019 0.018 0.016 o l e 1402.1F 146o.n~ 1463.9F 0.021 0.32P 0.017 014 1396.9F 0.019 o i a 1396e9F 166306F 1466.3F 0 . 0 2 2 0.023 t l 023 1401 e7F 1 0 6 5 . ~ r 1471 .3F 0 024 O r O Z 7 01n 3 8 026 0.026 0.025 7 20 1397 7F 1664.7F 1467r3F 0 e 028 0.02A 0 026 021 1399.5F 146R * 2 F 1469mlF 0.027

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023 1406 *OF 1469.5F 1473 o4F 0.031 0.029 O f 4 1399rlF I 460 *OF 1464 e7F 0.031 1465.ZF 0 0 0 3 3 0.032 02s 1395.6F 1463.4F 0 033 6.039 0 035 026 14OP.lF I 4 6 9 1F 1469.1F 0 035 27- 58 L i I 1 ! ? !

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27-60 W L E 27-16 TEI4lW?&E% DIST.!UBUTIONS FOR BEADEB CRIPPLING PANEL cw 152 CH 154 Cw 155 Cu 156 CH 157 1-1 1-2 T - 4 1-5 7-6 1-7 DEG F O F G F OEG F DEG F DEG F DEG F 11 16.8 1378e2F 1375 *2F 1407e3F OO? 1387 *OF 1412 .bF 1392e6C 1117.9 1387e5F 1396.5F 1376e3F 1*14*7F 004 1 *08 6F 1390 e4F 1118.7 1380.4F 1381*6F 1373e4F 141 1 7F i 406eSF 1 3 9 4 e 3 F 1119.6 1460 *4F 1369 *5F oca 1391 * Y t 1414.YF 1406*OF 1392e6F 008 1122.6 1394 e?F 1401*7F 1372 6f 1*06*5F 1403*9F 1385eOF 010 liP2e8 1386e2F 1 1 0 0 e O F 1376e5F 1 4 12 *6F 1410eOT 1390e8F 01P 1123.9 1390 e OF Z 386 e 2F 1381.2F 1412elF 140beSF 1394 e 3F 014 112b.7 i 4 0 2 *6F 1385e4F 1378.2F lul1.3F 1406eOF 139Qm8F 016 1?25e5 1386 *2F 1403 *5F 1403.9F 13R8a7F 1390*4F 136heOF 1388.7F 1368*6F 1411e7F O l P 1126.2 ?389*5F 1421.3F 1394o’IF 020 1127.0 139% 6F 1390 4F 1371.3F 1 &lO*ClF 140107F 1390e4f 021 1127.7 1378e2F 1393 *9F 1389elF 1410.4F 1403*0F 1394 e 3F 022 lli’0.4 1413eOF 1399.1F 135605F 1406*5F 1390 *F 1382e9F CH 159 CH 160 CM 261 CH 162 CU 163 1-9 t-12 Y-13 1-8 T-10 T - 1 1 OEG C DEC f DEG F DEG F DEG F DEG f 1116.8 1400 a 4F 1394 7F 1401*3F 1403 4F 1408 6F 1403eOF 1117.9 1393*9F 1403eOr 004 1392.1F 1 403 9F 1407.8F 1421e3F 1118.7 006 1397 e8F 1397*8F 1406*5F 1410a 8F 1412elF 1396e9F 008 lS19.6 1399mfF 1398.6F 1407e3F 1*12*1F 1423 OF 1406eOF 1121e6 139Ra6F 003 1395 e 2F 1400e8F 14lO.bF 1423. O F 140309F 010 1122.8 1399elF 1395e6F 1399mlF 1406eOF 14 1 4 e 7F 1406e5F 012 1123.9 1397.3F S 400 *OF 1407 08F 1407.8F 1402rlF 1410 4F 11.Z4.7 140?*6F 1395 *2F 1393-3F 1412*6F 014 1404- 5F 1399elF 1125.5 I393 *OF 1406aOF 016 1393.4F 1401*7F 1405*6F 1416mOF 01n l l t 6 . 2 1396e5F 1392 *6F 1409*5F 1404 ?F 1*C%*3F 1406aOF 0 20 1127.0 1395 6F 1401e3F 1407*8F 141Oe4F 141; = 1F 1396eOF 1404e7F 021 1127.; 1 4 G O *4P 1403e4F 1410 4F 1 I 16 5F 1402alF 02P flioU.4 1392.1F 1399elCI 1402*6F 1402*lF 1414 o 7 F 3402. IF CH 181 c w 182 Cw 165 CH 179 CH 180 7-15 LVOT-3 LvDT-4 LvDl- 1 LVDT-?

DEG F I FICHES ! W H E S INCHES INCHES 1116.8 0e000J o*ooo J 002 1428ePF 0.000J O*oooJ 0 a 004 004 1117.9 1433 *P 0.004 0.004 0 . 0 0 4 006 1518e7 1*27*8F ocoo9 0.01C 0 . 0 0 9 0.008 1119.6 1 4 4 1 a T t 0.312 0.010 008 O*OlO 0.013 llE1.6 O V O l l Ob8 i * % ? * l p 04011 0 ~ ~ 1 4 0.013 010 llL2.8 1434e3F 00015 0.014 0.015 0.057 012 1123.9 0.020 Ob018 1427 e 3F 0.019 0 e 022 014 lie4.7 1430o4F 0 622 0.025 0.024 0 b 021 016 11tS.5 0.030 0.026 0 025 1432.1F 0.026 1126.2 O r ’ 3 3 P ob029 018 1433e9F 0.027 0.034 020 1127rO 0.034 0.040 0.038 0.034 1432e6F 021 PlP7e7 0 a 040 0.036 l435e2C 0.036 0 042 1128.4 1430*0F 0*041 0.049 0.044 0 e 042

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a Q 27-62 - T e . -- - ... - I - . . -- 4 * 27-63 T4BU 27-19 TEMpEFVAVR?3 DISTRIBUTIONS FOR TUBUUR CRIPPLING PANEL LOA9 cu 152 CH 154 CH 155 CU 156 C I ' 157 7-e I - 4 T -6 1 - 7 7-5 CEG F OCC F OEG F OEO F OFG F 1427.8F 1350.4F 002 1362-6F 1413e9C 141 * . 7 F 1435 *2F 1353 .4F 1364 3F 141 7 03F 004 i421.7r 006 14P6.5C 1350- BF 1360eOF 1400.5F 1412mlF 008 1431e7F 1345.0F 1361.3F 1405e6F 1*06*OF 1434.7F 134SoOF 1360oHF 010 141 0.4F 1 4 1 0.4F O l P l430sOF 1351.3F 1363 PF 1408.6F 1412.6F 14 me0F 014 1352.6F I 3 7 O * W 1415.6F 14090 1F 016 l b 3 l e 7 F i 3 r n . 7 ~ 1410.4F 135703F 1407eIF 1354.7F 01- 1473rOF 1346 *6F 1403eOF l40609F 1436e5F 1353 -4F 020 136201F 141708F 1 4 2 5 d F 0 P2 1431 e3F 1350.4F 1358 6F 1b10.4F 1415-2F 024 1433s4F 1345.4F 1360 8F 1409.1F 1414e3F OZb 1433m4F 1349.5F 136403F 1409 -5F 14lte6F OP8 lC33-9F 133806F I 364.3r 1402.6r 1403.4F 0 30 lb35.2F 1340eBF 136201F 14'26.SF 1406.5F 031 1438 2F 14 15 e2F 135502F 1360cBF 14i6eOF 032 jb38-2F 1326 e6F I 356e5F 1417e3t 1411-7F 033 1432-1F 134004F 135806F 1399.1F 1405e6F 034 143?*1F 135201F 1365.2F 141905F 1423 9F L W I tu 159 cn 160 Cd 161 Cw 162 CH 163 TO9 1-10 1-11 1-12 1-13 DEQ F OEG F DCG F DEG F DFG F oor 1373.4F 140309F 1363-9F 1349e5F 1394.3F 1366.9F 004 137806F 1405 6F 1372.6F 1395-6F 1 3 7 0 e O F 1400.43~ 1358. t F 1344.1F 138R.7F 008 136806F 1396.9F 1364 -3F 134405F 139004F 1379elr 1356.W 010 1403.OF 1366.CF 1389.1F 012 1370e4F 14OO*OC 136506F 1354e3F 13860 6F 7 C 014 1374 3F 1399.1F 1364 1356.5F 1390.4F 016 1376*% 1402-IF 1368.2F 1353e4F 1388.3F oio ¶374*7T 1398 e6F 1366.0F 135407F 1390 O F or0 1385 *OF 1407- 3F 1373.4F 1361.3F 1393e9F or2 1380eOF 140304F ~ 66*5F 1360r4F 1388.7F 1377.3F 1367.5P 1388.3F 024 1401*7F 1363.4F O S t 3 7 7 . 8 ~ 1401 *7F 1366eSF 1354.7F 1390eOF 1370.4F 1394-3F 1360.4F 1343-3F 139004F 030 1375-2F 1400 S f 1365.6F 1353.9F 1390 e8F 1384.5F 031 l405*2F 1369.5F 136004F 1391.3F OYC 1380 w 4F 1406.5F 1374 3F 1355 * 2 t 1389olF 033 1374 e7F 1*02*6F 1368e6F 135605F 1367.9F 034 1380*3F 1405 *6F 1370 I F 136004F 1392alF cw 181 CII 182 CH 165 CH 179 Cn 180 LOAD ~ ~ 0 7 - 3 LvDT-b T-15 LvDT-1 LvOT-2 1 NCHES INCHES OEQ F INCHES INCHES 0.000J 1409eSr 0100OJ 0 e 0004 0.000J 0.000 0.001 1417.8C 0 . 0 0 1 . 0.002 0 . 0 0 5 Os007 0.005 0.006 14lO.OF 0.008 1409o'Jt Oe007 0.010 0 - 0 0 6 0.011 0.008 0.009 1414*1F 0 * 008 141JeSf 0.011 0.010 OlF 0 009 0.013 0.015 0.012 Olb ¶ % i 0 *bF 0.011 0.012 0.012 0-017 0.013 0.014 016 1411 *7F 0.015 1417-3F 0.013 0.018 01015 o i n 0.014 0.017 0.017 010 1+15*ZF 0.019 1417.3F OIOl9 0 023 0.019 0.018 0 t 2 0.02P 0.020 1416.5F 0.018 00023 0.020 0.026 0.024 0.021 026 1422.6F 0.023 1419.1F 0.021 0.028 0.075 02n 0.084 1421.3F O e O P 3 0.030 0.027 3 30 0.024 0.027 0.025 031 14%?rlF 0.031 1420rOF 0.024 0*03? 08029 0.026 m i ?

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I , . - 27-69 I - - - - - X I 27-70 TABLE 27-26 TEMPERATURE DETRIBUl!IONS I%R TRAP%ZOII)AL CO€?RUGATION COMPRESSION PANEL LOAD T l M E CU 47 CU 48 CH 4 9 CH 50 CH 51 CH 52 L V O T - LvOT LVOT .

LVOT.

PT-2 P1.1 Pt-3 PT-4 T C - 2 TC-1 1NCWES 1 NCHES 1 NCHE S 1 NCHES DEG C DEG F 2231.3 0.0004 O*OOOJ 0.000J 0.000.J 1340eJC 002 1337.3: 004 2232.1 0.010 3.012 0.010 0.010 1337.3C 1339.5F 006 2232.8 0.018 0 020 0.018 0.010 i338.6: 1339.1F oon 2233.5 0.024 0.026 0.026 0.026 1337.3: 1340.W 010 P234.e 6**@30 0.033 0.033 0.033 1338.2F 1339.5F 012 2E34.Y 04037 0.039 0.041 0 040 1337.8F 1 3 h 1F 014 r m . 6 0.043 0.047 0.047 0.047 1339.5F 134 1 e6F 016 2236.3 0.049 0 8 054 0.053 0.054 1339.1F 1341.2F o i n PF 2t36.8 0.056 00060 0.011 0.061 133806F 13b1 019 2237.2 0.060 0.064 0.065 0.064 1340.4F 1342.0F le-4 TC-3 TC-5 TC-6 TC-7 TC -8 OEQ C DE0 C DE6 F OEO F DE5 C OEG C P 3 1 e 3 008 1316.51 1405.W 1347r9C 1 bO5*Ef 1344 r5F 1329.1C 2 m . l 004 13b6 *OF 1406.5F 1347.5t 1405.6F 1347 OF 1330.8C 006 2232.8 1366.9F 1405r6F 1350.8F 1406.5F 1 347 SC 1330 OF 008 2P33.5 1367 *8F 1604 r7F 1353 * O f 1404.7F 1344rlC 1329. IC 2134 r 2 1366eOF 1405 r6F 1345r8F 1328.W 010 1351.7F 1402.1F 2 P 3 4 . 9 1406 *OF 012 1369.W 1350r8F 140 7 3F 1345.8C 1326.0F 1344.51 014 2 t 3 5 *6 1365 o2C 1405.2C 1348*7F 1402 6 : 1329. I F 016 tP36 3 136609C 14r?6*5F 1350*0F 1405*2F 1345*4F 1328dF 2P36.8 1367.8F 1405.2F 13$1*IC 14064F 1341.bP 1325*6F zm7.e 1367 8C 1407.81 1350 81 14074C 134606F 1332.61 0 % ) LOAD TIHE CH 59 Cu 60 CW 61 Cw b2 Cw 63 CW 1 4 YC-10 TC-13 W - l I DE0 C oco 1 DEQ 13Y3.W 1379.51 ooi? 1603.W 1377r81 004 1 3Y* 41 l404*71 006 13Y4.71 1403.H 1378e61 008 1394 r 3 1 1403.01 1 376rYF 010 13Y3. b? 1376r51 OIOI*71 1394r3F 1 400 8F 1377 * 11 01s 13Y5mt1 1376.01 014 1401*71 016 1391rlF 1403.W 1378*11 018 13% *OF 1401.3t 137% 11 B Z ? 13YZ011 1 O l O . O ~ lbO@*W TCog5 I C 0 1 1 tC.18 T C m l Y TC-LO DE0 C DE0 F o m F DL0 1 DE0 F ooe rrw.3 1326 801 13PZ e 3 1 1 l m 4 P 1251r7t 1204*11 004 e232rl I328 *2F 1320rlF 1124r51 123OrOF 12OfrOt 006 2232.8 1326 v 9P 11Z3r7F 1 23O.OF 1320r01 1206 * 61 12280 3t 001 l233rS 1329 *SF 1328.01 1206r6P 1123. 7F 010 2234.2 1328 2F 1316 e 9 1 1207 OF 1124.51 122Cr61 012 2P34r9 1324 9 78 131) e l ? ' 120696: 1126.2f 1 Z27r d l 1326.9F 1319.51 1123*3F 12i?lr7?

014 223906 iZO4* 1F 016 2 2 3 6 . 3 11 25r8F 1 2tl.OC 1326.01 9 32014l 1209*11 018 2 m . a 1326 61 t3EO.41 120Y.lt 1127r51 12t8r71 O? Y EP37 2 132s r ZF 1321 r7C 1205r O F llE8.8P 1 2 3 O r O ~ 27-71 --I- -

L

.- VI m :: .- N Q 00 . a - ?

m -4 8 '

s

27-72 WLE 27-28 ! C H I C - S ERmW- FCR BEADED COMPRESSION ? A N J 3 L ( 2OOM TEMFElUTW) I !

N . +- + - * --- - 0 o N = E

t"------o-- I -1 9 9

-----

- E L $ 9 - 9 . . + . .

N m ~ N N = i ? 9 b - - + ! ! . - 27-74 I W I X 27-30 TEMPERATURE DISTRIBUTIONS FOR TUBULAR COMPRESSION PANEL Ct4 49 LbAO T I H E C U 47 Cu 48 Cu 50 CU 51 t u 5 2 LVOT. LV01. LVOT .

C V D I .

P I . ? PI-1 PI-3 Pl.4 1 C . l TC.2 I hcuts I LICUES DEC F OEG F IhCUES I N C ~ 3 1380.8f OO? 1447.5 OmOOOJ O.OooJ 0.oOoJ 3.oooJ 1406.5F 1383.3F 1615.6f 094 1448.1 0 - O O R 0.008 0 . 0 0 7 0.006 0.013 1393 OF 1413.9f 000 144s.9 0.013 0.014 OaCl?

0 . 0 2 1 0.020 0.019 1 3 9 3 e O f 1419.w 001 1449.5 0.019 1450.1 0.024 0.027 0.026 0 . 6 2 5 1375.2f 1412.bF 0.032 0.030 1407.31 1 * P O .OF O I P 1450.8 0.029 0.032 013 1451.3 0.031 C.035 0.033 0.03F 1409.11 1421.3F 1 b G 7 0 3 F 01. 145?.0 C*03* 3.037 0 . 0 3 6 0.036 1423.9F 1493.7 0.036 0.04c 0 . 0 3 9 1400.0f 1426.01 015 0.039 0.042 1402.6~ 1418 *6F 016 1453.3 0.039 7.043 0 . 0 4 2 1453.8 0.041 3.046 0.046 O . O v 5 13611 PF 1421.3F 1454.4 O e C 4 4 0.048 0.047 0.047 1418.2F 1428.2C 1455.0 0.045 0.050 0 . 0 4 9 0.048 1417.3F 1418.6F 1499.7 0.047 0.053 0 . 0 5 1 01050 1416a9F lb31.7F 1456.2 0.050 0.055 0.054 0.053 14 1708f 14P304f 141 7 . RF 1 4 4 i . 3 F 02? 145608 0.052 0.057 0.056 0.055 0 . 0 5 9 1420.0f 1418.2f 023 1457.3 0.054 0 . 0 5 9 0.058 O.OS6 0 . 0 6 1 0 . 0 6 0 1422.1~ 1433.01 024 1457.9 0 . 0 6 1 1 4 5 8 . 8 0 . 0 6 0 0.067 0.066 0.065 1416 * S f 1420.8F 1459.3 0.063 0.069 0.069 0.067 1424 3F 143806f 028 i459.7 0.065 0 . 0 7 1 0 . 0 7 0 0.069 14ZP.lF 1437.8f 029 1S00.2 0 . 0 6 6 0.073 0 . 0 7 . 0.071 1423.4F 1419.6F 0.075 030 l S O O . 8 0.069 0.016 0.073 1410.0F 1432elf 0.078 1437.3f 031 1501.3 OeO?;! 0 . 0 7 9 0.016 1424.W 012 l t 0 1 . 7 0 . 0 1 4 0 . 0 8 1 0.080 0.079 141 8 . 2 f 1426 e 5 f 0.083 14P3.0F 1436.5F 033 1502.1 0 . 0 7 6 0 . 0 8 5 0 . 0 8 1 034 1502.5 0 . 0 1 8 9 . 0 8 6 0.016 0 . 0 8 4 1 4 10- OF 1408 4 1 035 1503.0 0 . 0 8 1 0.090 0.087 C.086 1425 *6F 1433.9f 036 150305 00083 0 . 0 9 1 3.089 1416.9F 1431. 7 F 0.092 037 1 5 0 3 . 9 0.086 0 . 0 9 5 0 . 0 9 4 0.092 l+iS-If 1428obF lb30.8~ 038 1504.2 0.088 c .097 0 . 0 9 6 0.094 1421.3f 0.099 1422.11 1434.3p 039 1 5 0 4 . 7 0 . 0 9 1 0.099 0.096 0.104 1424 3f 1421.31 041 1 5 0 5 . 4 0.095 0.105 0.102 1421.31 1434.71 042 1501.9 0 . 0 9 8 0.108 0.108 0.104 LOA0 T W C CH 53 CM 54 cw 55 Cw S b t u 17 EM 18 lC-3 TC-4 T C - I 1e.r l C . 7 l C . 8 OLO f D t B 1 DE0 f DL5 1 oca r ota I OOP 1432061 1 4 ? O . O F 143107f 1253 *If 1372.11 004 1450-4F 1 4 6 b r ) f 141703F 1252.51 1376.91 006 lbhl e7t 1473hf 1442elF 1262 61 1304 -51 008 1443.41 1474e7C 1435.2f 1254.11 1381.81 010 1476.bf 166OebC 1 4 4 3 . 0 f 1264e7f 1341.61 012 14450 b? 14b4.3f 1446eOf 12% 11 1383.71 1 4 6 5 . 2 1 1474*7l 013 1447.8F 1266*9? 1390.41 014 1470.8f 1470e4f 1 4 4 3 . 9 f 1242.lC 1 367.W 01% 1444.71 1475s6C 1 4 5 0 . 0 f 126% 11 130Oe01 016 146407t 147304f 127P.W 1385.01 1 3 9 8 . 2 f 1443-4f 017 1465 8 bf 147609F \274*11 1393-4C 1472 r 6 t 1275.01 1393aOF 018 1471e7F 144206F 019 1461.31 1265.6f 1479.5F 1b4407F 1393.41 020 1468-6f 1476.9F 144609F 1276.2f 1390dF 021 1466.01 1472e6F 1?32*1F 1273*7f 1339blf O ? t 1463.91 1479*SF 1 4 4 8 . 6 F 1280 001 1388r7f 023 1462.11 1468.2F 1443eOf 1277 9 f 1391.3t or4 146905f 148300f 1446.9f 1278*3f 1395r2F OLb 1470.8f 1470.1~ 1283. Y?

1442.1F 1 3 8 7 . 5 F 1470e~t 1466e6F 1379elt or7 1435 *6f 1280.

1476*sf 1477.8F 1443.bf 1281.11 139103F 029 1 4 b 9 . 1 f 147708F 1449*1f I r78.3r 139baOC 030 r471.3f 148b.Sf 1444e3f 127a.71 1894. 31 o n 1471 e31 lb01e7f 1437.3f 1282.11 13Y6.01 0 3 1 1471.7F 1267 31 1383. 31 1466r5f 1437e8f 033 1468.21 1478.61 14*0.0f 1286 01 1381e6l 034 1 470*41 1466 10f 1442.6f ll84r71 1381.51 14> *?f 035 1455 *Or 1480.4f 127brrt 1393h41 036 i466.a~ 1466.W 1446r91 lL64-71 1374 e 71 1 b 6 0 e W 147be3F 14*0*8f 1279.V 1386.4C 038 145@*bf 1472 abf 1439mlr 1281.71 1389eW Of) 14bS e 0 1 1479ebf 1440*4F lt80.4t 1391031 041 1483e4C 1450 e)? 1b43.01 lL76ebt 1390eOl Obl? lbbbe7t lb7b.Of 1 (I79 e a ?

1bSb.W 1383r3?

27-75 L O A 0 T l Y t CU 59 CU 60 CH 61 CH 62 CU 63 CH 64 1C-9 TC.10 OEG F OEG F 1408.61 ooe lbb7.5 1368.6f 1389.1F 1431.3F 14hR.l 136i.3F 1145.4F 1419.1F 004 1435.6F 144r.9 1361 -3F 1387eOC 1425.PF 1 4 13rOF 1449.5 1315.2F 1389 e5F iu*9.5r 1413.91 010 1450.1 1340sOF 1119.5F 1439.1F 1402-bF 1430.8 1371.7F 1367 3F 1442.15 14930OC 139905f 017 1451 - 3 1356.9C 1393.4F 1450.0F 1452.0 1362.1F 1-22.1F 1408.6F 014 1390 * 8 F 1492.7 1370.W 1376 .OF 1451 -3F 1423.01 1451.3 1350.0F 1391.7F 1432.6F 1*09*11 1413oOF 017 1453.8 1373.4F 1397.3F 1451.8F 1414.71 018 1454.4 1378.61 1396.9F 1442.BF l450.4F 1418e6F 019 1455.0 1379.5F 1390.8F 144882F 1420.8F oeo 1.55.7 1382.OF 14OO.OF l45B.2 1378.2F 1453.61 1413.4F 021 1395.6F 1456.8 1376.SF 1395.2F 1 447.3F 142300C 1457.3 3372.1F 1384.5F l454.5F 1429.21 024 1457.9 1375.6F 139P.lF 1 4 4 2 . 1 ~ 1416e9F 026 14%.8 1382.5F 1400 r4F 1455.9F 1420.41 1417.3F 1459.3 .384*5F 1368 -2F 1462.1F oz7 1459.7 135802F 1402.1F 1453elF I42O.W 1500.2 1370.8F 1398.6F 1651 -3F 1blbrOF O W I408.2F 030 1500.8 1389.SC 1400.8F 145500F 1393 e 4F I413.W 031 1501.3 1378.6F 145000F 1501.7 1350.W i 3 a 5 . 4 ~ 14 13.9t 032 1410*0F 1502el 1381.2F 1396.0C 1443.9F 1400 e 4F 0 33 150205 1348.3F 1396.9F 1h49.5F Ibl3.4F 1503.0 1379.1F 1400.4F 1432.1F 1413.4F 036 1503r5 1370.RF 1391.3F X442.6F 143004F 037 1501.9 135609F 139201F 1442.1F l 4 l l . 7 1 03R 1504.2 1374.3F 140806C 1377.8F lO*Q*OF 1504.7 1364.7F IJTI.-C 1036.2F 1415.21 1505.4 1376.9r 1393.9F lbb4o3F 1396.9F Ob? 1505.9 1360-4F 1385.0F 1442.16 l4P0.91 Lnm CU 66 CH 67 C H 68 CU 69 CU '10 TC-15 TC-16 DEI3 F DE0 F 1407.5 1427eIF 002 1378.61 1448.1 139609F 1397.31 cob 006 1448.9 1413ebF 1315.41 008 1449.5 14POeUF 1386.2r 14110.1 13We4F 010 139502F 1450.8 1405v6F 012 138P.O: lb51.3 1415w6F 013 1405.21 1452.0 13Y700F e 6 1 014 1398 015 1452.7 1426rOF 1405 2 1 016 1453.9 142OoOF 1381 061 1453.8 14JOr8F 017 1401 *3?

lb54.4 t4E4.7F 011 136008F 1455.0 1424.9F 019 1401 e7C 1445.7 1428r6F 1396.5F 021 1446.2 1429elF I404e7F 022 1456.8 1435r2F 1402 *6F 0?3 1437.3 1431e3F 1407 e8F 1457.9 1431.3F 024 1404r7C 026 145a.a 1 4 2 7 . 8 ~ 140309F 027 1459.3 1400wW 1399.IC 028 1459.7 1b28.6F 1405.2F 029 150002 1434r3F 1403e9F 030 150008 1431e7F 1393-91 031 1501.3 1428e6F 1406.9: 032 1501.7 1415.21 1395eL: 0 33 1502.1 I42685F 1404*36 034 1402.5 1418.21 a395 *ZF 035 150300 1433eOC 1382.W 036 1503.5 lbl4m3F 1386a6t 037 1503.9 1426.01 l3bB.56 038 150402 148boOF 1378.PF 039 1504.7 1519.511 t 396 r9P 041 1505.4 1480.4? 1390eOP Obi? 150509 1392.11 1397r3C 27-76 2 I

R

Q

t '"

27-77 27-78 - *

/ o E

I I 27-79 TABLE 27-34 S U M M A R Y CORRELATION OF STRUCTURAL Ec;EMENT TESTS rubular Closeout RT 105 300 L Crlppllng RT 105 300 L Crlppling 1400 78 500 L Panel(c) RT 105 300 L

j ~ a n e ~ ‘ ~ ) 1 I 78 500 L

-_- - -- 3eaded closcuut 130 000 L Crippling 130 000 L 1 Crippling 1400 92 500 L P’nnrl

I 27

:orrugetlon- tlffened

Crlppling I 1400 1 16 200 I., 4 1 500 Cl 30 000

43 700 Unknown 39 600 P a n e P ) RT 22 ZOO L 92 600 P 24 700 Eccentrlc end loading and n panel bowing Imperfectlon of 0.10 meeeured at midpanol Penel@) 1400 16 200 L 17 300 32 000 Some portbuckllng behevlor 24 00OP rrrperoldnl 92 400 None Crlpyllng RT 69 600 L 86 600C 69 600 orrugation 66 800 None 75 600 None 4 9 800 Panel lnatablllty with poWble lnlcractlon with lnltlal bucklln(l !Iroulw ah, 40 noo~ L‘orrugullun None Shear penel 38 4ouj _I_-.

per cap 127 200 Slight eccontrlc cap londlng Crippling

-

:Code for type of buckllng: L local. P panel, C crlpyllng.

Tested with clamped londed e a e a : all olhor types of psncln tealed with slmplc support-londod edges.

‘All panels tested for panel buckling were 30 In. long.

27-80 TABLE 27-35 C O M P A R I S O N O F TUBULAR AND BEADETI CONFIGURKTION INITIAL BUCKLING TEST REsUIlIs WITH PREDICTIONS Panel concept Tubular Beaded T e s t type Panel Crippling Crippling -~ Test temperature RT 1400?F EIT L400°F R" 14OO0F ~ ~~ Avg t e s t i n i t i a l

(4

Buckling s t r e s s (psi) 88 ooo 66 700 3 0 200 96 700 65 400

73 800 Calculated i n i t i a l Buckling stresses ( p s i )

-

12-14, a r c -buckling 105 300 .05 300 130 ooo 92 500

18 500 78 500 (local) Test/Pred.

0.84 o .85 0.70 1.02 0.75 0.71

- -

Interrivet buckling(a) 82 500 60 500 82 500 50 500

- -

Test/Pred. 1.07 1.10 0 .go 1.33 Buckling of flat (b) 76 600 56 ooo 76 ooo 5 6 000 97 500 71 200 Text /Pr ed.

1.15 1.19 1.43 0.99 0.92 0 097

-

- - - -

Comment s fietachec spots aBased on one loose spotweld i n each row of double row, located side-by-side; S = 0.5 in., K = 3.5.

bBased on t r e a t i n g one sheet i n t h e P l a t as a place with no spotwelds.

CUnknown amount of bending was applied 27-81

. E

-

i

I

.R

I

i a !

I

ti

I I I I I / I

i

I I I !

I

I I i \

I

i I I ‘i

i I

I I I O h I i i

‘ I

i I

I

I

I !

J-‘

27-82 Inconel bearing plate and pytoform blocks used fcr elevated temperature test setup. Nichrorne heating elements are inserted into precast holes, Typical elevate4 temperature test s e t UP FJ-gme 27-35.

f o r 30-inc3 compression panei ? { - 1 7 3 0 0.004 008 0. 012 0.016 0.020 0.024 Strain, in./in.

Figure 27-38, Tensile stress-strain cuxves for .016 gwe Reng 41 compression panel sheet material, longitudinal grain direction 27- 106 , . r-{ Q ' .. .

: ! L >, .'.'

N 27- 107 S t r a i n gage locations i o ~ cwrugation s t i 3 f e n t d Figure 27-41.

skin end-closeout panel rrfI. 108 .

27-11.0

t

T

27-111

r

I - -

E

I

27-112 27- 113 -P

*

-

C -z 20 m

[

U 0 0.002 0.004 0 . 0 0 6 0.008 0.01 0 0.012 Panel shortening, A L / l , in./in.

- . . . - __

- - - - .. - _ _ - .

. .

Figure 27-43. Panel shorteiing curve AL/L for corrugation-stiffened end closeout panel, room temperature Figure 27-45. Strain gage locations for beaded end-closeout panel 27-116 F Y X I I 27- 120 n

-

-

C .o 20 Y) Y)

E

U 5 ' 0.001 0 . OM 0 . 0 0 3 Panel shortening, AL/L, in./in.

Figure 27-47. Panel shortening curve AL/L for beaded end-closeout panel, room temperature 27-125

r9

FS sure 27-49. Strain gage locations f o r tubular end-closeout panel c .- C .- a “ 0 X 27- 126 a c' .- 0,

.-

-a xi

-

.- 6 40 In In e

E

V 0 0.001 0.002 0 . 0 0 3 Panel shortening, AyL, in./in.

- Panel ehortening curve ,AL/L for tubular end-closeout panel, Figure 27-51, room temperature 27-128 *

19.0 -----

/”

0 Total no. of gages = 14.

0 Gages 11, 12, and 15 located directly below gages 7, 8, and 13.

Figure 27-53. Strain gag@ locatfons for corrugation-stiffened crippling panel

Q

n c .- , C .- i c' .- e n Q - .- e

Q

N C .- a c' .- e c m

-

t

n

0 1 8

0 N 8 h 0 27-132 27-133 1 ) - d -

.- 40

v) u l Q

Eo

v 0.001 0.002 0.003 Panel shortening, AL/L, in./in.

Figure 27-55. Panel shortening curve AL/L f o r corrugation-stiffened crippling panel, room temperature 27-135 19.0

--I

Figure 27-57. Thermocouple locations f o r the corrugation-stiffened crippling panel

-

.- 5 20

g

f

V 0.001 0 . 0 0 2 0 . 0 0 3 0.004 0.005 0 . 0 0 6 Panel shortening, AL/L, in./in.

Figure 27-58.

Panel shortening curve f o r corrugation-stiffened dL/L crippling panel, 1400° F , I Figure 27-60. S t r a i n gage locations f o r trapezoidal corrugation crippling panel )QOD .-

m 0 1 1

- C

. : 20

v) ? !

a .

u .0005 ,001 ,0015 .002 ,0025 ,003 Panel shortening, A L/L, in./in.

Figure 27-62 Panel shortening curve 'L for trapezoidal corrugatlon crippling panel, room temperature 27-145 Figure 27-64 Thermocouple Locations for the trapezoidal corrugation crtgpling panel 4 0 x IO I 1s 0,001 0 , OM 0 . 0 0 3 0 . 0 0 4 0 . 0 0 5 0.006 Panel shortening, AL/L, in./in.

. . . . . ..

Figure 27-65 Panel shortening curve AL/I. for trapezoidal corrugation crippling panel 27-248 1 1 . 1 2

f

Note: iyguxe 27-67 Strain gage loca$ions for beaded crippling paylel

I I I

t d

*

-

s 20

! L

V 0 0,001 o b 0,003 0 . 0 0 4 0,005 o b 906 Panel shortening, hL/L, in,/inb

--

. - Figure 27-69 Panel shortening curve AL/L for beatled cr5pp5ing pane3, room temper&%we c Figure 27-71 Thermocouple locations f o r the beaded crippling panel 27-156

-

.- 0 ' 20

m H

g

v

0.001 0. w 0.003 0.004 0.005 0.006

Pnnal shortening, Al/L, in&.

Figure 27-72 Panel shortening curve AL/L f o r beaded crippling panel, 1400°F Note : 0 Total no. of S.G. =: 16.

0 Gages 7 and 8 located 1/2 distance from tube to flange.

0 Gages 13, 14, 15, 16 located directly be F ow 9, 10, 7 and 8.

_- Figure 27-74 S t r a i n gage locations f o r tubular crippling, panel I .- - -.

27-159

I

T

-0 -: 27- 160 x m ! 3 I

Y

y

q

5: h

Q c

I VI h x I 27- 162

s

a’ B

-

.-

u l ul

g

V 0.001 0.002 0,003 0.004 0.005 0 . 0 0 6 Panel shortening, AL/L, in,/in.

Figure 27-76 Panel shortening curve AL/L f o r tubular crippling panel, room temperature 27-1.63 \ 9 1 1 2 J Figure 27-78 Thermocoupling l o c a t i o n s f o r the t u b u l a r : c r i p p l i n g panel

4 0 x 10 -

f !

-

g 20

.-

H m E

t

W 0.001 0 . 0 0 2 0 . 0 0 3 0.004 0005 0.006 Panel shortening, AL/L, in,/in.

-7 Figvre 27-79 Panel shortening curve AL/L f o r tubular c r i p p l i n g panel, !.4000F 27-166 15.5

- 5.5 -

I

c

1 !

Figure 27-81 S t r a i n gage l o c a t i o n s for t h e spar cap cri@ing speci-rer-s 27-168 h o_

-

5 40

1 - Q) u I 0.001 0 . 0 0 2 0 . 0 0 3 0 . 0 0 4 0.005 0.006 Panel shortening, d L/L, in./in.

- .- ____ - - Figure 27-83 Spar cap shortening curveAL/L f o r 3/8 inch flange speaimen, room temperature 27- 170

x I

I

I I

I /

I

p '

1 1 ' -

I '

I '

i 1

FigvrE 2,-S5 S t r a i n gage locations f o r corrugation -

stiffened skim compression panel @-I

I 0

!

1-

T

i

T-

I I I

t

C C

P

m P

d

R 27- 180 27-181 Figure 27-99 Thermocouple l o c a t i o n s f o r t h e corrugation- s t i f f ened- skin compression panel 27- 1.83

I

40 I O -

-

.- 0 ‘ 20

H H a l L

! $

U 0.001 0 . OM 0 . 0 0 3 0 . 0 0 4 0 . 0 0 5 0.006 Panel shortening, AL/L, in./in.

-__.- - Figure 27-91 Panel shortening curve AL/L f o r corrugation- stiffened-skin compression panel, l4OO0F t e s t . ..__ I 27-184

-

5 (6 3.4 30.0 i Figure 27-93 Strain gage l o c a t i o n s f o r trapezoidal corrugation compression panel 27-186

I

I

oaoD

. ._I

2 12 v, 0 0 3 3 w ON

r-r

> I 27- 188 27-193 27-195 !n m 0 m

IO-/ ' ( 7 J b o

Thermocouple l o c a t i o n s f o r t h e t r a p e z o i d a l Figure 27-98 corrugation com,pression panel 27- 198 ,001 ,002

Strain, ALA, i n h

Figure 27-99 Panel shortening curve A L/L f o r t r a p e z o i d a l corrugation compression panel, 140Q°F 27-203 n

t -

27-204 a . - 27-206

t

cu r: F- cu a , k 27-208 - O d U D I_-

i

I - -

t

s : 0 n l?

-

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0.0 m * G 7 3 % I I - 2 27-210 . . -

3-7 - 211

Y I6 x IO

a

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a a U Panel expansion, in.

.

_I_._ - - Figure 27-19s Expansion of beaded compression panel due t o a x i a l compression loads, room temperature-- - - - - - __ - -- . -- 27- 212

I----- Lt.5

Figure 27-107 S t r a i n gage l o c a t i o n s f o r tubular compression panel 27-214 27- 21.6 27- 217 27- 218 27-219 O Q O D o x

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P ' 0 I ?

PI

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27-220 X 27- 221

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27- 222 X P

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k

V 0~0005 0,001 0.001 5 0.002 0,0025 0 . 0 0 3 0 Panel shortening, AL/L, in./in.

Figure 27-109 Panel shortening curve AL/L f o r tubular compression

panel, room temperature 27-223 30.0 i

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-i I

. - - - ..

Figure 21-111 Thermocoup!.e location for the tubular compressior.

panel, 2b0°F t e s t 27-225

s

-

.- 40

W B

{

0. o001 0. oooz 0.003 0. oO04 0. oO06

Panel shortening, AL/L, in./in.

- _ _ ~ . . -~ 'Ygure 2'7-112 Panel shortening curve AL/L for tubular compression panel, 14000 t e s t rl-226

T I I

7.75

. - . L

Figure 27-114 Strain gage locations for the circular arc corrugation shear panel 27- 228 ~ ~~ 400 800 I200 1600 ism 2403 2 m 3200 Principal rhoin, p in&.

Figure 27-115 Relationship of principal s t r a i n s and applied v e r t i c a l cantilever loading for circular arc corrugation shear panel (TIG w e l d w i t h Rene' 41 filler wire), r o w temperature 0 500 1000 1500 2000 2500 3MH) 350 #ooO Moximum shear strain, p in./in.

Figure 27-116 Relationship of shear s t r a i n and applied v e r t i c a l cantilever loading for circular arc corrugation shear panel (TIG weld with Rem' 41 filler wire), room temperature IOM) Principal &train, fi in./in.

F'igure 27-118 Relationship of principal strains and applied v e r t i c a l cantilever loading for circular arc corrugation shear room temperatme panel (TIG weld with Bastelloy W f i l l e r wire) 0 500 loo0 1500 2My) 2500 b i b x i r u m shear stmin, p in./in.

Figure 27-119 Relationship of shear strain and epplied v e r t i c a l cantilever loading for circular arc corrugation shear panel (TIG weld w i t h Hastelloyh "iller wire), room temperatwe 27-231 40A

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u

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ti L

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c - c -

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ld 0 1.0 2.0 3.0 4.0 b/o Figure 27-121 Degree of conservatism in the wide-column analysis as applied to compression panels VS. width-to-length r a t i o 27-233 r l J- v) a , c, M) I= .rl r-i ."

...

.-!

u L o I n 3 d 9 9 1 .o .9 .8 .7 .6 g " s 2 .5 .- U .- - t n.

.4 .3 . 2 * 1 90 100 110 120 130 140 150 160 lo3 ,.rers, Ib/in2 Figure 27-125 P l a s t i c i t y factors f o r 0.060-in. Rene' 41 sheet at room temperature