NASA Contractor Report 3517 

Development of an Advanced 

Composite Aileron for the 

L-101 1 Transport Aircraft 

C. F. Griffin and E. G. Dunning 

CONTRACT NASl-1506y 

FEBRUARY 1982 
TECH LIBRARY KAFB, NM 

I 1#111 Ill11llll Ill11llll Ill1llll HI II 

0062222 

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I 
-- -y 
NASA Contractor Report 3517 

Development of an Advanced 

Composite Aileron for the 

L-101 1 Transport Aircraft 

C. F. Griffin 

Lockheed Corporation 

Burbank, California 

E. G. Dunning 

Avco Corporation 

Nasbde, Tennessee 

Prepared for 

Langley Research Center 

under Contract NAS l-15069 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1982 

FOREWORD 

This report was prepared by the Lockheed-California Company, Lockheed 

Corporation, Burbank, California, under contract NASl-15069 and it is the 

summary report for the aileron program. The program is sponsored by the 

National Aeronautics and Space Administration (NASA), Langley Research Center. 

The Program Manager for Lockheed is F. C. English and the Project Manager 

for NASA, Langley is H. L. Bohon. The Technical Representative for NASA, 

Langley is H. A. Leybold. 

The following Lockheed personnel were principal contributors to the 

program: C. Griffin, Engineering Manager; L. Fogg and J. Pearson, Structural 

Analysis; R. Stone, Materials and Processes; J. Ekvall, Design Allowables; 

J. Soovere, Sonic Fatigue Analysis; J. Salvaggio and W. Parks, Mass Properties; 

D. Paschal, Design; S. Bocarsley, S. Langenbeck and D. Thompson, Testing; 

W. Newcomb, Producibility. 

The following Avco personnel were principal contributors to the program: 

W.L. Cobbs, Program Manager; E. Dunning, Project Engineer; R. Legg, Process 

Development; R. Autery, Design. 

iii 

TABLE OF CONTENTS 

FOREWORD. .............. 

LIST OF FIGURES ........... 

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

SUMMARY ................ 

INTRODUCTION ............ 

MEASUREMENTVALUES. ......... 

COMPOSITE AILERON DESIGN. ...... 

General Description. ....... 

Design Criteria. ......... 

Detailed Design. ......... 

Analysis ............. 

VERIFICATION OF STRUCTURAL INTEGRITY. 

Design Data. ........... 

Concept Verification ....... 

Ground Tests ........... 

Flight Test. ........... 

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

AILERON FABRICATION ........................ 

TOOLING .............................. 

PROCESS DEVELOPMENT ........................ 

DETAIL FABRICATION. ........................ 

Ribs .............................. 

Spar .............................. 

Covers ............................. 

ASSEMBLY .............................. 

Main Rib Subassembly ...................... 

Spar Subassembly ........................ 

Final Assembly ......................... 

QUALITYASSURANCE ......................... 

iii 

Vii 

viii 

1 

1 

3 

3 

3 

4 

5 

9 

10 

10 

13 

17 

20 

21 

21 

21 

21 

23 

24 

25 

25 

25 

25 

26 

27 

V 
TABLE OF CONTENTS (Continued) 

MANUFACTURING COST ...................... 28 

Producibility Cost Analysis ................ 29 

Projected Production Cost ................. 31 

CONCLUSIONS ......................... 32 

REFERENCES .......................... 34 

vi 
LIST OR FIGURES 

Figure 

1 Masterschedule ...................... 2 

2 Inboard aileron - existing metal design. ......... 3 

3 Advanced composite aileron assembly. ........... 5 

4 Aileron cover design ................... 6 

5 Hinge/actuator fitting installation. ........... 7 

6 Aileron assembly ..................... 8 

7 Typical defect and damage tolerance data ......... 12 

8 Summary of subcomponent structural tests ......... 13 

9 Summary of assembly structural tests ........... 14 

10 Setupforribstatictest ................. 15 

11 Rib/spar/fitting specimen installed in test machine. ... 16 

12 Aileron ground test set-up ................ 19 

13 Flight conditions investigated .............. 20 

14 Manufacturing flow diagram ................ 22 

15 Riblayup ......................... 23 

16 Spar preplying operation ................. 24 

17 Spar preply template and cure tool ............ 26 

18 Drill motion with hydraulic check attachment ....... 27 

19 Actual learning curve development. ............ 31 

vii 
LIST OF TABLES 

Table 

4 

8 

9 

Comparison of Composite Aileron to Aluminum Aileron . . . . 9 

T300/5208 Graphite/Epoxy Fabric Characterization 

TestData......................... 11 

T300/5208 Graphite/Epoxy Fabric (45"/O"/135"/O"/45") 

Laminate Data . . . . . . . . . . . . . . . . . . . . . . . 11 

Graphite/Epoxy Tape and Fabric Lamina Design 

Allowables. . . . . . . . . . . . . . . . . . . . . . . . . 12 

Frequency Comparison. . . . . . . . . . . . . . . . . . . . 17 

Torsional Stiffness Comparisons . . . . . . . . . . . . . . 18 

Aluminum Aileron Cost Compared to Composite Aileron 

Cost Based on Current Materials Cost. . . . . . . . . . . . 30 

Labor Cost Comparison . . . . . . . . . . . . . . . . . . . 30 

Projected Cost Comparison . . . . . . . . . . . . . . . . . 32 

viii 
SUMMARY 

This report summarizes the engineering and manufacturing development 

activities of an advanced composite inboard aileron for the L-1011 commercial 

transport aircraft. 

The composite aileron is a multirib assembly with graphite/epoxy tapesyntactic core sandwich covers, a graphite/epoxy tape front spar, and 

graphite/epoxy fabric ribs. The design has a weight savings of 23 percent 

compared to the aluminum aileron. Due to the simplicity of the design, the 

composite aileron has 50 percent fewer parts and fasteners than the metal 

aileron. 

The structural analyses conducted on the composite aileron in combination 

with materials tests and concept verification tests have verified that the 

structural integrity of the aileron meets or exceeds the design requirements. 

Final substantiation of structural integrity was obtained by static and 

damage growth/fail-safe ground tests and a flight check-out. 

Very simple and practical manufacturing techniques have been developed 

for the composite aileron. Utilization of male tools, curing to net part 

size, and use of a single cure cycle for all of the aileron composite parts 

has resulted in the fabrication of high quality parts. Drilling and machining procedures developed for the composite aileron have been proven to be as 

simple and reliable as techniques currently employed for aluminum structures. 

INTRODUCTION 

The broad objective of NASA's Aircraft Energy Efficiency (ACEE) Composite 

Structures Program is to accelerate the use of composite materials in aircraft structures by developing technology for early introduction of 

structures made of these materials into commercial transport aircraft. This 

program, one of several which are collectively aimed toward accomplishing 

this broad objective, has the specific goal to demonstrate the weight and 

cost/saving potential of secondary structures constructed of advanced composite materials. The secondary structure selected for the program is the 

inboard aileron of the Lockheed L-1011 transport aircraft. 

Use of trade names or names of manufacturers in this report does not 

constitute an official endorsement of such products or manufacturers, either 

expressed or implied, by the National Aeronautics and Space Administration. 
The scope of this program is to design, fabricate, qualify, and 

certificate a composite inboard aileron; to test selected subcomponents 

to verify the design; to fabricate and test two ground test articles; to 

fabricate and install five shipsets of inboard composite ailerons; and to 

gather flight service data on the five shipsets. (The original contract 

required fabrication of ten shipsets, but a contract change reduced this to 

five.) 

The Lockheed-California Company is teamed with Avco Aerostructures 

Division of Avco Corporation to accomplish program goals. Lockheed designed 

the aileron, conducted the materials, concept verification, and ground tests, 

and will evaluate inflight service experience. Avco developed the manufacturing processes, fabricated the test specimens, and fabricated the ground 

test and flight articles. 

As shown on the master schedule, figure 1, the program was conducted in 

six nonsequential tasks. Task I, Engineering Development, and Task II, Design 

and Analysis, are the portions of the program wherein the composite aileron 

design was formulated and subcomponents fabricated and tested to verify design concepts and fabrication procedures. During Task III, Manufacturing 

Development, and Task IV, Ground Test and Flight Checkout, production 

quality manufacturing tools were constructed, and two full-scale ailerons 

were fabricated and tested. A production run of five shipsets were fabri+ 

cated during Task V, Aileron Manufacture, to provide manufacturing and cost 

information. In Task VI, Flight Service, inspection and maintenance data 

TASK 1977 1978 1979 1980 1981 1982 1983 1984 1985 1986 

. 

: 

. 

I ENGINEERING : 

DEVELOPMENT : 

: 

. 

: 

II DESIGN AND : 

ANALYSIS : 

: 

. 

I : 

Ill MFG : 

DEVELOPMENT : 

: 

IV GROUND TESTS 

V MANUFACTURE 

VI FLIGHT 

SERVICE I 

: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 

:: 

:: : 

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..J 

: 

Figure 1. - Master schedule. 

2 
will be gathered on the five shipsets to assess their potential for economical operation in routine service. The work performed during this program 

is intended to provide the data required to progress toward a production 

commitment. 

This report summarizes the major technical achievements of the program. 

Additional detailed information pertaining to the various tasks within the 

program may be found in References 1 through 4. 

MEASUREMENT VALUES 

All measurement values in this technical report are expressed in the 

International System of Units and customary units. Customary units are 

used for the principal measurements and calculations. 

COMPOSITE AILERON DESIGN 

General Description 

The inboard aileron is located on the wing trailing edge between the 

outboard and inboard trailing edge flaps, as indicated in figure 2. It is 

ALUMINUM 

L.E. SHROUD 

\ er’n EIYU ‘^‘SINGS rnln 

FIBEF-’ ’ HINGE AND / - 

ACTUATOR FITTINGS 

Figure 2. - Inboard aileron - existing metal design. 

3 
supported from the wing at two hinge points and is actuated by three hydraulic 

actuators. It is a wedge-shaped, one-cell box, thinning slightly from root to 

tip. At the front spar the aileron is 2337 mm (92 in.) in length and approximately 254 mm (10 in.) deep. The width of the aileron is 1270 mm (50 in.). 

The current aluminum aileron, shown in figure 2, consists of a front 

spar, a rear spar, and upper and lower skins, joined by three hinge ribs and 

fifteen intermediate and end ribs. A trailing edge wedge, leading edge 

shroud, and end fairings complete the assembly. 

Design Criteria 

The approach used for determining the composite aileron design concept 

and materials was to define the design criteria, create alternate designs, conduct comparative tests, evaluate the alternatives against the criteria, and select and implement the best alternative. A brief description of the criteria 

established for the Advanced Composite Aileron (ACA) is presented below: 

l The ACA must be a direct replacement for the current aileron in form, 

fit and function. 

l A minimum of 20 percent weight savings must be obtained. 

l The ACA must be cost competitive with the existing aileron on a 

production basis. 

l Materials selected for the ACA must not be severely degraded by 

environmental effects such as temperature r219.3 K(-65'F) to 335.4 K 

(180°F)] or moisture (1 percent by weight minimum). 

Aileron external loads are related to hinge moment capabilities. 

Deflection of the aileron causes an airload hinge moment which is reacted 

by the actuators. Larger deflections give larger airloads until the maximum 

hinge moment capable of being reacted by the actuators is reached. In 

addition to airloads, the aileron must be designed to withstand acoustic 

loading up to 135 dB/Hz. 

Typical of secondary structural components, the aileron is not a 

strength critical component. Instead its design is governed by stiffness. 

The composite aileron must have the same torsional stiffness as the metal 

aileron to preclude self-induced vibrations. 

The aileron structure is designed to be fail safe for limit flight loads 

in accordance with Federal Aviation Administration (FAA) requirements. Also 

it is designed not to separate from the aircraft after any one of several 

failure or jam conditions premised in FAA requirements. 

4 
Detailed Design 

Design trade-off studies and materials evaluation tests were conducted 

to determine component designs and materials which best meet the program 

goals. The selected composite aileron design, illustrated in figure 3, is a 

multirib configuration with single piece upper and lower covers mechanically 

fastened to the substructure. Three composite materials are utilized in the 

aileron design: Thornel 300/Narmco 5208 graphite/epoxy unidirectional tape 

LO.19 mm/ply (0.0075 in./ply)]; Thornel 300/Narmco 5208 graphite/epoxy bidirectional fabric; and Hysol ADX-819 syntactic epoxy. The selected resins 

have a cure temperature of 450 K (350'F) for excellent environmental resistance. The [0.19 mm/ply (0.0075 in/ply)] tape thickness was selected for the 

aileron for optimum weight savings and minimum lay-up time. The bidirectional 

fabric offered lower production costs through reduced number of plies required 

and improved drapability. Syntactic epoxy , which is a film adhesive filled 

with hollow glass microballoons, is used as a core to create a lightweight 

sandwich structure. 

The aileron covers are approximately 2438 mm (96 in.) by 914 mm (36 in.). 

The structure, shown in figure 4, is a syntactic core sandwich consisting of 

three plies of 0.19 mm (0.0075 in.) thick graphite/epoxy tape on each side of 

a 0.953 mm (0.0375 in.) syntactic/epoxy core. The orientation is 45"/0"/135" 

with the 0" ply in the spanwise direction. The syntactic core over the three 

main ribs is replaced by five-ply unidirectional tape doublers. Additional 

doublers are cocured to the inner surface at the ribs and spars to assist in 

BACK-UP I 

ALI 

FllTlNGS 

U”‘NUIM / 

T.E. WEDGE 

KEVLAR AND 

IRE 

FRONT SPAR/ RIBS ’ 

GR/EP TAPE GR/EP FABRIC 

Figure 3. - Advanced composite aileron assembly. 

5 
SYNTACTIC CORE 

(45/O/l 35/SY N)s 

SURFACE DOUBLERS 

GWEPR/ TAPE 6 FABRIC 

5$9+/l 36T/goF) 

SECTION A-A 

INNER DOUBL 

GR/EP TAPE 

(95/85/90/85/95) 

Figure 4. - Aileron cover design. 

load transfer between covers and substructure and provide thickness buildup 

required for countersunk fasteners. Forward of the front spar datum, the 

syntactic core is omitted and inner surface doublers are added to help carry 

the spar bending loads near the main ribs. 

The front spar is a constant thickness channel section constructed of 

ten plies of unidirectional graphite/epoxy tape LO.19 mm/ply (0.0075 in./ply)] 

laid up at (45"/0"/135"/90"/0") where the 0" direction is spanwise. The 

flange width of the spar caps is" increased locally at the main rib locations 

to provide mounting surfaces for the main ribs and rib backup fittings. 

Flanged holes are provided in the spar web for access and inspection purposes. 

Front spar bending moments are shared by both the spar cap and the 

leading edge of the cover which has local doublers at the main rib-front 

spar intersections. This design approach yields a very efficient fail-safe 

design. The ribs, hinge/actuator fittings, and leading edge shroud supports 

act as stiffeners on the spar shear web; thus, no additional stiffening 

elements are necessary. 

The rear spar is a constant thickness El.02 mm (0.040 in.)] channel section fabricated of 7075-T6 clad aluminum. It has aft facing-flanges to which 
both the covers and trailing edge wedge are attached. Aluminum was selected 

for this component because utilization of composites would be too costly for 

the small amount of weight saved. 

The main ribs are located at each of the three hinge/actuator fitting 

locations to transfer concentrated loads from the fittings to the aileron 

covers and spar. Aileron hinge/actuator fittings attach to the forward face 

of the front spar web, and the rib to spar joint is completed by means of 

aluminum backup fittings which attach to each side of the ribs as shown in 

figure 5. 

The main ribs are a channel section configuration constructed with four 

plies of graphite/epoxy bidirectional fabric IO'.36 mm/ply (0.014 in/ply)] 

oriented at 45'/90°2/45“ where 0" is the lengthwise direction of the rib. 

Five plies of unidirectional graphite/epoxy tape are added to the rib caps to 

increase the stiffness and strength of the ribs. 

In addition to the three main ribs, the aileron has five intermediate 

ribs and two closeout ribs which support the covers and react the air 

pressure loads. Intermediate and closeout ribs are constant thickness 

channel sections consisting of five plies of graphite/epoxy bidirectional 

fabric oriented at 45"/9O"/135"/9O"/45" where the 0" is again the lengthwise 

direction of the rib. Each rib contains five flanged lightening holes. The 

ribs are flanged on all four edges to eliminate the necessity of using separate 

clips to attach the ribs to the spars. 

Figure 5. - Hinge/actuator fitting installation. 

7 
A photograph of the aileron assembly is shown in figure 6. The upper 

surface, ribs, and spars are permanently fastened with titanium Triwing 

screws and stainless steel Hi-Lok collars. The removable lower surface, 

trailing edge wedge, leading edge shroud , and end fairings are attached 

with the same type of screws but with stainless steel nut plates attached 

to the substructure with stainless steel Cherry rivets. The trailing edge 

wedge, front shroud and fairings are identical to those used on the current 

metal aileron. To preclude galvanic corrosion, aluminum parts within the 

aileron are anodized, primed with epoxy and then given a urethane topcoat. 

Graphite/epoxy parts in contact with aluminum parts are also painted with a 

urethane topcoat. Faying surface sealant is used at the interface of all 

aluminum and composite parts. After assembly the aileron is primed and 

painted. Lightning strike tests indicated that no lightning protection was 

required for the aileron assembly. 

Thermal expansion differences between the graphite/epoxy aileron and 

the aluminum wing were calculated over the operating temperature range. 

Expansion differences were found to be less than the available lateral play 

at the inboard hinge. However, to maintain the same amount of lateral play 

a slight modification was made to the hinge fitting at the inboard hinge 

location. 

The simplicity of the composite aileron design and the unique combination of materials utilized have resulted in a very efficient structure 

Figure 6. - Aileron assembly. 

8 
which is 22.9 percent L14.6 kg (32.1 lb)] lighter than the metal aileron. 

Since the composite aileron has 50 percent fewer parts and fasteners than the 

metal aileron (see table l), it is predicted to be cost competitive in a 

production environment. 

TABLE 1. - COMPARISON OF COMPOSITE AILERON TO ALUMINUM AILERON 

Weight, kg.@) 

Percent Composite Material 

Weight Saved, kg (lb) 

No. Ribs 

No. Parts 

No. Fasteners 

Aluminum 

Aileron 

63.4 (139.9) 

5.8 

18 

398 

5,253 

Composite 

Aileron 

48.9 (107.8) 

64 

14.6 (32.1) 

10 

205 

2,574 

Analysis 

Internal loads, deflections, and structural influence coefficients for 

the composite aileron were predicted using a three-dimensional NASTRARl 

finite element model. Using this information a stress analysis was made 

of each component to determine ply level strains. This analysis indicated 

that the maximum strain within the aileron occurs in the +45" fabric plies 

of the main rib web at Inboard Aileron Station (I.A.S.) 102.7. The longitudinal (fill direction) ply level strain in this region is 0.0035 tension 

(ultimate). 

A finite element stability analysis was performed on the upper and 

lower covers and front spar web. The combined loads from the NASTRAN 

finite element model ultimate internal loads results were applied. Initial 

buckling occurs for the upper and lower cover at 24 percent and 20 percent 

of design ultimate load, respectively. The spanwfse loading is the major 

contributor to buckling of the covers. 

It is required that the aileron be capable of supporting limit loads 

(considered as ultimate) with a critical element broken or inoperative. Failsafe conditions were analyzed for the basic limit design conditions. Internal 

loads for these conditions were compared to those for the basic ultimate 

loading conditions. All of the fail-safe condition strains were less than 

those for the ultimate design load conditions. 

Vibration analyses of the composite aileron were performed using the 

NASTRAN three-dimensional finite element model. Results of these analyses 

INASTRAN: Registered trademark of the National Aeronautics and Space 

Administration. 

9 
were compared to measured vibration characteristics of the metal aileron 

and it was concluded that the behavior of the composite aileron was similar 

to the metal aileron. 

VERIFICATION OF STRUCTURAL INTEGRITY 

Design Data 

Fundamental material behavior was determined by a large number of 

coupon tests at several environmental conditions. Ply level data for graphite/ 

epoxy tape and fabric were obtained for use in the analytical predictions of 

laminate behavior. Laminates representing the covers, ribs, and front spar 

were tested to verify material behavior and to establish strength adjustment 

factors for environmental conditions, notches, and damage tolerance. In 

addition, specimens representing various mechanical joint areas in the ACA 

design were tested to verify joint strength. 

The environmental extremes for the aileron are temperature [219.3 K 

(-65OF) to 355.4 K (180°F)] and various fluids including jet fuel, hydraulic 

fluid, paint stripper and water. Screening tests were conducted to determine 

which of the above environmental extremes caused the greatest degradation of 

mechanical properties. These conditions were then used for testing to establish design allowables. Examples of coupon data for the ply level properties 

and typical laminate data for Thornel 300/5208 bidirectional fabric 

(24X23 - 8HS) are given in tables 2 and 3. Note that typically the 

219.3 K (-65°F) dry condition results in the greatest reduction in tensile 

strength and that the 355.4 K (180'F) wet (one percent moisture by weight) 

condition is the most severe for compression loading. For the notched 

laminate tests a 4.76 mm (3/16 in.) diameter open hole was used since this 

is the most common fastener size utilized in the aileron assembly. 

Defect and damage tolerance of the aileron was determir.ed by conducting 

tests on representative laminates. Defects and damages investigated included 

voids, notches, and impact damage. Information from these tests was used to 

establish the manufacturing accept/reject criteria and laminate design 

allowables. Typical defect and damage tolerance data for the ACA cover 

laminate and a graphite/epoxy tape laminate are presented in figure 7. 

Results of all of the tests indicated that for tensile loads an open hole 

degraded the strength much more than a void or impact damage. For compression 

loads, impact damage caused the greatest reduction in strength. 

For the graphite/epoxy composite materials used for the aileron, the 

design allowables were established to reflect the worst environmental condition in combination with the largest defect allowed by the accept/reject 

criteria, a 4.76 mm (3/16 in.) diameter open hole, or impact damage which 

cannot be located by visual inspection. All of the laminate data were 

statistically analyzed and design allowables established. Allowables for 

T300/5208 tape and fabric are presented in table 4. 

10 
TABLE 2. - T300/5208 GRAPHITE/EPOXY FABRIC 

CHARACTERIZATION TEST DATA@ 

O” Tensile 

Strength - MPa (ksi) 

Modulus - GPa (msi) 

O” Compressive 

Strength - MPa (ksi) 

Modulus - GPa (msi) 

+45O Tensile 

Strength - MPa (ksi) 

Modulus - GPa (msi) 

O” Short Beam 

Shear Strength - MPa (ksi) 

297 K (75OF) Dry 

Unnotched 

591 (85.77) 

71 (10.29) 

583 (84.62) 

63 ( 9.16) 

203 (29.45) 

20 ( 2.83) 

49 (7.1) 

219.3K (-65OF) Dry 

Unnotched 

487 (70.64) 

65 ( 9.46) 

621 (90.03) 

65 ( 9.42) 

217 (31.52) 

21 ( 3.04) 

37 (5.4) 

@ Data Normalized to a Cured Ply Thickness of 0.36 mm (0.014 in) (61% Fiber Vol.) 

@ 1% Moisture By Weight 

355.4K (18OOF) Wet @ 

Unnotched 

557 (80.73) 

71 (10.30) 

433 (62.80) 

62 I 9.05) 

175 (25.43) 

16 ( 2.29) 

‘43 (6.2) 

Tensile 

Strength - MPa (ksi) 

Modulus - GPa (msi) 

Strain - lo6 

Compressive 

Strength - MPa (ksi) 

Modulus - GPa (msi) 

Strain - 1 O6 

TABLE 3. - T300/5208 GRAPHITE/EPOXY FABRIC 

(45"/0"/135"/0"/45") LAMINATE DATAa 

297K 

(75OF) Dry 

Unnotched 

403 (58.4) 

43 ( 6.2) 

9 430 

497 (72.1) 

40 ( 5.8) 

13,540 

297K 

(75OF) Or 

Notched b 

240 (34.8) 

41 ( 6.0) 

5 790 

318 (46.1) 

41 ( 6.0) 

7 810 

219.3K 

(-650~) Dry 

Unnotched 

383 (55.6) 

43 ( 6.3) 

9 240 

482 (69.9) 

42 ( 6.1) 

12 060 

219.3K 

(-650~) Dr 

Notched d 2 

219 (31.7) 

42 ( 6.1) 

5 110 

365 (52.9) 

43 ( 6.3) 

7 970 

1803r?)~t~ 

Unnotched 

408 (59.2) 

43 ( 6.2) 

10 070 

364 (52.8) 

41 ( 5.9) 

9 990 

:1,Ei4L@ 

Notched @ 

- 

243 (35.3) 

41 ( 5.9) 

5 880 

262 (38.0) 

39 ( 5.7) 

6 950 

@ Data Normalized to a Cured Ply Thickness of 0.36 mm (0.014 in) 

0 Notch = 4.76 mm (3116 in) 

@ 1% Moisture by Weight 

11 
TAPE/SYN (COVER) LAMINATE 

(45”/0”/ - 4WSY N/ - 45”/0”/45”) 

TAPE LAMINATE 

( + 45”/03”/ T 45VO”)s 

TENSION COMPRESStON TENSION COMPRESSION 

Figure 7. - Typical defect and damage tolerance data. 

TABLE 4. - GRAPHITE/EPOXY TAPE AND FABRIC LAMINA DESIGN ALLOWABLES* 

Material 

Load 

Direction O” Property Tape 

Modulus - GPa (msi) 141 (20.5) 

Tension Strength - MPa (ksi) 671 (97.4) 

Strain - 1 O6 4 750 

Compression Modulus - GPa (msi) 128 (18.5) 

Strength - MPa (ksi) 510 (74.0) 

Strain - 1 O6 4 000 

*Worst Environmental Condition, 4.76 mm (3/16 in.). Diameter Notch or Impact Damage. 

Fabric 

67.6 ( 9.8) 

263 (28.2) 

3 900 

60.3 ( 8.7) 

240 (34.8) 

4 000 

12 
Concept Verification 

Structural tests were conducted on ACA subcomponents and assemblies, as 

summarized in figures 8 and 9, to investigate the various design details and 

to verify that strength is adequate. These tests included static tests, and 

spectrum fatigue tests for two lifetimes followed by residual strength tests. 

The cover specimen which was tested included local regions where the 

syntactic core is replaced by strips of graphite tape doublers which form part 

of the rib cap and spar cap structure. After confirming the predicted shear 

stiffness of the cover, these doublers were investigated by loading the cover 

in combined shear and compression (and later, tension) to 112 percent of design 

ultimate load without failure. Subsequently, the fail-safe capability was 

demonstrated by cutting a 305 mm (12 in.) long slot in the most highly loaded 

region of the cover (adjacent to the rib/spar intersection). Limit load (shear 

and spanwise tension) was applied, but no damage growth occurred. The load was 

increased to 169 percent of limit load with no damage growth or failure. This 

confirmed the damage tolerance/fail-safe characteristics of the cover design. 

COVER RIB 

1 STATIC 

5 LIGHTNING 1 STAT 

STRIKE 

FRONT SPAR 

IC 

Figure 8. - Summary of subcomponent structural tests. 

Intermediate and closure rib webs contain flanged lightning holes which 

result in a 13 percent weight savings compared to a solid web rib. Analytical 

investigations indicated that the loss of shear strength in the web due to a 

cutout could be recovered by adding a flange to the periphery of the hole. 

Experimental verification of the shear strength of a web containing a flanged 

lightning hole was obtained by statically testing a segment of an intermediate rib. The specimen, shown installed in the picture frame test fixture 

on figure 10, failed at a load of 2168 kg (4780 lb), which is approximately 

three times greater than design ultimate load (D.U.L.). Initial buckling of 

13 
RIB/SPAR/FTG BOX ASSEMBLY 

2 STATIC 

1 FATIGUE 

FATIGUE 

GROUND TES 

ARTICLES 1 STATIC 

1 DAMAGE GROWTH/ 

FAILSAFE 

Figure 9. - Summary of assembly structural tests. 

the specimen occurred at 771 kg (1700 lb). Both the initial buckling load 

and failure load of the specimen were analytically predicted to within six 

percent of the measured loads. 

The front spar together with mechanically fastened segments of the 

upper and lower covers and the hinge/actuator fitting was tested both 

statically and in fatigue. A download of 123 percent D.U.L. was applied 

to the hinge/actuator fitting (with flexures to react loading eccentricities) 

without failure. The additional 23 percent over ultimate load was intended 

to account for strength differences between the environmental extremes and 

the ambient testing conditions. The spar was then subjected to two lifetimes of fatigue loading (the equivalent of 72 000 flights). There was no 

evidence of damage or degradation during or after the fatigue test. A 

static upload was then applied until failure occurred at 182 percent D.U.L. 

Failure occurred by interfastener buckling of the lower cap flange and 

cover and buckling of the web. Failure had been predicted at 169 percent 

D.U.L. in the spar web. These tests verified the structural integrity of 

the front spar design concept. 

The hinge/actuator fitting at the main rib introduces a concentrated 

load of over 12 250 kg (27 000 lb) into the rib and covers. The rib/spar/ 

fitting specimen was tested statically and in fatigue to verify the design 

concept. As shown in figure 11, a fixture introduced the loads into the 

14 
Figure 10. - Setup for rib static test. 

hinge/actuator fitting, and the loads were reacted through the covers and 

the rib web. The specimen was loaded in compression to 110 percent D.U.L. 

without visible damage. Two lifetimes of fatigue loading were applied 

to a second specimen without visible damage. Then it was loaded statically 

in tension, and a bolt attaching the fitting to the aileron failed at 

140 percent D.U.L. which substantiated the analysis which had predicted 

failure at 139 percent D.U.L. These tests verified the structural integrity 

of the fitting load transfer design concept. 

To verify resistance to sonic fatigue, a full scale section of the composite aileron was mounted in an acoustic progressive wave tunnel and tested 

at an accelerated spectrum level 5 dB above the design noise environment of 

135 dB/Hz for 10 hours. This conservatively results in the equivalent of one 

lifetime, including environmental effects. The specimen was instrumented to 

measure critical strains. The highest value measured was 310 microstrain 

root mean square (rms> which gives a substantial margin against the allowable 

15 
Figure 11. - Rib/spar/fitting specimen installed in test machine. 

level of 600 microstrain rms. Following the test there was no evidence of 

damage or degradation. 

The aileron is in the zone 2 swept stroke region of the wing. To 

determine the effects of lightning a full scale section of the aileron was 

subjected to four swept stroke-restrikes of 100 000 amperes and two direct 

strikes of 200 000 amperes. The box was then inspected ultrasonically to 

map the damage size. The worst swept stroke-restrike damage occurred between 

ribs resulting in a 6.35 mmAO.25 in.) diameter penetration, burning, and a 

16 
delamination region of 380 mm (15 in.) by 76.2 mm (3 in.). The damage was 

within the limits of the fail-safe criteria and repairable. 

Ground Tests 

The first set of ground tests consisted of vibration and stiffness 

tests for both a metal and composite aileron to enable a relative comparison. 

The aileron was mounted by the hinges in a reaction test frame, and various 

actuator configurations were simulated with adjustable links. An electrodynamic shaker, attached by a bonded pad to the outboard closeout rib, was 

used to apply various frequencies and amplitudes to the aileron. A reference 

accelerometer together with a roving accelerometer were used to monitor the 

frequency responses summarized on table 5 below. Since the frequencies were 

so comparable, flutter-free performance of the composite aileron was 

predicted. 

A pair of actuators attached by pads to the rear spar were used to load 

the aileron to determine chordwise bending and torsional stiffness. The 

rotation and deflection at nine points on the aileron surface were measured 

both optically and with linear transducers. The average chordwise bending 

stiffness of the composite aileron was 27 percent less than that of the 

metal aileron. 

Torsional stiffness testing produced some unexpected results. The 

composite aileron stiffness was multilinear and generally increased as the 

load increased. Furthermore, the stiffness increased during each of three, 

or four, load applications; finally stabilizing after four, or five, load 

applications. To test the hypothesis that this phenomenon might be due to 

the clearance fit of the fasteners used for assembly the load was reversed. 

That is, the load was applied at station 57.1 and reacted at station 102.7 

instead of vice versa. The initial loading resulted in 20 percent less 

stiffness than in the other direction. However, after four more load applications the stiffness stabilized at 96 percent of the stiffness in the 

other direction. To further test the fastener fit hypothesis, the second 

ground test article was tested in the same manner. It was known that the 

fastener fit for the second article was better than for the first article 

due to improved hole drilling methods. The fastener fit hypothesis was 

TABLE 5. - FREQUENCY COMPARISON 

Mode ,,I Actuator Condition 1 Metal Freqrncv Composite 

Flapping 

Flapping 

Flapping 

Torsion 

17 
confirmed by the fact that its stiffness was higher, had little variation 

with load magnitude, and stabilized after only three load cycles. The 

torsional stiffnesses of the three ailerons tested are summarized in 

table 6. 

Following the stiffness tests , pads were bonded to the upper surface 

over ribs and spars to introduce the loads from four hydraulic cylinders 

through a whiffle tree system as shown on figure 12. The four hydraulic 

cylinders were computer-controlled through a hydraulic servo system to 

apply the correct proportion of loading throughout the load range. The 

aileron actuators were simulated with three instrumented links that were 

adjustable to simulate the angular position of the aileron. Strains were 

monitored with 27 axial or rosette strain gages. Deflections were monitored 

with eight linear transducers. 

Two loading conditions were tested. Condition 4 loads, representing 

the 12" down aileron position, were applied to 124 percent D.U.L. without 

failure. The 24 percent factor on condition 4 loads resulted in loads on 

the hinge/actuator fitting fasteners equal to the design ultimate load for 

another loading condition that was not tested. There was no visible damage 

from this test. 

The condition 1 loads, representing the 20" up aileron position, were 

applied to 117 percent D.U.L. without failure. The 17 percent factor 

accounted for the strength differences between the environmental extremes 

and the ambient testing conditions. The condition 1 loads were increased 

to failure which occurred at 139 percent D.U.L. Failure was a result of 

postbuckling of the webs of the hinge and actuator backup ribs at stations 

102.7 and 107.1. 

A second ground test article was tested to verify the damage tolerance 

and fail-safe characteristics of the composite aileron. Four critical 

locations were selected for inflicting damage: the upper cover was impacted 

to obtain visible damage, the main rib cap was partially severed at a 

fastener, the front spar web was impacted to obtain visible damage, and the 

TABLE 6. - TORSIONAL STIFFNESS COMPARISONS 

Torsional Stiffness 

IO3 N-m2 (IO6 lb-in*) 

Required 

Predicted 

Metal 

861 (300) 

1068 (373) 

Measured I 1105 (385) 

GTA I GTA 2 

861 (300) 

1059 (369) 

1079 (376) 

18 
Figure 12. - Aileron ground test set-up. 

forward spar cap flange of the lower cover was impacted to obtain visible 

damage. The design requirement was that the damage not grow to catastrophic 

size between normal major inspection periods of approximately 20 000 hours. 

To demonstrate this, the aileron was subjected to one lifetime o# flightby-flight fatigue loading. Inspection of the aileron after fatigue loading 

indicated only slight growth in the damage to the upper cover. Limit load 

with a 17 percent environmental factor for both condition 4 and condition 1 

was sustained demonstrating that damage had not reached critical size. 

The aileron was subjected to fail-safe tests to demonstrate that it 

could complete its flight after suffering substantial damage from an obvious 

discrete source. The damage selected was that resulting from sweptstroke 

lightning. The damage to the full scale section of-the aileron during the 

lightning test was simulated on the second ground test article by impacting 

the cover and burning it with a welder's torch. This resulted in a burnthrough and a delamination over an area 381 mm (15 in.) by 76.2 mm (3 in.). 

To demonstrate the ability to complete the flight a load equal to 70 percent 

of limit load (the maximum reasonable expected during completion of the flight) 

with a 17 percent environmental factor was applied in both the condition 4 

and condition 1 configurations. There was neither failure nor evidence of 

damage growth. 

19 

I 
Following all these tests, the aileron was loaded to failure to 

determine its residual static strength. Failure occurred at 130 percent 

of design ultimate load in essentially the same mode as occurred with the 

undamaged static test article. 

The successful conclusion of the static tests and the damage tolerance 

and fail-safe tests demonstrated the structural integrity of the composite 

aileron. 

Flight Test 

The flight test of the advanced composite inboard ailerons was successfully completed in June 1980. The first shipset (left and right) of inboard 

composite ailerons was installed on the Lockheed L-1011 flight test aircraft. 

The aircraft takeoff gross weight (TOGW) was 160 121 kg (353,000 lb). -The 

ailerons were ballasted with lead tape to achieve a worst-case hinge unbalance 

of 23 043 mm-kg (2000 in-lb): Ground servo stability during engine run up, 

level flight, and high-speed descent tests were performed (figure 13). During 

these tests one or two of the three hydraulic servos were shut off to simulate 

fail-safe conditions. Pulses were applied to the control column to give a 

momentary longitudinal pitch or lateral right or left roll impulses. 

M FT 

15,000 - 

ALTITUDE 

20,000 

5,000 - 

10,000 

M = .95 DIVE 

KNOTS - EQUIVALENT AIRSPEED 

Figure 13. - Flight conditions investigated. 
The advanced composite aileron response was somewhat better than the 

metal aileron during ground engine run up, and the damping characteristics 

of the composite aileron were comparable to the metal aileron during all 

tests. The composite aileron was flutter-free throughout the flight 

envelope. 

AILERON FABRICATION 

Tooling and processing parameters were developed for the composite 

aileron. These procedures were used to manufacture five shipsets of these 

ailerons in a production environment. A flow diagram defining the basic 

manufacturing operations necessary to perform this task is shown in 

figure 14. 

TOOLING 

Prior to the initiation of the production composite detail fabrication 

effort an evaluation was made of male vs female tooling philosophy for 

producing channel configuration details such as spars and ribs. The effectiveness of each tooling concept was evaluated for tool cost (recurring and 

non-recurring), detail dimensional repeatability, cost, and physical and 

mechanical property consistency. The male tool concept was chosen because 

less tool development was required, lower tool fabrication and maintenance 

cost, lower composite fabrication costs and the necessary dimensional 

repeatability could be maintained on fabricated details. 

PROCESS DEVELOPMENT 

The processes developed for fabrication of the graphite parts utilized 

the following significant achievements: 1) a common cure cycle for all 

details, 2) use of low resin content material, 3) incorporation of a no 

trim technique of cure. In addition, the assembly techniques developed 

showed that composite details can be drilled with hand tools at a production 

rate and maintain the high quality standards required. 

DETAIL FABRICATION 

The 13 graphite composite parts required for each composite aileron 

were fabricated in production environment using the developed tools and 

processes. 

21 
COMPOSITE 

MATERIALS 

BOND 

BOND SHOP SHOP Q.A. 

LAYUP tXJF?E TRIM 

I I 

INTERMEDIATE --- ASSY SHOP 

RIBS 

Q.A. 

BATCH 

ACCEPTANCE --- 

FRONT SPARS- PREPLY TRIM LAYUP CURE 

L 4 

ASSY 

-a- SHOP 

MAIN RIBS A RIB SUB 

ASSEMBLY 

4 

1 ASSY SHOP 

MACHINE SHOP 

FITTINGS 

METAL COMPONENTS 

REAR SPAR FAB 

CLIPS FAB 

PURCHASED PARTS I 

Figure 14. - Manufacturing flow .diagram. 
Ribs 

The three types of ribs were fabricated in a like manner. The required 

quantity of graphite/epoxy fabric plies were preplied and then hand trimmed 

to net size using a template and shop knife. This preplied material was 

laid up onto the male cure tool. Bleeder/breather materials were then 

applied to the layup. Bleeder/breather arrangement for the ribs was as 

follows: one peel ply, one ply Mochburg CW 1850 paper, and Armalon strips 

to provide a vacuum path to the vacuum source. A silicone rubber bag was 

applied, which acts as a dam at the part edge for control of resin flow and 

allows a more even distribution of pressure on the laminate. This entire 

layup with a product control panel was sealed under a nylon vacuum bag. 

This arrangement is shown in figure 15. The ribs were cured to a net 

configuration, eliminating subsequent trimming operations. Resin flash 

was manually removed using 200 grit abrasive paper. 

MOCHBURG 

CW 1850 

AND 

PEEL PLY, / 

SILICONE RUBBER BAG 

AIRWEAVE “A” FR (2 PILES) 

/ 

I I 

GLASS BAG 

FABRIC SEALER 

STRIP 

Figure 15. - Rib layup 

23 
Spar 

The ten-ply graphite/epoxy tape spar laminate was preplied in four sets 

as follows: three plies/two plies/two plies/three plies. Each set of preplied material was hand trimmed to net size using a template and shop knife 

as shown in figure 16. The spar cure tool (see figure 17) was heated to 

325 K (125°F) to enhance drapability of the graphite tape. The four sets of 

preplied material were oriented on the cure tool by aligning the prepreg 

index holes on the tool targets during layup. Bleeder/breather materials, 

a silicone rubber bag and nylon vacuum bag were then applied. A process 

control panel was included and laid up on the same tool. The spar was alsocured to a net configuration eliminating any trimming operations. 

Figure 16. - Spar preplying operation. 

24 
r 

Covers 

The covers were fabricated by first preplying the graphite/epoxy tape 

material required for the two face sheets and the doublers. This preplied 

material and the syntactic core was then hand trimmed using templates and a 

shop knife. The trimmed material was laid up on the flat cure plate starting 

with the outer surface plies. The syntactic core and inner doublers were 

then laid up, followed by the three inner surface plies and the inner 

surface doublers. Tooling holes in tabs were used to index the laminates to 

targets on the cure plate. Bleeder/breather materials were applied to the 

layup as follows: tool side of laminate - one ply of Armalon; bag side of 

laminate - one ply each of Armalon, A4000P3, Armalon, A4000P4 and Airweave 

SS FR. A nylon vacuum bag was then applied. The product control panel for 

the cover was cured integral to the part. After cure, the cover was 

mounted in a router fixture and the excess was removed by hand routing 

using a diamond bit. 

ASSEMBLY 

The aileron structure requires three assembly operations: main rib 

subassemblies; front spar subassembly; and final assembly. To assure the 

required structural integrity of the composite aileron assembly, high 

quality close tolerance holes are mandatory for the mechanical fasteners. 

The drilling system developed to maintain these standards utilizes a hand 

held drill motor equipped with: 1) a hydraulic cylinder to control feed 

rate, 2) a bushinged foot to stabilize drill bit and drill angle, 3) a 

coolant system for cooling and flushing the graphite, 4) a vacuum system 

to remove shavings and excess coolant, and 5) a dagger type drill which 

actually shaves the hole to the close tolerance requirements. This system, 

shown in figure 18, has produced excellent quality holes without the use 

of a drill back up. 

Main Rib Subassembly 

There are three different main rib subassemblies, each consisting of 

one main rib and four aluminum backup fittings per subassembly. A bench 

type fixture was loaded with these components and all attachment holes 

were drilled. After drilling, the fittings were deburred, faying sealants 

were applied, and fasteners were wet installed. 

Spar Subassembly 

The main and intermediate ribs were located and clamped into the spar 

assembly fixture. The spar was then loaded. The hinge and actuator fittings, 

feedback fittings and shroud ribs were positioned. All holes were then 

drilled full size. Metal parts were deburred. Faying surface sealants 

were applied and all parts assembled with the required mechanical fasteners. 

25 
Figure 17. - Spar preply template and cure tool. 

Final Assembly 

The final assembly of the aileron consists of the spar subassembly, 

rear spar, upper and lower covers, trailing edge, end fairings and leading 

edge shroud. The assembly sequence was as follows: 

1. Load front spar assembly, rear spar and closeout ribs 

2. Drill attach holes through rear spar and ribs 

3. Load upper cover and drill attach holes 

4. Remove upper cover, load lower cover and drill attach holes 

5. Remove lower cover and permanently fasten upper cover 

26 
Figure 18. - Drill motion with hydraulic check attachment. 

6. Install lower cover 

7. Remove assembly from fixture and place in handling fixture 

8. Locate and drill leading edge shroud panels and doors, trailing 

edge and end fairings. Installation of these parts then follows. 

QUALITY ASSURANCE 

Prior to utilization of any of the composite raw materials, batch 

acceptance testing was conducted as required by the material specifications. 

These tests include visual, mechanical and chemical tests that must be conducted on each batch of material prior to production use. 

The quality assurance requirements for both fabrication of composite 

parts and assembly of the parts were established and controlled by process 

bulletins. These specifications define the process controls on material 

and labor, as well as nondestructive inspection and process control coupon 

test requirements for each detail. 

27 
Each part had thickness measurement determinations and was ultrasonically 

inspected for voids. The process control panel for each part was cut into 

specimens and tested for resin content, void content and short beam shear 

properties. 

Manufacturing routing sheets were used throughout the fabrication and 

assembly areas to provide work instructions to the latest engineering and 

specification requirements. The inspectors verify and document these operations. Detail fabrication documentation includes: material batch acceptance 

logs, ply orientation and sequence, cure records, process control test 

results, ultrasonic inspection, and visual and dimensional inspection. 

During assembly operations, the quality assurance inspectors check for 

part numbers, detail acceptance stamps, proper location and fitup as the 

assembly fixtures are loaded. After drilling and removal from the fixture, 

each hole was inspected for size, tolerance and quality. Any condition that 

could not be reworked in accordance with engineering requirements is placed 

on a Nonconforming Material Review (NMR) tag for disposition. In addition 

to the typical quality surveillance, the final assembiy of each aileron 

received FAA conformity inspection , which included supporting -data such 

as material acceptance logs, test results of each composite detail, weights 

and copies of all processed NMR tags. 

MANUFACTURING COST 

The manufacturing costs associated with the advanced composite aileron 

have been developed through three phases. The first phase, prior to fabrication consisted of a cost estimate of producing both the composite and 

metal aileron configuration based on previous experience. The second 

phase involved the tracking and documentation of the actual costs to manufacture components of 12 composite ailerons. These data were used to 

determine the current cost of manufacturing composite ailerons. During 

the third phase, actual production data developed during phase two was 

used to establish a cost estimate for producing advanced composite ailerons. 

Further analysis compares the projected composite costs, assuming automated 

equipment and material costs currently being quoted, with the metal aileron. 

During the fabrication and assembly of two ground test articles (GTA's) 

and ten production ailerons, labor and material costs were tracked and 

compiled for the front spar, a main rib, anlntermediate rib, the upper 

cover, and assembly. From these documented data the estimated cost to 

produce an average GTA unit, the average estimated cost to produce an 

aileron, a projected improvement slope (learning curve), and the projected 

cost to produce 100 ailerons (using similar tooling, equipment and 

facilities) were computed. 

28 
A cost analysis, using the fabrication and assembly data developed on 

the composite ailerons, was conducted to compare the cost of composite 

aileron with the aluminum aileron. 

Producibility Cost Analysis 

An analysis was made to compare the costs of manufacturing the 

composite aileron and the sheet metal configuration. The composite 

aileron cost is based on the manufacturing data which were accumulated 

and documented. The corresponding costs for fabricating the sheet metal 

configuration were determined using a standard cost estimating methodology. 

The comparison, shown in table 7, was based on the following premises: 

l Costs are stated in 1980 dollars 

l Costs include recurring materials, production labor and quality 

assurance only 

l Labor rates and material costs per AVCO study 

l Labor calculated as a cumulative average of 100 units using the 

learning slopes experienced during fabricating and assembling 

advanced composite and sheet metal components (see figure 19). 

l The recurring production materials and their corresponding costs are: 

Graphite/Epoxy Tape $123/Q ($ 56/lb) 

Fabric $lSO/kg ($ 68/lb) 

Syntactic $ 81/m2 ($7.50/ft2) 

This comparison demonstrates that the cost of the composite material 

is the major cost driver. The quantity of fasteners was reduced from over 

5000 to less than 3000. However, in most instances, because of the need 

for material compatibility, the low-cost fasteners used in aluminum 

structure had to be replaced with more expensive titanium fasteners. 

In summary, the costs saved by eliminating fasteners and the corresponding 

reduced installation manhours were not sufficient to overcome the higher 

material and fastener costs. This situation will be partially remedied 

in the future through increased quantity procurement. 

A comparison of the labor costs, including inspection, of the redesigned 

aileron components, shown as percentage of the total, are given in table 8. 

These data show that the major cost driver in labor is installing the 

fasteners during final assembly. 

29 
TABLE 7. - ALUMINUM AILERON COST COMPARED TO COMPOSITE AILERON COST 

BASED ON CURRENT MATERIALS COST 

Component 

Covers 1 799 

Ribs (Main) 698 

Ribs (Close Out) 235 

Ribs (Intermediate) 911 

Forward Beam 756 

Aft Beam 44 

F.B. Fittings* 1 966 

Shroud/Fairing* 2 647 

T.E. Wedge* 3 922 

Assembly 2 541 

Fasteners 500 

Paint 33 

Mfg. Variance 0 

Total 16 052 

*Supplied by Short Bros. 

Sheet Metal 

Configuration 

cost - $ 

- 

Adv. Composite 

Configuration 

cost - $ 

Amount 

of Change 

cost - $ 

3.958 +2 159 

1 597 + 899 

633 + 398 

1 560 + 649 

701 - 55 

44 0 

1 966 0 

2 647 0 

3 922 0 

1 850 - 691 

3 127 +2 627 

33 0 

0 0 

22’038 +5.986 

TABLE 8. - LABOR COST COMPARISON 

Sheet Metal Fabrication 

Composite Fabrication 

Sub Assembly 

Final Assembly 

Quality Control 

Metal Configuration 

32% 

21% 

35% 

12% 

100% 

Composite Configuration 

4% 

23% 

54% 

19% 

100% 

30 
2500 

2000 

z 

z 

is 

1000 

3 

G 500 

E 

E 

200 

LEGEND: 

2187 HRS CUM AVG ATOTAL LAB01 

2 GTA UNITS Cl ASSY. LABOR 

1439 HRS CUM AVG 

859 HRS CUM AVG 

2 GTA UNITS 

873 HRS CUM AVG 

PROJECTED AT 1DOTH UNIT 

566 HRS CUM AVG 

10 PROD UNITS 

343 HRS CUM AVG 

PROJECTED AT 1OOTH UNIT 

1 2 5 10 20 50 100 200 

UNITS 

Figure 19. - Actual learning curve development. 

Projected Production Cost 

The projected cost of the composite aileron was calculated based on 

the accumulated cost data. For this analysis it was assumed that the 

aileron would be one of many composite structures being fabricated. 

The following assumptions were made: 

l Buy preplied materials and eliminate many manhours of hand 

layup time. 

l Buy materials in sufficient quantities to reduce the cost 

to $86/kg ($39/lb). 

l Automate the cutting methods to reduce the averaged manhours 

during fabrication by 30 percent. 

l Standardize and procure the mechanical fasteners in larger 

quantities. 

l Automate the assembly operations to reduced manhours during 

assembly by 30 percent. 

31 
A comparison of the projected cost of the composite aileron and the 

metal aileron is shown in table 9. This comparison assumes both the metal 

and composite ailerons are at the same point on the learning curve. 

TABLE 9. - PROJECTED COST COMPARISON 

Component 

Sheet Metal 

Configuration 

cost - $ 

Covers 1 799 

Ribs (Main) 698 

Ribs (Close Out) 235 

Ribs (Intermediate) 911 

Forward Beam 756 

Aft Beam 44 

F.B. Fittings 1 966 

Shroud/Fairing 2 647 

T. E. Wedge 3 922 

Assembly 2 541 

Fasteners 500 

Paint 33 

Mfg. Variance 0 

Total 16 052 

Adv. Composite 

Configuration 

cost - $ 

Amount 

of Change 

cost - $ 

2 378 -I- 579 

985 + 287 

212 - 23 

463 - 448 

427 - 329 

44 0 

1966 0 

2 647 0 

3 922 0 

1 276 - 1 265 

1 866 +l 366 

33 0 

0 0 

16 219 + 167 

CONCLUSIONS 

Results obtained in this program indicate that significant improvements 

in structural efficiency can be achieved by the utilization of advanced composites for construction of aircraft secondary structures. Careful evaluation of alternate designs and materials for the L-1011 advanced composite 

inboard aileron has led to the selection of several unique material combinations and easily manufactured structural configurations. 

The advanced composite aileron is a direct replacement for the metal 

aileron with a weight savings of 23 percent. Due to the configurational 

simplicity of the components within the composite aileron, and because it 

contains 50 percent fewer parts and fasteners than the metal aileron, it 

is predicted that the composite aileron will be cost competitive with the 

metal aileron in a production environment. 

32 
Structural analysis of the composite aileron, in conjunction with the 

design data, concept verification, and ground tests, 'indicates that the 

composite aileron design meets or exceeds structural requirements. 

The simple and practical manufacturing techniques developed for the 

composite aileron have been verified by the production of high quality 

parts. Reliability of the manufacturing process was demonstrated through 

the production of five shipsets of ailerons. Based on the manufacturing 

cost information accumulated during the production of 12 ailerons it is 

predicted that the composite aileron can be produced at a cost competitive 

with the aluminum aileron. 

33 
. . . . . . . . .---.. ,.-. ~~... 

REFERENCES 

1. Griffin, C. F., Fogg, L. D., Stone, R. L., and Dunning, E. G.: Advanced 

Composite Aileron for L-1011 Transport Aircraft - Task I, Final. NASA 

CR-145370, July 1978. 

2. Griffin, C. F.: Advanced Composite Aileron for L-1011 Transport 

Aircraft - Ground Tests and Flight Evaluation. NASA CR-165664, 

February 1981. 

3. Griffin, C. F., Fogg, L. D., Dunning, E. G.: Advanced Composite Aileron 

for L-1011 Transport Aircraft - Design and Analysis. NASA CR-165635, 

April 1981. 

4. Dunning, E. G., Cobbs, W. L., Legg, R. L.: Advanced Composite Aileron 

for L-1011 Transport Aircraft - Aileron Manufacture. NASA CR-165718, 

June 1981. 

34 
1. Report No. 2. Government Accession No. 

NASA CR-3517 

-.-. _--. -: 

4. Title and Subtitle 

DEVELOPMENT OF AN ADVANCED 

COMPOSITE AILERON FOR THE L-101 1 

TRANSPORT AIRCRAFT 

7. Author(s) 

C. F. GRlFFlN”AND E. G. DUNNING** 

9. Performing Organization Name and Address 

LOCKHEED-CALIFORNIA COMPANY ** ;‘;;;;R2F;zRATloN 

P. 0. BOX 551 

BURBANK, CALIFORNIA 91520 AND N’ASHVI LLE, TENNESSEE 

37202 

2. Sponsoring Agency Name and Address 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 

WASHINGTON, DC 20546 

.- .- 

3. Recipient’s Catalog No. 

6. Performing Organization Code 

76-23 

6. Performing Oqanization Report No. 

LR 29911 

10. Wcwk Unit No. 

11. Contract or Grant No. 

NAS l-15069 

5. Supplementary Notes 

14. Sponsoring Agency Code 

LANGLEY TECHNICAL MONITOR: HERBERT A. LEYBOLD 

FINAL REPORT 

j. Abstract 

Results obtained in this program indicate that significant improvements in structural efficiency 

can be achieved by the utilization of advanced composites for construction of aircraft secondary structures. 

Careful evaluation of alternate designs and materials for the L-101 1 advanced composite inboard ailerqn 

has led to the selection of several unique material combinations and easily manufactured structural 

configurations. 

The advanced composite aileron is a direct replacement for the metal aileron with a weight savings of 

23 percent. Due to the configurational simplicity of the components within the composite aileron, and 

because it contains 50 percent fewer parts and fasteners than the metal aileron, it js predicted that the 

composite aileron will be cost competitive with the metal aileron in a production environment. 

Structural analysis of the composite aileron, in conjunction with the design data, concept verification, 

and ground tests, indicates that the composite aileron design meets or exceeds structural requirements. 

The simple and practical manufacturing techniques developed for the composite aileron have been 

verified by the production of high quality parts. Reliability of the manufacturing process was demonstrated 

through the production of five shipsets of ailerons. Based on the manufacturing cost information 

accumulated during the production of 12 ailerons it is predicted that the composite aileron can be produced 

at a cost competitive with the aluminum aileron. 

1. Key Words lSuggested by Author(s)) 16. Distribution Statement 

COMPOSITES, MATERIALS, AILERON 

SECONDARY STRUCTURE DESIGN, 

TRANSPORT AIRCRAFT, GRAPHITE/EPOXY 

FEDD DISTRIBUTION 

1. Security Classif. (of this report) 

UNCLASSIFIED 

1 

20. Security Classif. (of this pe@ 

UNCLASSIFIED 

21. No. of Pages 

41 

22. Rice 

Subject Category 24 

Available: NASA's Industrial Applications Centers 

NASA-Lang1 ey I 1962