Advanced composite vertical stabilizer for DC-10 transport aircraft

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

SECTION l

INTRODUCTION

The overall objective of this program is to accelerate the use of advanced

c_osite structures by developing technology and processes for early

progressive introduction of composite structures into production commercial

transport aircraft. Key steps in accomplishing this objective are:

(1) to develop low-cost design and manufacturing approaches which will

produce a cost competitive structure, and (2) to initiate commercial

airline service of a mid-sized composite primary structure, the DC-IO

con_oosite vertical stabilizer (CVS).

The Work Breakdown Structure (WBS) for the program is presently organized

in eight major tasks as follows:

I. Preliminary Design

2. Detail Design

3. Manufacturing Process Development

4. Con_osite Structure Fabrication

5. Subassembly, Subsystem, and Other Fabrication

6. Assembly

7. Verification Testing

8. Contract Management and Plans Development

In Task l, the Preliminary Design Synthesis culminated in selection of a

four-spar, multi-rib structural configuration similar in geometry to the

existing baseline metal stabilizer for structural interchangeability. The

composite skin panels between spars and ribs will be stiffened by honeycomb

sandwich construction for minimum weight and cost purposes, and the skin

panels wi II be mechanically attached to the spar-rib substructure to enhance

detai I part fabri cati on, assembly, inspection, and maintenance access.

The selected design concepts are presently being verified through testing of

the concept and joint development component.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

In Tasks 2 and 3, currently active, the structural detailed arrangements,

tooling, and manufacturing processes are being developed through engineering

and tooling design for the full-size CVSand through experimental development

and testing of structural components representative of critical design

features. The final development componentwill be a box-beam approximately

eight feet long representing the lower portion of the CVS. The ribs, spar

segments, and skin panels of this component will be fabricated and assembled

using the full-scale tools for the CVSand will serve as a tooling and

processing verification componentas well as a structural test component.

Following completion of the box-beam tests, eight stabilizer units will be

constructed in a serial production mode in Tasks 4 through 6. The first two

of these units will be used in Task 7, Verification Tests, for the ground-

based static and repeated load tests. The third unit will be flight-

tested as part of the Federal Aviation Agency (FAA) certification testing

and, together with the remaining five units, will be introduced into

commercial airline service after receipt of FAAcertification. Task 8

includes program management functions and the formulation of the plans

necessary for development, certification by the FAA, and in-service

inspection and maintenance of the CVS.

This report describes work accomplished during the seventh quarterly period

of the program. Work continued on structural design and analysis; weight

and cost analysis; and tooling, fabrication, and test of the structural

development componentsand specimens. Detail design of the composite

structure was continued and tool design was initiated. Overall schedule

status is summarizedin Figure I.

The activities during the quarterly period are described under the headings

Detail Design, Concept Development Components, Joint Development Components, Mechanical Property Testing, Design Verification Test Components, Tool Design,

Cost Analysis, and Quality Assurance. Engineering drawings of the composite

skin panels and selected spar and rib assemblies are included in the appendices.

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Douglas Aircraft Company ACEE-O3-PR-9642

Contract NASl-14869

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ACEE_)3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

The measurement values in this report are expressed in the International

System of Units (SI) and also U.S. Customary Units in some cases. U.S.

Customary Units were used for the principal measurements and calculations.

D .J

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NAS-1-4869

SECTION 2

DETAILDESIGN

Detail design of the CVScontinued with engineering drawing preparation,

concurrent structural analysis, and updating of the weight status report

based on released engineering drawings.

DESIGN DEVELOPMENT

A CVSdrawing list consisting of 89 detail, assembly, and installation

drawings was completed and drawing numbers were assigned. Twenty-six of

the required drawings are identical to or require only minor modifications

to existing DC-IO drawings. Twenty-seven of the required drawings are new

drawings of conventional metal details. The remaining 36 are new drawings

defining the advanced composite structural elements.

The spar and rib locations of the exi sting metal stabi li zer were retained

in the CVSfor interchangeability with the fixed-fin structure and the rudder

system. The locations and drawing numbers of the major composite structural

elements are shown in Figure 2. With the exception of the two uppermost

ribs which were identified in the vertical stabilizer coordinate system

(subscript V), the rib stations were identified in the forward rudder

coordinate system (subscript FR). Engineering drawings of the skin panel

assemblies, the substructure assembly, and typical rib and spar

installations are included in Appendix A. Design features of the major

composite structural elements are described in the balance of this section.

Skin Panels Left and right-hand skin panels will each be made as a single molded honeycomb sandwich assembly. Spar and rib cap laminates will be included within the sandwich as shown in Figure 3. To maintain structural continuity at spar and rib cap intersections, the caps will be laid up with graphite unidirectional tape in a pseudo-isotropic pattern. The spaces between the caps will be filled by Nomex 4.0 pcf honeycomb core, 0.30 thick with a ACEE-O3-PR-9642 ORIGIHA_ P._

Douglas Aircraft Company

OF POOR QUALITY

Contract NASI-14869

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ACEE-O3-PR-g642 Douglas Aircraft Company Contract NASI-14869 I/8 inch cell size. Syntactic foam will be used around the periphery of the core to stabilize the edges and to provide a transition region between the solid cap laminates and the thin sandwich facings. The gaps between the core edges and the surrounding laminates will be filled with foaming adhesive to ensure good shear connections at the interfaces.

The sandwich facing material will consist of graphite bi-weave cloth layers in a ±45 ° orientation with respect to the box axis. Near the root-end, some additional 0/90 ° layers will be added to help align the CVS box axis with the lower vertical stabilizer box axis. FM 300 K adhesive will be placed between the facings and the core, and the whole assembly will be cured in a single autoclave operation. Recesses for leading and trailing-edges and for access panels will be molded net, and the panel edges will be machined as the final fabri cati on operati on.

Spar Assemblies The four spar assemblies will have several design features in common. Each spar will have a pair of titanium fittings bonded within the graphite/epoxy at the root-end. The entire load of the stabilizer will be transferred into the lower vertical stabilizer through these fittings and their associated attach bolts. A detail view of the fitting installation is shown in Figure 4, and a more general view of the root-end of a typical spar in Figure 5.

Each titanium fitting will be basically a "tee" section where each of the three legs will be tapered to a narrow edge at the upper end, interfacing with the composite spar cap in a scarf joint. This joint will be co, cured and adhesively bonded during the autoclave cycle in which the entire spar assembly is cured. A secondary load path, capable of transferring design limit load, will be provided by means of mechanical fasteners. The cross- sectional area of the spar caps will be reduced rapidly away from the root- end as the load is transferrec_ into the skin caps.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

OR!G|NAL OF POOR BASE SPAR ASSY SKIN PANEL TITANIUM FITTING ADHESIVE SCARF JOINT REGION FIGURE 4. SPAR-ROOT FITTING SPLICE_JOINT DETAILS ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 oRIGINAL PAGE IS OF POOR QUALITY >- .J .m (( mr (( eL .J (3 >.

F- w m- CO m LL Douglas Aircraft Company ACEE-O3-PR-9642 Contract NASI-14869 Localized honeycomb sandwich stiffening will be used near the root-end to stabilize the shear web against buckling. In most other regions of the substructure, sine-wave webs will be used to provide stabilization as shown in Figure 5.

The sine-wave geometry is more properly described as intersecting circular arcs. A single wave geometry was selected for use throughout the entire substructure to reduce tool machining costs. This type of stiffening has been found to offer many advantages in cost and weight studies. Some of the weight advantage will be lost in reinforcing access holes and in mounting brackets and clips for rudder attachments and systems installations.

Plain holes in sine-wave webs were found to be inefficient because of the out-of-plane forces around the periphery of the hole. During the test program, a reinforced hole configuration was successfully developed in which a 0.50 inch wide flange was incorporated at the edge of the hole.

The required doubler material around the holes was determined from test and analysis. The required local web thickness enabled the panel to remain flat in the region of the hole as shown in Figure 6.

In the lower segment of the front spar, the number of reinforced holes became so extensive that plain thick laminates with unreinforced holes were used with some cost saving and little weight penalty. Each of the three longer spars was divided into upper and lower segments to facilitate fabrication. The lower spar portion is shorter in each case but contains the thicker laminates and the titanium fittings. The upper portion will be primarily of sine-wave construction.

At the mounting stations for the rudder hydraulic actuators, cut-outs will be provided in the rear and aft center spar webs as shown in Figure 5.

These cutouts will be plain holes in the thick laminate material.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

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Provisions will be made for access into each bay of the substructure by means of 4.5 inch diameter access holes. This internal access will assist in initial fabrication and inspection and in subsequent inspection and repair.

Base Ri b The base rib will consist of two thick laminate segments joined at the rib center-line as shown in Figure 7. Since the adjacent rib in the lower vertical is capable of transmitting the re-distributing shear loads, the web of the base rib will be largely removed by three oval holes. The remaining material will provide load paths between the root attach bolts, and the skin and spar panels. The base rib edges will be flanged for attachment of the titanium spar fittings and the skin access panels. Projections at the aft end of the rib will support the hinged trailing-edge panels.

Ri b 295 This rib is typical of the three ribs which do not incorporate rudder support fittings. The simple sine-wave webs without holes, Figure 8, are representative of many rib segments throughout the substructure. The skin edges of the webs will be reinforced and flanged to form "tee" section caps for attachment to the skin panels. Drainage holes will be provided in the valleys of the sine,waves to prevent standing water from accumulating.

Actuator Ribs Actuator ribs are located at stations ZFR = 314.000, 328,000, 430.500, and 444.500. The rib at station 314, illustrated in Figure 9, is typical of this group. The web construction will be similar to that of Rib 295 except that in this case, an access hole will be provided in the center segment. A cover-plate over this hole will prevent loose objects from falling into the bay below during hydraulic actuator maintenance.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

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ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASl-14869 The aluminum alloy actuator bracket will locate the rudder hinge point and the mounting provisions for the rudder hydraulic actuator trunnion. The actuator bracket differs only slightly from the bracket used on the metal stabilizer. Separate fittings, forward and aft of the rear spar, will transfer the rudder loads into the composite box structure. The shear webs at this fitting interface, and at the front spa_ will be reinforced with doubler layers to accommodate loads from the rudder and from the control quadrant mounting on the front spar. This particular rib will also incorporate an oval cutout for hydraulic piping.

Tie-Rod Ribs Tie-rod ribs are located at stations ZFR = 350.319, 424.219, 473.581, and in the tip region. These ribs differ from the actuator ribs in the manner in which the tie-rod brackets transfer loads into the composite box-structure.

A typical fitting arrangement is shown in Figure lO. Fail-safe requirements make dual load paths necessary and hence two fittings run along the rib caps at each skin panel. Since it is not convenient to attach a sine-wave web to the face of one of these fittings, a thin honeycomb sandwich panel will be used in this rib segment.

At the upper tie-rod station, a separate structural sub-assembly will be added to the box-structure as shown in Figure If. This complex region will provide fail-safe load paths for loads arising from the rudder balance weights.

Seven existing metal fittings will be contained within this sub-assembly, including the brackets for two hinges and two tie-rod attachments.

Plain Hinge Ribs The ribs located at station ZFR = 398.681 and 51g.081 support simple hinge brackets. Since the rudder loads are smaller at these ribs than at actuator or tie-rod ribs, the load transfer to the composite box structure will be accomplished by simple fittings mounted on the aft side of the rear spar as shown in Figure 12.

ACEE-O3-PR-9642 Douglas Aircraft Company Contract NAS1-14869

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Douglas Aircraft Company

Contract NASI-14869

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D + ..J li,I e_ D < u.I -r ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASl-14869 f , f Substructure Assembly The spar and rib assemblies will be joined to form a substructure assembly as shown in Figure 13. The attachment at the intersecting corners will be effected by graphite-epoxy angles which will be co-cured and adhesively bonded in place.

Box Assembly The skin panels will be attached to the substructure assembly with titanium bolts in accordance with normal DC-IO fastener usage policy. Nut-plates and channels will be used in all places where access does not permit the installation of nuts.

Access Doors Access doors for the root-attach and actuator regions will be simple sandwich panels as shown in Figure 14. The basic construction will follow skin panel practice. The panel will be thick enough to sustain shear, compression, and lateral pressure loads without buckling. Solid laminate material will be used at the edges where the doors bolt to the skin panels.

Trailing Edge Panels The design concept for the graphite-epoxy trailing-edge panels is shown in Figure 15. This type of composite panel has already beendeveloped and several panels are presently in regular DC-IO airline service. The CVS panels will be hinged at the forward edges to permit opening for maintenance.

The hinge-line was moved slightly aft in comparison with the metal stabilizer.

when closed, the panels will be attached to supports mounted to the rudder hinge brackets or to the box skin panels.

Systems Installation Mounting brackets and clips will be provided for installation of the hydraulic, control, electrical, and avionics systems. The avionics VOR/localizer antennas contained in the tip and door panels will be retained without change. To maintain satisfactory avionics performance, the outside surface of ORI_NAL PAGE IS ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUALITY

Contract NASI-14869

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Douglas Aircraft Company

Contract NASI-14869

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Douglas Aircraft Company

ACEE-O3-PR-g642

Contract NASl-14869

the graphite-epoxy skin panel and the inside surfaces of the antenna bay will be flame sprayed with a 5-mil aluminum coating. The exterior aluminum surface will also provide protection against the l}ghtning travelling-stroke phenomena. Direct lightning paths will be provided across the tip, and down the metal leading-edge and rudder assemblies.

Special care will be taken to ensure electrical continuity across panel joints and at access doors. This continuity will be achieved at the access door by coating the inner rather than the outer surface of the door with the aluminum spray. Electrical continuity to the leading-edge will require local removal of the paint finish. The metal will then be specially treated to avoid galvanic corrosion.

Drawin9 Releas e Status A total of 28 new drawings were released during the reporting period, and changes were made to four existing drawings to adapt them to the composite stabilizer configuration. Slightly more than 50 percent of the total required drawing effort has been completed. Released drawings are listed below.

AMC 784O Skin Panel Assembly AMC 7844 Substructure Assembly AMC 7845 Front Spar Assembly AMC 7846 Front Spar Attach Fitting AMC 7847 Forward Center Spar Assembly AMC 7848 Aft Center Spar Assembly AMC 7849 Lower Rear Spar Assembly AMC 7850 Hinge Support Fitting AMC 7851 Cant Rib Cap Fitting AMC 7852 Forward Center Spar Attach Fitting AMC 7853 Base Rib Installation AMC 7854 ZFR 295 Rib Installation AMC 7856 Hinge Bracket Assembly AMC 7857 Hinge Support Fitting

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

AMC7859

ZFR350 Rib Installation

AMC7862

Rib CapFitting

AMC7863

Rib CapFitting

AMC7865

ZFR 375 Rib Installation

AMC7869

ZFR 424 Rib Installation

AMC7871

Hinge Support Fitting

AMC7872

Rib CapFi tti ng

AMC7873

Rib Cap Fitting

AMC 7878

Rear Spar Attach Fitting

AMC7882

Aft Center Spar Attach Fitting

AMC7892

Hinge Support Fitting

AMC7893

Upper Rear Spar Assembly

S00202

Laminated Graphite-Epoxy Angle

S00203

Laminated Graphite-Epoxy Angle

The appropriate changes have been madeto the following existing hinge and

tie-rod bracket assemblies: AMC7029, 7031, 7073 and 7074.

Drawing preparation is continuing on the remaining ribs and their associated

fittings, on the access door assemblies, and on the trailing-edge installation.

Douglas Aircraft Company

ACEE-O3-PR-9642

Contract NASI-14869

STRUCTURAL ANALYSIS

The structural analysis effort has been devoted to completion of the NASTRAN

internal loads analysis, derivation of test conditions for the Z5943454

box-beam verification test component, and strength analysis in conjunction

with the engineering drawing release activities.

The NASTRAN model for the final internal loads analysis is illustrated in

Figure 16. The rudder modules are complete and operational. The lower

vertical stabilizer and aft fuselage module is complete and operational.

This latter module will be also used to establish the support flexibilities

for the Z5943454 box-beam verification test component. Physical and

material properties for the upper stabilizer module are approximately

30 percent complete.

The fatigue loading spectrum for the Z5943454box-beam component has been

completed. The identical spectrum will be used on the Z5943452 spar root

splice specimens. The spectrum is based on that used on the DC-IO aft-

fuselage full-scale fatigue test with the loads modified to reflect the

latest external and internal load distributions as developed for this program

(Reference l). The completed test spectrum load exceedance chart is shown

in Figure 17. The gust exceedance theoretical line was plotted directly

from the DC-IO aft fuselage test report, Reference 2, and the maneuver

line was obtained from References 3 and 4.

The analysis method for flanged access holes was substantiated during the

reporting period by successful testing of the Z5943434sine-wave shear-web

component (see Section 3 of this report). The analysis method was described

in the prior quarterly Progress Report, Reference 5, Appendix B. Accordingly,

standard access hole doubler arrangements were formulated for each sine-

wave laminate patterns of interest. The doublers are capable of transmitting

the maximumshear-flow in the most critical panel of each patternj (see Figure 6).

2?

ACEE-O3-PR-9642

Douglas Aircraft Company

PAGE

ORIGINAL

Contract NASI-14869

OF. POOR QUALITY

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ORIG|_IAL PAGE IS

ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUALITY

Contract NASI-14869

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Z5943454 COMPONENTS

Z9

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

Analysis of designed components is proceeding concurrently with the engineering

drawing release activity. A summary of analyses completed during the

reporting period is presented in Table I.

WEIGHT STATUS

The predicted weight of the composite stablizer was revised based on

calculated weights for engineering drawings released to date. The revised

weight comparisons are shown in Table 2. The current predicted weight saving

is 21.7 percent.

A weight change summary for the current reporting period is shown in Table 3.

A net weight increase of 0.2 kilograms (0.5 pounds), primarily in shear

webs, resulted from recalculation of weights based on the released drawing

configurations. The CVS weight distribution by material is summarized in

Table 4. A weight-time history for the composite stabilizer is shown in

Figure 18.

ACEE-03-PR-9642

Douglas Aircraft Company

Contract NASI-14869

ORIGINAL PAGE IS

OF POOR QUALITY

TABLE 1 SUMMARY OF CRITICAL MARGINS-OF-SAFETY FOR DC-10 COMPOSITE VERTICAL STABILIZER COMPONENTS AMC MINIMUM DRAWING MARGIN- NUMBER PART DESCRIPTION OF-SAFETY CRITICAL MODE 784O SKIN PANEL ASSEMBLY 0.15 SHEAR-COMPRESSION INTE RACTION 7845 FRONT SPAR ASSEMBLY 0.52 SHEAR STRESS IN ADHESIVE BOND- FORWARD CENTER SPAR ASSEMBLY 7847 LINE AT INTERFACE BETWEEN 7848 AFT CENTER SPAR ASSEMBLY TITANIUM FITTING AND COMPOSITE 7849 REAR SPAR ASSEMBLY SPAR CAP -- INCLUDES 133-PERCENT FITTING FACTOR 7850 0.92 TENSION FITTING --STA ZFR 316 HINGE 0.07 BENDING FITTING -- STA ZFR 316 CANT-RIB CAP 0.01 THERMAL BOLT LOAD BEARING 7853 BASE RIB INSTALLATION 0.07 BOLT BEARING 0.07 WEB SHEAR AT CUTOUT STA ZFR 295 RIB INSTALLATION OK STRUCTURALLY IDENTICAL TO BRACKET -- STA ZFR316 AND 325 HINGE EXISTING DColO PART, AMC 7031 7857 0.24 TENSION FITTING -STA ZFR 351 HINGE 0.02 7859 BOLT BEARING STA ZFR 350 RIB INSTALLATION 0.12 7862 BOLT BEARING FITTING -- STA ZFR 351 RIB CAP 0.02 BOLT BEARING FITTING - STA ZFR 351 RIB CAP 0.24 TENSION 7871 FITTING -- STA ZFR 423 HINGE 0.02 BOLT BEARING 7872 FITTING --STA ZFR 423 RIB CAP 0.12 BOLT BEARING 7873 FITTING -- STA ZFR 423 RIB CAP

ACEE-O3-PR-9642

Douglas Aircraft Company

oRIGINAL PAGE IS

Contract NAS1-14869

OF POOR QUALITY

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OR!_NAE PAGE F3 ACEE-03-PR-9642

Douglas Aircraft Company

Contract NASI-14869

OF POOR QUALITY

TABLE 3 WEIGHT CHANGE SUMMARY COMPOSITE VERTICAL STABILIZER WEIGHTCHANGE ITEM KILOGRAMS POUNDS SPAR CAPS RELEASE OF PRODUCTION DRAWINGS REFLECT CURRENT WEIGHTS --73 -17.1 INTERSPAR SKIN PANELS i DELETION OF ANTENNA PANEL FASTENERS (2.4 LB). -16.0 -35.3 ESTIMATED WEIGHT FOR PANEL HAS BEEN REPLACED BY COMPOSITE PANEL CORE (HONEYCOMB).

RELEASE OF SKIN PANEL DRAWING REFLECTS THE CURRENT WEIGHT. IN ADDITION, TRANSFER OF SOME SKIN PANEL TO INTERSPAR RIB WHERE APPLICABLE.

SPAR WEBS +12.3 +27.1 RELEASE OF SPAR WEB PRODUCTION DRAWINGS REFLECT CURRENT WEIGHT INTERSPAR RIBS PARTIAL RELEASE OF PRODUCTION DRAWINGS PROVIDES +11.7 +253 CURRENT DESIGN AND WEIGHTS FOR RIBS TOTAL WEIGHT CHANGE +0.2 +0.5 Q

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

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Douglas Aircraft Company

OR!G|NAL PAG_ I_

Contract NASI-14869

OF POOR QUALITY

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Douglas Aircraft Company ACEE-O3-PR-9642 Contract NASI-14869 SECTION 3 CONCEPT DEVELOPMENT COMPONENTS The concept development component testing for the program was completed during the reporting period with completion of the Z5943434 sine-wave spar- web test and the Z3943451 lightning protection system tests. Previously completed tests in this component group included stiffened compression and shear panel tests (References 5 through 8) a honeycomb stiffened spar-web test (References 5 and 8), and galvanic effects tests (Reference 5).

Test setups and results of the sine-wave spar-web and lightning protection system tests are described in this section.

SPAR-WEB COMPONENT The Z5943434 sine-wave shear web component was redesigned and remade as a result of previous test component failure (see Reference 5). The redesign provided for two flat areas in the web for incorporation of ll.4 cm (4.5 inch) diameter access openings. The component included one access opening having flanged edges as shown in Figure 19. The other flat area in the specimen web was left blank to permit subsequent test evaluation of an unflanged cutout, see Figure 20.

The testing was accomplished in two steps. The first test subjected the component to ultimate design shear loading (600 pounds per inch) in the web.

Load was applied to the component as a simply supported beam as shown in Figure 19. The component successfully sustained the loading without failure.

The maximum tensile strain in the flange of the cutout was 2495 microinches per inch at 152 percent test limit load (TLL).

The component was removed from the test fixture and a circular opening cut in the blank web (Figure 21). The component was re-installed in the test fixture and the test loading sequence repeated. The web failed at the unflanged cutout at I04 percent TLL (420 pounds per inch). Failure resulted from P_C_D_N.G PAGE BLANK _._O_T_CE_ ACEE-O3-PR-9642 ORIGINAL

Douglas Aircraft Company

Contract NASI-14869

OF POOR FIGURE 19. SINE-WAVE SHEAR WEB COMPONENT IN TEST FIXTURE ACEE-03-PR-9642 Douglas Aircraft Company Contract NASI-14869 ORIGINAL P_G_ ;_ OF POOR QUALITY m lu i ul z i o ,_, @,j I,_ N Z I,M n- I-- ,=,I I,,I,, I,M I,I.

i ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 FIGURE 21. UNFLANGED ACCESS OPENING IN SINE-WAVE SHEAR WEB COMPONENT 4O ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 excessive circumferential tensile strains at the edge of the cutout as shown in Figure 22. A tensile strain of 5925 microinches per inch was recorded at the edge of the unflanged cutout at failure. The test indicated the need for local reinforcing flanges at the shear-web cutouts to meet design shear load requirements.

LIGHTNING TEST PANEL Current transfer and lightning restrike tests were completed on the Z3943451 lightning evaluation panel. The 5-mil aluminum spray coating was adequate for all lightning current transfer and restrike tests.

The test panel configuration is illustrated in Figure 23. The metal tabs at the left hand edge of the panel represented candidate joint configurations at the metal leading edge structure interface with the graphite stabilizer box-structure. The joint was effected with flush screws at normal installation torque and at two higher torque values. The metal tabs at the right-hand edge of the panel represented candidate joint configurations at the trailing edge structure interface with the graphite stabilizer box-structure in the vicinity of a rudder hinge bracket. The simulation of the metal piano- hinge which will attach the trailing edge structure to the graphite_box structure was attached with various rivet combinations. Tabs IA, IB, 3A, 3B, 5A and 5B all had metal-to-metal contact between the tab and the metal spray coating. The remaining tabs had a faying surface seal (PRC 1431G) between the tabs and the metal spray coating.

Lightning Current Transfer Tests During a severe 200 kiloampere (KA) peak-current lightning-strike at the tip of the stabilier, the current transferred in the skins of the composite vertical stabilizer will be approximately 0.5 KA per centimeter of width.

The 9.5 centimeter wide joints of the test panel, therefore, must be capable of transferring about 5 KA. The current transfer tests were made with the panel mounted in the test fixture as shown in Figure 24.

Douglas Aircraft Company Contract NASI-14869 ACEE-O3-PR-9642 GF POOR QUALITI FIGURE22. FAILURE OF SHEAR WEB AT UNFLANGED ACCESS OPENING ACEE-O3-PR-9642 DOUGLAS AIRCRAFT COMPANY OF POOR QUALITY COIITRACT NASI -I 4869 r FIGURE 23. SKETCH OF LIGHTNING PANEL SHOWING TEST POINTS ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPANY

CONTRACT NAS1-14869

OF POOR OUALI T_i FIGURE 24. TEST SETUP - SIMULATED LIGHTNING CURRENT TRANSFER TEST ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 The initial test was a discharge of 5.9 KA peak current from tab IA to IB.

The action integral was .0489 x lO 4 ampere squared seconds. The same test was made from tab 6A to 6B. Additional tests were made between the other tab pairs with peak currents of ll.6 KA and 16.9 KA. Typical current waveforms for the discharges made during the tests are shown in Figure 25.

The discharge paths, current levels, and resistance values are tabulated in Table 5.

There was no measureable difference in the resistance values obtained before and after test or between the various joints. These tests indicate that all joint configurations tested are acceptable for transferring the necessary lightning current densities. The joints with the faying surface seal are preferred, because of the greater corrosion resistance.

Simulated Li_htnin 9 Restrike Tests Photographs of the test panel mounted in the test fixture before and after the ll6 KA lightning restrike test are shown in Figure 26. The test waveform is shown in Figure 27. The action integral of 0.55 x I06 ampere squared seconds was greater than the required test value of 0.25 x lO 6 ampere square seconds and the peak current of ll6 KA was greater than the required level of I00 KA. The discharge path was from the center of the panel to tabs IB, 2B, 3B, 4B, 5B, and 6B which were clamped together. There were no changes in the resistances of the panel caused by the high current discharge (see Table 6).

Photographs of the test panel showing the area where the metal spray was vaporized are shown in Figures 28 and 29.

Upon completion of the lightning tests, the lightning panel was examined using x-ray and ultrasonic through transmission NDI techniques. This examination revealed that the high current lightning strike to the panel caused a delamination of approximately 7.5 centimeters (3 inches) diameter between the outer facing and the honeycomb core in the area where the metal spray was vaporized. The damage was repairable and considered to be acceptable for composite vertical stabilizer from a safety standpoint.

0RI_NAL ACEE-O3-PR-9642 DOUGLAS AIRCRAFT COMPANY OF POOR CONTRACT NAS1-14869 VERTICAL SCALE: 2.11KAPER DIVISION HORIZONTAL SCALE: 10p.SEC PER DIVISION PEAK CURRENT: 5.9KA DISCHARGE PATH: POINT 1ATOPOINT 1B VERTICAL SCALE: 4.22KAPER DIVISION HORIZONTAL SCALE: 10p_ECPER DIVISION PEAK CURRENT: 11.6KA DISCHARGE PATH: POINT 2A TO POINT 2B VERTICAL SCALE: 10.55KAPER DIVISION HORIZONTAL SCALE: 10w.SECPER DIVISION PEAK CURRENT: ,16.9 KA DISCHARGE PATH: POINT 1ATOPOINT 1B FIGURE 25. TYPICAL TEST WAVEFORMS - CURRENT TRANSFER TESTS _] m_" DOUGLAS AIRCRAFT COHPANY C., _N_,.. P._G_ _,_ ACEE-O3-PR-9642 CONTRACT r, IASl-I4869 OF POOR QUALITY TABLE 5 PANEL RESISTANCE MEASUREMENTS CURRENT TRANSFER TESTS RESISTANCE (MI LLIOHMS) DISCHARGE RESISTANCE DISCHARGE CHECK BEFORE AFTER CURRENT PATH POINTS TEST TEST (KILOAMPE RES) 10.0 6.0 5.9 1A-1B 1A-1B 6.0 5.0 2A-2B 4.5 5.0 3A-38 7.0 T 1-T2 6.5 5.0 9.0 T3-T4 T2-T4 6.0 7.0 1A -T 2 3.0 2.0 1B-T4 5.0 2.0 6A-66 4A-4B 9.0 4.5 5 A_5B 13.0 4.0 6A-66 27.0 5.0 T 1-T2 7.0 6.5 T3-T4 9.5 5.0 T 1-T 3 8.0 7.0 6A-T 1 8.0 1.0 66-T3 20.0 3.0 2A-2B 1A-1B 6.0 6.0 11.6 2A-26 5.0 4.5 3A-36 5.0 4.5 T 1-T2 7.0 6.5 T3-T4 9.0 7.5 2A-T2 3.0 3.0 4.0 26-T4 3.5 4.5 3A-36 2 A-2B 4.5 4.5 3A-38 . 4.0 4A-4B 4.5 4.5 T 1-T2 6.5 6.5 T3-T4 7.5 8.0 3A-T2 4.0 4.0 3B-T4 4.5 4.5 16.9 1A-1B 1A-1B 6.0 5.5 2A-2B 5.0 5.0 3A-3B 4.5 4.5 T1-T2 7,0 6.5 T2-T4 7.0 7.0 6A-6B 4A-46 4.5 4.0 5A-56 4.0 4.0 6A-6B 5.0 5.0 T 1-T2 8.5 6.5 T1-T3 7.0 6.0 2A-2B 1A-1B 6.0 6.0 2 P_2B 4.5 4.5 3A-3B 4.5 4.5 T 1-T2 6.5 7.0 3A-3B 2A-2B 4.5 4.5 3A-3B 4.0 4.0 4A-46 4.5 4.0 T 1-T2 6.5 2.0 4A_B 3A-36 4.0 4.0 4A-46 4.5 4.0 5A-5B 4.5 4.5 T 1-T2 6.5 6.5 5A-5B 4A-48 4.0 4.0 5A-5B 4.5 4.0 6A-6B 5.5 5.5 T 1-T2 6.5 4.0 OR{G|NAL PAGE _J ACEE-O3-PR-9642 DOUGLAS AIRCRAFT COMPANY OF POOR QUALITY CONTRACT NAS1-14869 BEFORE TEST AFTER TEST FIGURE 26. TEST SETUP - SIMULATED LIGHTNING RESTRIKE TEST TO COMPOSITE PANEL (116-KA PEAK) DOUGLAS AIRCRAFT COMPANY ACEE-O3-PR-9642 CONTRACT NASI-14869 OF r;_-_R QUALITY VERTICAL SCALE: 40 KA PER DIVISION HORIZONTAL SCALE: 100/_SEC PER DIVISION PEAK CURRENT: 116KA FIGURE 27. TEST WAVEFORM - SIMULATED LIGHTNING RESTRIKE TEST TO CENTER OF PANEL ORIGINAL pAGE ACEE-03-PR-9642

DOUGLAS AIRCRAFT COMPANY

OF PoOR QUALrrY

CONTRACT NASI-14869

TABLE 6 PANEL RESISTANCE MEASUREMENTS LIGHTNING RESTRIKE TEST RESISTANCE (MILLIOHMS) RESISTANCE DISCHARGE BEFORE AFTER DISCHARGE CHECK CURRENT TEST TEST PATH POINTS (KILOAMPERES) 6.0 CENTER 1A-1B 6.0 1 16.0 5.0 OF 2A-2B 5.0 5.0 PANEL 3A-3B 4.5 T 1 -T 2 7.0 7.0 TO T3-T4 9.0 7.5 TABS T2-T4 7.0 4.5 1B,2B, 1B-T4 2.5 3.0 3B,4B, 4.5 3.0 4 A-4B 5B,6B 4.0 5.5 5A-5B 6 A-6B 5.0 5.5 7.0 T 1-T3 7.0 6B-T3 3.0 3.0 5O GFCT"':'T"_ _"J >_ '_'"'_" ._¢,11 ACEE-O3-PR-9642 DOUGLAS AIRCRAFT COMPANY OF POOR QUALITY CONTRACT NASI-14869 U m j _m FIGURE 28. COMPOSITE PANEL AFTER 116-KA SIMULATED LIGHTNING RESTRIKE TEST ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPANY ORIG'!NAE PAGE rS

CONTRACT NAS1-14869

OF POOR

QUALITY # 4,

¢

% J; h, FIGURE 29. COMPOSITE PANEL AFTER 116-KA SIMULATED LIGHTNING RESTRIKE TEST (CLOSE-UP VIEW) 5Z ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 The metal tabs were all removed from the lightning panel for visual examination.

There was no burning or other evidence of lightning damage at any of the joints. One joint that had only an edge seal was filled with fluid. The fluid entered the joint through a break in the seal when the panel was submerged for ultrasonic test. This incident emphasized the importance of using a faying surface seal to eliminate moisture intrusion and the resulting corrosion.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract IIASI-14869

SECTION 4

JOINT DEVELOPMENT COMPONENTS

The joint development component testing for the program will be completed

on successful testing of the Z5943453actuator (-l) and tie-rod (-501)

rudder fitting components. These tests will simulate critical load conditions

in the actuator ribs and tie-rod ribs discussed previously in Section 2,

Detail Design. Previously completed tests in this componentgroup included

the leading edge splice tests, leading edge attachment fatigue tests, spar

cap to cover panel attachment tests, and major attach fitting tests (see

References 7 through 9).

Detail part fabrication was completed for both the Z5943453-I and -501

specimen configurations during the reporting period and final assembly was

completed for the -l (actuator) component. The completed -l component is

shown in Figures 30 and 31. The -501 component during setup for final

assembly is shown in Figure 32. Test fixture installation and testing for

both components will be completed in January 1979. Fabrication of the full-

size rib tooling for the CVS will not be started until these tests are

successfully completed.

P_RECEDING PAGE BLANK NOT. FILterED"

ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 oRiGiNAL pAGE IS OF poOR QUALITY FIGURE 30. Z5943453-1 ACTUATOR HINGE RIB COMPONENT - VIEW LOOKING FORWARD AT SIMULATED REAR SPAR ACEE-03-PR-9642 Douglas Aircraft Company Contract NASI-14869 ORIC:_IAL RAGE ;3 OF POOR QUALITY \ \ ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 ORIEBNAL PAGE IS OF POOR QUALITY !

FIGURE 32. Z5943453-501 TIE-ROD RUDDER FITTING COMPONENT DURING SETUP FOR FINAL ASSEMBLY

Douglas Aircraft Company

ACEE-O3-PR-9642

Contract NASI-14869

SECTION 5

MECHANICAL PROPERTY TESTING

Mechanical property testing for the program was completed during the

reporting period with the completion of laminate fatigue testing and fracture

mechanics data testing on a variety of damaged and debonded specimens.

Previously completed laminate property tests for static tension and

compression allowables, fatigue data, and bolt bearing and shear-out data

were reported in References 5 and 8.

Test conditions and results for the completed fatigue tests and the fracture

mechanics tests are described in this section.

LAMINATE PROPERTY TESTS

Fatigue testing was completed on the Z3943432-505 quasi-isotropic laminate

sandwich beam specimens. The results of these tests are plotted in Figures

33 through 36. The test data are tabulated in Tables B-l and B-2 in

Appendix B.

A plot of the fatigue data at a stress ratio of R = -l.O is shown inFigure 33.

Tests were conducted at temperatures of 219°K (-65°F), ambient, and 350°K

(170°F). All specimens included a central hole of 0.635 cm (0.250 in) diameter

providing a width-to-hole-diameter ratio of 6. The average fatigue strength

exhibited by the specimens tested at ambient and 350°K was approximately 185

megapascals (26800 psi) at the one-life equivalent of 130,000 load cycles. The

specimens tested at 2190K exhibited a somewhat higher average fatigue strength

of 225 megapascals (32700 psi) or about a 22 percent increase in fatigue

strength.

A plot of the fatigue data at a stress ratio of R = 0.05 is shown in Figure 34.

Tests were conducted at temperatures of 219°K, ambient and 350°K as before.

These specimens also included a central hole of 0.635 cm. The data at this

ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUALI='_/

Contract NASI-14869

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Douglas Aircraft Company

Contract NASI-14869

stress ratio do not show the drop in fatigue strength with increasing load

cycles that was evident at R = -l.O. The data all fall within the scatter

band of the static tensile strength data obtained on similar sandwich beam

specimens tested in the same environments (Reference 5). It is apparent

that cyclic loading at a stress ratio of R = 0.05 has a negligible effect

on laminate strength.

The results of the ambient fatigue tests and the results of tests on similar

sandwich beams from the DC-IO composite rudder program IReference lO) are

compared in Figure 35. The test results from the rudder program are for

T300/5208 laminates made from unidirectional tape as opposed to the bi-

woven broadgoods used in the composite stabilizer sandwich beams. The data

show a difference in fatigue strength between the two types of prepreg of

approximately 39 percent at 130,000 cycles in favor of the unidirectional

tape. Unidirectional tape material has now been specified for the CVS

spar cap laminates.

The same data in terms of strain levels for each beam specimen at failure

are shown in Figure 36. No failures occurred below the design limit micro-

strain of 2000 even for those cases where the load cycles exceeded the one-

life equivalent of 130,000 cycles. Strain levels were computed using the

modulus values from the static tensile strength data (Reference 5).

It was concluded that the selected quasi-istropic laminate of T300/5208

unidirectional tape will provide more than adequate fatigue life for the

expected in-service loads and environments of the CVS.

FRACTURE MECHANICS TESTS

Fatigue testing was completed on the Z3943442-I debond tension specimens

and on the Z3943442-505 damaged tension specimens. The test data are

tabulated in Tables B-3 and B-4 in Appendix B.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

Photographs showing a typical test set-up and a close-up of the anti-buckling

plates used to prevent lateral instability at the laminate under compression

load are shown in Figures 37 and 38.

A plot of the fatigue data for the Z3943442-I debond specimens is shown in

Figure 39. All testing was conducted at a stress ratio of R = -l.O and at

temperatures of 219°K, ambient, and 350°K. All specimens included a laminate

debond area of approximately 1.27 cm (0.50 in.) diameter located in the

center of the test region. All specimens were subjected to 130,000 load

cycles (equivalent to one life-time of structural loading) at the design

limit microstrain of 2000 prior to undergoing the additional fatigue loading

shown. No failures occurred during the first life cycle tests and no changes

to the debond area were noted.

The second life cycle tests were run at higher load levels to establish the

fatigue characteristics of the laminate in the presence of debonds. None of

the debond specimens failed through the delaminated area but through the

minimum section adjacent to the tangent point of the 4.0 inch shoulder radius.

The X-ray, Figure 40, clearly indicates a significant stress concentration in

this area. An opaque liquid was applied to the specimen edges to accent the

damage. Figure 41 shows that no significant, consistent difference exists

between the results for the moisture conditioned specimens and those tested

"dry".

A plot of the fatigue data for the Z3943442-505damagedspecimens is shown

in Figure 42. All testing was conducted at a stress ratio of R = -l.O and

at temperatures of 219°K, ambient, and 350°K as before. All specimens included

a damagedarea in the center of the test region to provide a w/d ratio

(specimen width to damagesize) of 6.0. This ratio was selected to provide

a similar stress concentration effect to that in the sandwich beamfatigue

test specimens. The tests were conducted at strain levels ranging from

2000 microstrain (design limit strain) down to 1500 microstrain (84 occurrances

in one lifetime). None of the specimens failed at this strain level even

when tested for 260,000 cycles (two lifetimes).

DOUGLAS AIRCRAFT COMPANY

ACEE-O3-PR-9642

ORIGINAL PAGE IS

CONTRACT NASI -I 4869

OF POOR QUALITY

iIiIiiIlall !

FIGURE 37. TEST SETUP FOR DAMAGE AND DEBOND TEST

ACEE-O3-PR-9642

ORIGINAL

DOUGLAS AIRCRAFT COMPATIY

CONTRACT NAS1-14869 OF POOR

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FIGURE 38. CLOSE-UP VIEW OF PLATES USED TO PREVENT BUCKLING OF DAMAGE AND DEBOND SPECIMENS

ACEE-03-PR-9642

Douglas Aircraft Company

ORIG!NAL PAGE I_

Contract NASI-14869

OF POOR QUALITY

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DOUGLAS AIRCRAFT COHPANY

ACEE-O3-PR-9 cA°

COIITRACT NASI -I4869

OF POOR QUALITY

FIGURE 40. X-RAY OF DEBOND SPECIMEN SHOWING FATIGUE DAMAGE

ORIGINAL P&,GZ _

ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPArlY OF POOR QUALI_!

COrITRACT NASI-14869

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FIGURE 41. X-RAY OF DEBOND SPECIMEN AFTER FATIGUE TEST SHOWING STATIC FAILURE

ACEE-03-PR-9642

Douglas Aircraft Company

OF POOR (_UALITY

Contract NASI-14869

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ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

ORIGINAL PAe-E I_

OF POOR QUALITY

Trial impact energy tests were conducted on a piece of 0.052 inch thick quasi-

isotropic laminate to determine what energy levels would be required to

produce the desired damage level (Figure 43). These tests were conducted

using a standard Gardner impact tester with a 0.50 inch diameter anvil.

The desired level of damage was a 0.50 inch diameter area on the impact

side with complete penetration through the laminate. The energy levels

were varied from 10 inch-pounds to 70 inch-pounds. An energy level of

40-inch-pounds produced the desired damage and this level was used in

damaging the Z3943442-505 test specimens.

Fatigue damage sustained by the damage specimens progressed fairly rapidly

under cyclic load with failures occurring through the minimum net section of

the damage area. Damage propagation initially appeared as surface pitting

accompanied by a brownish discoloration of the epoxy with subsequent localized

buckling of the graphite woven fabric. Figures 44 and 45 show typical fatigue

effects. Figures 46, 47, and 48 show typical failures of the damage and debond

specimens together with C-scan records after test.

The results of the damaged specimen fatigue tests and the sandwich beam

fatigue tests are compared in Figure 49. An apparent 36.6 percent reduction

in strength is indicated at 130,000 cycles from 185 megapascals (26800 psi)

for the sandwich beams (drilled holes) to ll7 megapascals (16990 psi) for the

impact damaged specimens. The theoretical stress concentration factor from

Reference II for the sandwich beams is calculated as follows:

kte = 2 + (l - d/w) 3 = 2 + (l - 0.25/1.50) 3 = 2.58

The composite stress concentration factor from Reference II is:

ktc = 0.73 + 0.27 kte = 0.73 + 0.27 x 2.58 = 1.43

ACEE-03-PR-9642

Douglas Aircraft Company

Contract NASI-14869

CR,_G!t_AL p ....

OF POOR QU.,_L_T.y

A. IMPACT SIDE B. REVERSE SIDE FIGURE 43. TRIAL IMPACT ENERGY TESTS

OR:GII_!AL pAGE _3

ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPANY

OF POOR QUALITY

CONTRACT NASI -I 4869

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FIGURE 44. CLOSE-UP VIEW SHOWING FATIGUE DAMAGE IN IMPACT-DAMAGED SPECIMEN

DOUGLAS AIRCRAFT COMPAt|Y

ACEE-O3-PR-9642

COt|TRACT _la, Sl -14860

OE POOR QUALITY

FIGURE 45. X-RAY OF IMPACT-DAMAGED SPECIMEN SHOWING FATIGUE DAMAGE

ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPANY.

ORIGINAL F_ ;_'

CONTRACT NASI -I 4869

pOOR QUALITY

FIGURE 46. TYPICAL FAILURE MODES OF DAMAGE AND DEBOND SPECIMENS

ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPANY

CONTRACT NASI -I 4869

OF POCR QUALITY

PANEL 10-5

NOT [: FIGURE 47. TYPICAL C-SCANS SHOWING FATIGUE DAMAGE SUSTAINED BY DAMAGE AND DEBOND SPECIMENS

ACEE-O3-PR-9642

DOUGLAS AIRCRAFT COMPANY

ORIGINAL

CONTRACT NASI -I 4869

QUALI,'

OF POOR

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FIGURE 48. C-SCANS OF DAMAGE SPECIMENS AND MOISTURE CONTROL COUPONS'

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

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ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

The average measured stress concentration factor for the damagedspecimens

from Table B-4, Appendix B, is 1.75. From these factors, the strength

reduction from the drilled hole to the damage-induced hole is 18.3 percent

(l - 1.43/I.75). The damageis apparently more severe than is indicated

by the strain gage readings. This is probably due to local invisible damage

to the epoxy around the edge of the damagedarea.

Testing of the three Z3943443axial load cover panels with a transverse

center slit was completed. The data obtained from the tests are tabulated

in Table 7. One of the specimens mounted in the MTStest machine is

shown in Figure 50. The photograph also shows the environmental chamber used

for testing at 350°K (170°F) or 219°K (-65°F). Figure 51 is a photograph of

the ambient specimen after test indicating crack growth as a function of the

number of cycles. Final rupture occurred at lO,OOOcycles. A similar failure

occurred in the specimen tested at 350°K at 4670 cycles. The specimen tested

at 219°K did not fail after two lifetimes of load cycling. The first life cycle

test was run at a load level of I0231 Newtons (2300 pounds). This load level

was doubled for the second life cycle test in an attempt to induce crack

propagation. No failures occurred so the specimen was tested statically to

determine residual strength. Failure occurred at 194.04 MPa (28143 psi)

across the net section. It is evident from these tests that crack propagation

will not be a problem at low temperature (219°K) but will be a significant

consideration at room and elevated temperatures. These tests indicate the

need for further test and evaluation in this area.

Testing was completed on the three Z5943428-501damagedshear panels. The

data obtained from the tests are tabulated in Table 8.

The specimens were load cycled to the equivalent of two lifetimes (260,000

cycles) and then tested to failure to determine residual strength. The

testing was stopped after each 130,000 load cycles and the specimens visually

and sonically inspected to determine damage growth. Neither specimen exhibited

ORIGINAE P_-.Q_ i_

ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUAL],_

Contract NAS1-14869

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ORIGINAL PAGE IS

ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUALITY

Contract NASI-14869

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FIGURE 50. Z3943443 CENTER SLIT PANEL SPECIMEN IN TEST MACHINE

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ORIGINAL PAG_ _

ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUALITY

Contract NASI-14869

FIGURE 51. FAILURE OF Z3943443 CENTER SLIT PANEL SPECIMEN

OR;GIHAL PAGE I_

ACEE-O3-PR-9642

Douglas Aircraft Company

OF POOR QUALITY

Contract NASI-14869

TABLE 8 DAMAGED SHEAR PANEL TEST RESULTS Z5943428-501 _--DAMAGED AREA-'_ |(APPROX 2.54 CM) !

( _ (1.00 IN.) / I 0 762 CM :ii' i'i i!'i i!i'_'i:: :i! iii ii:'i i ili'i !: (0.300 IN.)_iiI:_l:_i:li:ili:l!:_:_if:il:il:ili:!f:il."_ p_ '_P

l ' I LDOUBLE-PLY OUTER LAMINATE ((00:00666 CNM.I

SINGLE-PLY INNER LAMINATE (0.033 CM) (0.013 IN.)

SECTION THROUGH DAMAGED AREA DAMAGED SHEAR STRESS SHEAR FLOW IN LAMINATE PERCENT ;RESIDUAL STRENGTH AT FAILURE AT FAILURE OF TEST AT FAILURE DESIGN p(1) TEMPERATURE PERCENT Nxy l"xy ULTIMATE PANEL MOISTURE LOAD (2) N/M LB/IN. MPa PSI oK o F NEWTONS POUNDS NUMBER CONTENT 202.77 29,410 1 1.31 219 -66 131,303 29,518 200,870 1147 162.29 23_38 AMBIENT AMBIENT 105,103 23,628 160,766 918 2 1.66 199.05 28,000 350 170 125,066 28,116 191,239 1092 3 0.99 i (1)RESIDUAL STRENGTH AFTER THE EQUIVALENT OF TWO LIFETIMES OF CYCLIC LOADING AT A LOAD RATIO OF R = --1.0.

(2)DESIGN ULTIMATE SHEAR FLOW = 132,746 N/M (758 LB/IN.)

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI -I4869

any change or growth in size of the damaged area. The residual strengths

were approximately 42 percent below the undamaged (and uncycled) panel at

ambient temperature and 22 percent below the undamaged panel at 350°K.

A typical failure mode for these panels is shown in Figure 52. All failures

occurred through the damaged area.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

•Q

FIGURE 52. FAILURE OF Z5943428-501 DAMAGED SHEAR PANEL

ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 SECTION 6 DESIGN VERIFICATION TEST COMPONENTS Tooling and detail part fabrication were continued on all elements of this component group. Final assembly operations were started as detail part availability permitted. The verification tests components are: • Z5943445 Concept Verification Panels • Z5943446 Concept Verification Spars • Z5943452 Attach Fitting Splice Specimens The fabrication status of these components is discussed in this section.

CONCEPT VERIFICATION PANELS Fabrication and assembly were continued on the Z5943445 combined load test panel (-l) and the acoustic test panels (-501). Current fabrication emphasis is on parts for the -l configuration, illustrated in Figure 53, because this activity is on the critical schedule path of the program. The -l panel will be tested in combined compression, in-plane shear, and lateral pressure to verify the CVS skin panel design concept. Fabrication of the CVS skin panel tooling will not begin until this test is successfully completed ....

Tool fabrication for the combined load test panel was completed on 6 November 1978. Some joggled regions were omitted on the sine-wave rib laminating mold and the initially cured rib-elements were rejected as a result. The mold was modified per engineering drawing requirements and replacement parts were fabricated.

Assembly of the Z5943445-I combined loads panel is in progress as shown in Figure 54. The cured skin panel, protected by peel plies on the outer surfaces, has been fit to the metal parts which will connect it to the test fixture.

The sine-wave rib and spar-web elements have been trimmed to size and fit together. Subsequent operations will cocure and bond the rib and spar-web junctions using graphite-epoxy attach angles, and mechanically fasten the skin panel to the sine-wave substructure elements. Finally, the graphite-epoxy panel will be mechanically attached to the test frame.

ACEE-03-PR-9642 Douglas Aircraft Company ORIGINAL Contrac. ilASI-14869 OF POOR QUALITY N X Uxy 72 INCHES II I I G _ Nx.y 61.5 INCHES

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J Z5943445 COVER PANEL COMBINED LOAD TEST SPECIMEN FIGURE 53.

ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 OF POOR QUALITY

\

ACEE-O3-PR-9642

Douglas Aircraft Company

ORIGINAL PAEL__S

Contract NASI-14869

OF POOR QUAL!_( Following completion of final assembly and instrumentation, the component will undergo a four-day moisture exposure period to achieve a minimum moisture content of one percent in the honeycomb sandwich skin facings. The component will then be tested under combined compression,in-plane shear, and lateral pressure. Test completion is presently projected at 23 February 1979.

CONCEPT VERIFICATION SPARS Tool fabrication for the Z5943446-I concept verification spar component was essentially completed during the reporting period. The machined aluminum alloy lower half of the laminating mold together with the mold end piece is shown in Figure 55a. The upper half of the laminating mold is shown in Figure 55b.

The upper mold half consists of a rigid aluminum alloy component (which in conjunction with the lower half will maintain the spar cap contours and bevels) and a cast rubber facing. Side pressure plates will also be provided to transmit autoclave pressure to the spar cap flanges during the cure cycle. This tooling concept is further illustrated in Section 7, Tool Design. Fabrication of the spar tooling for the CVS will not be started until the Z5943446 spar component is successfully tested.

ATTACH FITTING SPLICE SPECIMENS Fabrication of the Z5943452-I, -501, and -503 splice specimens, three each, was completed during the reporting period and the specimens were placed in a moisture chamber to start a 30-day exposure period. The completed -l and -501 specimens are shown in Figure 56. The specimens represent the bonded/ bolted splice joint at the root-end of a CVS spar between the laminates and the titanium alloy fitting as shown in Figure 4.

The thicker -l specimens, Figure 57, represent the joint interface between the thick laminated spar-web and the titanium alloy fitting. The thinner -501 specimen, Figure 58, represents a double-lap scarf-joint at the skin flange of the spar in the same region. The -503 specimen represents a single-lap scarf-joint at the skin flange of the spar. This single-lap scarf-joint configuration is being used in the CVS spar drawings to simplify the spar layup and processing.

ORIP=I,_IAL P._ _ ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 OF POOR QUALITY f _ -_, - (a) LOWER MOLD HALF AND END-PIECE (b) UPPER MOLD HALF FIGURE 55. Z594344.6 SPAR COMPONENT LAMINATING MOLD FIGURE 56. Z5943452-1 AND-501 SPECIMEN ORIGINAL PAGZ ;."_; ACEE-O3-PR-9642 OF POOR QUALITY Douglas Aircraft Company Contract NASI-14869 FIGURE 57. Z5943452-1 SPAR WEB SPLICE FIGURE 58. Z5943452-501 SPAR SKIN-FLANGE SPLICE Douglas Aircraft Company ACEE-O3-PR-9642 Contract NASI-14869 CF POOR QUALITY Resin and void content samples indicated good laminate quality in all specimens as shown in the following table.

Nominal Thi ckness- Resin Content Void Content Specimen Sample Number Number Layers (inches) Weight % V o l u_=_ ___ _ me_==%__ _=.=_ Z5943452-I 6O 1 25.49 1.70 | (0.780) 26.77 1.30 26.13 I 1.50 Average iZ5943452-501 28 1 28.02 0.90 2 28.41 i 0.90 (0.364) 28.21 , 0.90 Average ' 34 1 Z5943452-503 26.34 0.90 2 27.13 O. 57 (0.422) 26.74 0.74 Average all 27_+5 2.00 max IDPS Requirement l, Computer runs have been made using an infinite series solution to Fick's Second Law of Diffusion to show moisture distribution in different laminate thicknesses for various exposure times. Average moisture content vs.

thickness is plotted in Figure 59 for various times of exposure to 170°F and I00 percent relative humidity. The plot indicates that an average moisture content over one percent will be achieved in laminates up to a thickness of about 0.35 inches during the 30-day exposure period. Additional studies indicate that discernible moisture will not penetrate to the center of the thicker section (-I) in the lifetime of the DC-IO aircraft.

ORIGINAL PAGE |8 ACEE-O3-PR-9642 OF POOR QUALITY Douglas Aircraft Company Contract NASI-14869 1.5 A F- Z uJ (,3 ro- w F- -r (3 i W v F- 1.0 Z w I.- z (..)

uJ I- :E UJ (3 ,,( 0.5 rr uJ > 90 DAYS ,< 60 DAYS ) 30 DAYS 15 DAYS 0.9 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 LAMINATE THICKNESS (INCHES) FIGURE 59. AVERAGE MOISTURE CONTENT VERSUS LAMINATE THICKNESS FOR VARIOUS TIMES OF EXPOSURE TO 170°F AND 100 PERCENT RELATIVE HUMIDITY

Douglas Aircraft Company ACEE-O3-PR-9642 Contract NASI-14869 SECTION 7 TOOL DESIGN Tool design activities are proceeding on the released engineering drawings.

With the exception of the station 295 rib tools, fabrication of hard tooling will not be started until appropriate joint development or concept verifi- cation tests are completed. The station 295 rib tools will be fabricated directly on completion of the tool design. One set of station 295 rib parts for the box-beam verification component will be fabricated to prove the rib tooling design concept.

SKIN FABRICATION TOOLING The skin panel tool design is in progress based on the requirements of drawing number AMC7840. The tool concept is illustrated in Figure 60. The outer surface of the skin panel will be the tooled surface to facilitate net molding of the recesses for access doors, antenna bays, and leading and trailing edge attachments. The basic mold surface will be a I/4 inch thick steel plate rolled to contour. The plate will be stud-welded at the back surface and bolted to an egg-crate supporting structure. The supporting structure will be designed to facilitate air circulation and rapid heat-up in the autoclave. Honeycomb locators which index to the mold surface will be used to facilitate placement of the core segments during the skin panel buildup. A caul plate will be used during the cure cycle to maintain a smooth inner surface on the cured part.

SPAR FABRICATION TOOLING The design of spar tools is in progress based on the requirements of drawing numbers AMC7845, AMC7847, AMC7848, AMC7849, and AMC7893. The spar tool concept is illustrated in Figure 61. A machined aluminum alloy tool for the upper surface will be used. The titanium fittings of the spar assembly will be held in position by locating bolts through the end piece of the tool.

The lower half of the laminating mold will consist of a cast rubber mold ACEE-03-PR-9642

Douglas Aircraft Company

Contract NASI-14869

ORIGINAL PAG'_" _

OF poOR QUALIT'I' it ,,,// ", \ / / \ ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

\

FIGURE 61. TOOLING CONCEPT FOR COMPOSITE VERTICAL STABILIZER SPARS

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

backed up by a rigid aluminum alloy plate. The rubber mold will be Teflon

coated to preclude laminate sticking after a cure cycle. The side pressure

plates will maintain the lofted bevels and contours of the spar cap flanges

which interface with the skin panels.

RIB FABRICATION TOOLING

The design of the base rib tool is in progress based on the requirements of

drawing number AMC7853. The tooling concept is illustrated in Figure 62. An

aluminum female mold will be used to control the outer loft surface of the part.

A fiberglass caul plate will be used to control the internal surfaces to

insure a good fit with the spar ends. A split base rib was designed to

provide a tolerance take-up feature at a centerline overlap splice joint.

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

OF POOR QUAL;_-_/, i i /--- CAUL PLATE FOR BIN-5 / /---CAUL PLATE /FOR B/N-3 'N-1 PLM1 _'- iN-5 PLASTIC LAMINATING MOLD MAKE FROM -,._...

/ SECTION VIEW LAY UP MANDREL FIGURE 62. TOOLING CONCEPT FOR COMPOSITE VERTICAL STABILIZER BASE RIB

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

i I _ k ¸_

SECTION 8

COST ANALYSIS

An economic analysis was completed using the technical and manufacturing

information developed for the current structural configuration (e.g., mechanically

attached skin panels, sine-wave stiffened shear panels). The assumptions

and analysis guidelines reported previously (Reference 9) were applied in this

current analysis. Experience and information obtained in the composite upper-

aft rudder program (Reference 6) were also applied.

The economic analysis results are expressed in terms of the number of composite

stabilizers to be manufactured in a production mode to achieve cost parity

with the current unit costs of the metal stabilizer. This cost cross-over

point was estimated at lO0 composite stabilizer units at the start of the

program. The current analysis indicates that the cost cross-over point will

be achieved after approximately 32 units are produced (see Figure 63).

New labor estimates were made to reflect the current design concept. The

estimates associated with the economic analysis were selectively extracted

from the overall program labor estimates. Development costs were excluded.

The manufacturing data shown in Table 9 represents estimates of the eight

stabilizers incorporating the present design concept. These estimates will

be updated as actual cost data are accrued. In Table 9, the manufacturing

labor hours and the planning hours were derived from the current estimates.

The recurring tooling was allocated arbitrarily to the eight units as

sustaining tooling. Sustaining Engineering and Inspection/NDT were assumed

to be the same as shown in the prior analysis, Reference 9. Manufacturing Tl

labor hours (recurring labor hours for the first production unit) derived

from the base estimate are shown in Table lO together with production progress

(learning) curve assumptions used in the cost cross-over analysis.

P.RECEDIIN'G PAGE BLA_'_K NOT FIL'VIED _

ACEE-O3-PR-9642

Douglas Aircraft Company

Contract NASI-14869

ORIGINAL

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ORIGINAL PAGE iS

ACEE-O3-PR-9642

OF POOR QUALITY

Douglas Aircraft Company

Contract NASI-14869

TABLE 9 RECURRING LABOR HOURS FOR DC-10 COMPOSITE VERTICAL STABILIZERS MANUFACTUR ING NO. 2 NO. 3 NO. 4 NO. 5 NO. 6 NO. 7 NO. 8 TOTAL FUNCTION NO. 1 COMPOSITE STRUCTURE 11,153 9,050 10,073 9,348 8.968 8,692 8,475 8296 74,055 I_E/TIP/RUDDER (102) (99) 133 133 133 132 132 132 795 METAL DETAILS (387) (383) 512 512 511 511 511 511 3,068 STRUCTURE ASSEMBLY 286 286 286 286 286 286 286 285 2287 BOX ASSEMBLY 2,377 2,377 2,377 2.377 2,377 2.377 2.377 2.377 19,016 1.164 1,164 1,164 1.164 1,164 1,164 1,164 1,164 9.312 FINAL ASSEMBLY 542 462 _499 473 459 449 440 432 3,756 PLANNING 1,770 1,600 1,410 1,300 1250 1220 1,190 1,170 10.910 INSPECTION/NDT 700 700 700 700 700 699 699 699 5,597 TOOLING 260 250 240 230 220 190 150 140 1.680 ENGINEERING TOTA L 18 252 15 ,889 17,394 16.52 3 16,068 15,720 15,424 15,206 130,476

ACEE-O3-PR-9642

ORIGINAL PAGE IS

Douglas Aircraft Company

Contract NAS1-14869

OF POOR QUALITY

TABLE 10 PROJECTED T1 LABOR HOURS FOR DC-10 COMPOSITE VERTICAL STABILIZER T1 AVERAGE ESTIMATING DERIVED FROM CURVE AND MANUFACTURING FUNCTION UNITS 1 THROUGH 8 FACTORS COMPOSITE STRUCTURE 9257 80/84%* LE/TIP/RUDDE R 133 9O% METAL DETAI LS 511 9O% STRUCTURE ASSEMBLY 286 8O% BOX ASSEMBLY 2377 8O% FINAL ASSEMBLY 1164 8O% PLANNING FACTORED 6.7% I NSPE CTI ON/N DT FACTORED 13.6% TOOLING 8 K ENGINEERING 126 91% *ASSUMES AN 80 PERCENT LEARNING FROM T1 TO T100 AND AN 84 PERCENT FROM T101 TO T200

ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 SECTION 9 QUALITY ASSURANCE Non-destructive inspections (NDI) were performed on the Z5943434-501 sine- wave shear web component and the Z5943445-21 combined load test panel using ultrasonic, radiographic, and Fokker Bondtester NDI methods for the various types of construction (e.g., flat laminates, sine-wave laminates, and honeycomb sandwich regions).

The Z5943434-501 sine-wave shear web component was evaluated using an ultrasonic C-scan reflector technique to test the flat areas of the web, an ultrasonic C-scan pulse-echo technique to evaluate the solid laminate caps, and a pulse-echo digital thickness gage to contact scan the convoluted web areas. These ultrasonic tests indicated the test component to be of acceptable quality.

The Z5943445-21 combined load test panel was evaluated using X-radiography to view core quality and ultrasonic C-scan to detect discontinuities in the solid laminate and the skin-to-core bond lines. The solid laminate areas that could not be reached by the thru-transmission fixture were evaluated using an ultrasonic pulse-echo thickness gage. Skin-to-core bond areas were evaluated using the Fokker Bondtester. The core closures and skin- to-core bonds were judged to be of acceptable quality by the X-ray and Fokker Bondtester inspections.

Interpretation of the ultrasonic inspections was difficult because of the presence of peel plies on both surfaces of the panel and the lack of applicable inspection standards. The peel plies will not be removed until hole preparation is complete during final assembly of the panel and test fixturing. The panel was therefore examined ultrasonically for consistent homogeniety in both the solid laminate and honeycomb core regions. The inspection indicated several areas of porosity on the solid laminate. These regions were sufficiently porous to attenuate digital thickness gage readouts.

ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 Resin and void samples will be obtained from these regions after the panel tests are completed to help define QC acceptance levels in the CVS parts. These test results are also being considered during planning of the necessary CVS inspection standards.

Incoming quality assurance tests were conducted on 23.7 kilograms (52.2 pounds) of bi-woven material. The material met all specification requirements as shown in Table II.

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ACEE-O3-PR-9642 Douglas Aircraft Company Contract NASI-14869 SECTION lO REFERENCES l • Grubb, D. W., "DC-IO Composite Vertical Stabilizer Design Criteria and External Loads," McDonnell Douglas Corporation Report Number MDC J7718, September 1977.

o Abelkis, P. R., "DC-IO Aft Section Fatigue Test, Volume I, Spectrum Derivation," McDonnell Douglas Corporation Report Number DAC 67725, 23 July 1970.

Q Hunter, P. A., "An Analysis of VGH Data From One Type of Four-Engine Turbojet Transport Airplane During Commercial Operations," NASA TN D-4330, February 1968.

• Hunter, P. A. and M. W. Fetner, "Maneuver Accelerations Experienced During Routine Operations of a Commercial Turbojet Transport Airplane," NASA TN D-18Ol, May 1968.

• "Advanced Composite Vertical Stabilizer for DC-IO Transport Aircraft," Contract NASI-14869, Sixth Quarterly Technical Progress Report, McDonnell Douglas Corporation, Report Number ACEE-O3-PR-8549, 20 October 1978.

o "Advanced Composite Vertical Stabilizer for DC-IO Transport Aircraft," Contract NASI-14869, Second Quarterly Technical Progress Report, McDonnell Douglas Corporation, Report Number ACEE-O3-PR-7240, 15 October 1977.

• "Advanced Composite Vertical Stabilizer for DC-IO Transport Aircraft," Contract NASI-14869, Third Quarterly Technical Progress Report, McDonnell Douglas Corporation, Report Number ACEE-O3-PR-8332, 20 January 1978.

PRECEDING PAGE BLANK ": "r, • __.,O,, F,"r:_{_

Douglas Aircraft Company ACEE-O3-PR-9642

Contract NASI-14869

-% Q "Advanced Composite Vertical Stabilizer for DC-IO Transport Aircraft," Contract NASl-14869, Fifth Quarterly Technical Progress Report, McDonnell Douglas Corporation, Report Number ACEE-O3-PR-8484, 26 July 1978.

ga "Advanced Composite Vertical Stabilizer for DC-IO Transport Aircraft," Contract NASl-14869, Fourth Quarterly Technical Progress Report, McDonnell Douglas Corporation, Report Number ACEE_O3-PR-8394, 24 April 1978.

lO. Lehman, G. M., et al, "Advanced Composite Rudders for DC-IO Aircraft - Design, Manufacturing, and Ground Tests," NASA CR-145068, April 1976.

II. Hart-Smith, L. J., "Bolted Joints in Graphite-Epoxy Composites," Report NASA CR-144899, Contract NASl-13172, January 1977.

Ii0

Douglas Aircraft Company ACEE-O3-PR-9642

Contract NASI-14869

APPENDIX A

ENGINEERING DRAWINGS

III ORIGINAL PAGE IS OE POOR QUALITY Ill Itl I AMC7840 Dougl a.

Contra( 4 .3 ORIGINAL PAGE IS OF POOR QUALITY . r U (;

[]

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PRECEDI _ E

omGmSA_ pAGE is

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ll_ I is t

"NTER SPAR ASSEMBLY (SHEET 3 OF 10)

Douglas Aircraft Company

ACEE-O3-PR-9642

Contract NASI-14869

I 17 I

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Douglas Aircraft Company ACEE-O3-PR-9642 Contract NASI-1486g

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ACEE-O3-PR-9642

Contract NASI-14869

4.0 I 3_

OF POOR QUALITY TABLE NO. M I PLY LAYUP JEN_ PLY _ "IBER NO 3 "R E)_R _.45 P'M.I o/go P,..*O. _ 45 P_9.1 P'zB.I P_7.i

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