Composite issues in rotor hub structure design

Composires Engineering, Printed in Great Britain.

Vol. 2, Nos 5-7, pp. 321-328,

1992.

0961-9526/92 $5.00+ .OO Q 1992 Pergamon Press Ltd

COMPOSITE ISSUES IN ROTOR STRUCTURE DESIGN

HUB

W. S. CHAN Center for Composite Materials, Department of Mechanical Engineering, University of Texas at Arlington, Arlington, TX 76019, U.S.A. (Received

27 February

1992; accepted

27 February

1992)

Abstract-While the use of composites in rotor hubs has produced many benefits, it has also revealed weaknesses in application. In order to design a structurally efficient composite hub, weakness issues in structures and materials need to be addressed. This paper discusses the structural and material aspects of these issues. 1. INTRODUCTION

The technology of composites is extremely prolific, accordingly there is large numbers of such materials available for potential utilization as members in aircraft/aerospace structures. The desirable properties of structural composites such as high stiffness, low density, high resistance to environmental attack, favorable life cycle cost, design flexibility, etc., make composite materials ideally suited to certain structural applications. In the past three decades, composite rotor hubs appear to have offered improvements over conventional metallic rotor hubs in cost, weight, damage tolerance, radar detectability, and maintainability. The real thrust of using composites to replace metallic rotor hubs is to eliminate the flapping and the pitch change bearings, thus creating a fail-safe design. While the use of composites in hubs has produced many benefits, it has also revealed weaknesses in application. Thus, the application of composites materials to rotor structures becomes a most challenging task because of the design requirements for functionality as well as design soundness. Composite technologies in rotor structure have been described by Mayerjak (1978), Olster (1980) and Phillips and Merkley (1990). In the transition from metal to composite, the key point was to fully exploit the beneficial properties of composites in unique designs, not to merely replace those metallic structural rotor hub components that have isotropic loading and support the usual bearings and hinges of conventional hub designs. The objective of this paper is to address both the structural and material aspects that need to be considered in designing a structurally efficient composite hub. Since current hubs use polymeric matrix composites, discussion will be limited to such composites. 2. STRUCTURE

ISSUES

Composite rotors hubs are often made of laminates which are narrow compared to their thickness. These laminates must not only carry the axial centrifugal force but also withstand the beam-wise and chord-wise bending moments. Further, the thickness of structural laminates is being tailored to meet flapping-flexibility requirements. These structural configurations inherently have a high interlaminar stress gradient, which can cause delamination failure. Delamination is a unique issue in the use of composite laminates since composite materials have a low interlaminar allowable (stress or toughness). The following discussion focuses on the issues that affect structural performance. 2.1. Edge effect The free-edge effect in multilayered laminated composites exhibits a unique behavior that has not been observed in metallic structures. Edge delamination has long been considered a subject of academic research, and is well covered in the literature. However, 321

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322

s. CHAN

it is often ignored during design. The primary reason for this unusual indifference to a critical stress is that most laminates are very wide, and the edge effect does not affect the net stress significantly. Hub structures, however, are not wide and the edge effect should not be overlooked. Figure 1 show the edge failure mode of a composite hub that experienced high loads during experimental flight testing. The response at the free edge has been extensively reviewed by Pagan0 (1990). The associated failure is a delamination failure resulting from high interlaminar stresses at the edge. The presence of delamination causes a reduction in stiffness and a degradation of the compoiste materials strength. Chan et al. (1991) have developed an analytical method to analyze and predict the edge delamination of composite rotor structures. Although this failure may not immediately cause catastrophic failure of the hub, it often hastens repair or replacement which inhibits aircraft readiness and results in increased costs (O’Brien, 1990). Therefore, controlling delamination initiation and growth is clearly the key issue to improved damage tolerance and durability of composite hubs (Chan et al., 1986). Chan (1991) reviewed the material and structural approaches to increasing the resistance to edge delamination initiation and growth. He concluded that currently available toughened thermosets and thermoplastic composites can significantly improve delamination resistance for static, but not fatigue, loading. He also concluded that selective interleaving of a soft high-strain layer and terminating the critical ply in the structural edge are the most effective ways of increasing edge delamination resistance. 2.2. Taper effect Tapering in composite rotor hubs is achieved by internally dropping plies to increase the flapping capability. However, this creates geometric and material discontinuities that cause an eccentric load path, thus inducing interlaminar stress and initiating delamination. In the past few years, research effort have been focused on investigating the stress distribution in the vicinity of tapering (Curry et al., 1987) and the delamination characterists (Fish and Lee, 1988; Salpekar et al., 1991; King and Chan, 1991). The dependence of the interlaminar normal stress on the change in taper angle as well as on the ply orientation was investigated by Chan and King (1991). Figure 2 shows the dependence of the interlaminar normal stress at the interface between 0” and neighboring plies on to the change in taper angle for three of locations of the taper beam. It was found that the normal stress is more sensitive at the taper root (x = 0) for a given change in taper angle. This suggests that reducing the taper angle would increase the strength of the tapered beam. However, a longer taper increases the hinge offset which increases the mast and transmission loads, but reduces the overall structural efficiency. 2.3. Structural

corner

The intersection of the four arms of an in-plane four-bladed rotor hub presents a unique design challenge. A large corner radius reduces the stress concentration and is beneficial in metal structures, but is actually detrimental in composite structures. Figure 3 shows a composite rotor system manufactured by Bell Helicopter Textron. This hub was subjected to about 275% of the maximum level flight oscillatory loads. Delamination at the corners of adjacent arms of the hub was observed during testing. At the conclusion of the test, no significant loss in stiffness could be determined, demonstrating the fail-safe feature of this design. In this hub, the structural corners were formed by a large number of ply terminations. An analysis was conducted to better understand and improve the delamination characteristics of the hub. Figure 4 presents a comparison of finite element results for the hub with an without a corner under design load. As shown, at the area of highest ply stress, the transverse stress of the 0” ply in the material coordinate, 02, changes from tension in the cornered structure to compression in the uncornered structure. Furthermore, the interlaminar normal stress, rrZ, is also reduced. The change of o2 from tension to compression may delay or prevent ply splitting, and the reduction of oZ prolongs delamination life. The rationale for eliminating or reducing the corner radius is to provide a direct load path and reduce the load

Rotor hub structure design

Fig. 1. Failure mode of a composite rotor hub.

323

324

w.

s.

CHAN

CROSS SECTION VIEW OF LAMINATES PLY TERMINATION (30.

-302,30,

go), AS4/3SOl-6

Fig. 7. Photomicrograph

(+35,

WITH

0, 90)sAS4/3501-6

of a cross-section of a tapered beam.

Rotor hub structure design

1 -S

I -3

1 -1

325

I 1

I 3

1 5

change in taper angle (degree) Fig. 2. Dependence of normalized interlaminar normal stress at the interface between 0” and neighboring plies on taper angle variation.

YOKE ASSEMBLY LEAD-LAG

UPPER CLAMP

DAMP

PLA

CUFF ASSEMBLY

SHEAR

RESTRAIN

LEAD-LAG

DAMPER

f.r=

Fig. 3. Bell Model 680 advanced bearingless main rotor.

HIGHEST

WITH

2

rY

0

l 4s

x

FREE EDGE

PLY

CORNER

STRESSES

W/O

CORNER A

q

A 27.4

B -

712

27.8

9

2.1

-

-0.3

1.8

-

0.6

0,

23.8

30.2

15.5

9

7.1

8.8

5.0

7,2

4.0

4.8

5.0

9

-

1.71

1.23

Fig. 4. Comparison of finite element results of a composite hub with and without a structure corner.

326

W.S.

&AN

COMBINED IN CORNERS

CURRENT REDUCE CORNER RADIUS TO PROVIDE MORE DIRECT PATH TO CLAMP PLATE

Fig. 5. Illustration of composite hubs with and without a structure corner.

transferred to the angle plies which form the corner. This understanding was used to improve a new generation of hubs, as illustrated in Fig. 5. The delamination in the corners was eliminated. 2.4. Ply stacking sequence The ply stacking sequence plays an important role in, the strength of a structural laminate. Reducing the individual ply thicknesses and carefully dispersing plies of different orientation can minimize the Poisson’s ratio mismatch between the layers, thus reducing the interlaminar stresses. The effect of the ply stacking sequence on the strength and life of the structure is well perceived. However, designers are still often ignored. 2.5. Primary

load-carrying

ply

The 0” plies in a composite hub are the primary ones that carry the centrifugal force of the rotor blade, hence, they are often used as a continuous ply from one end of the hub to the other. They must follow a taper because of the need to control the bending stress as the flexural area approaches the hub. Angle plies, which are primarily for carrying shear, are placed between the 0” plies to create the taper. Figure 6 shows a comparison between two S2/SP250 coupons with different positions of 0” plies. The results indicate that providing a direct load path by minimizing the taper angle for the primary loadcarrying ply would increase the delamination strength.

OUTSIDE PLY DROPOFF

CL

-

*451 *451

b

0'

INSIDE PLY DROPOFF 0

t 20

I 40

60

STRESS,

KS1

60

Fig. 6. Comparison of delamination strength of ply dropoff for S2/SP250 glass/epoxy tapered laminates.

321

Rotor hub structure design 3. MATERIAL

FABRICATION

ISSUES

The weaknesses in using composites to replace metal hubs are the material allowables, material life and fabrication defects. Although these issues are not critical in most applications of composite, they must be correctly resolved for rotor hub applications. 3.1. Material

variability

Metal obtained from material manufacturers is produced by a highly controlled process, and therefore the resulting material properties are quite uniform. However, the properties of composite materials are dependent on the processing as well as the structural part configuration. Plastic matrix materials are manufactured using a wide range of chemical formulations and are available in a number of different forms (Richardson, 1987). Cured material is dependent upon the laminate and tool thicknesses. In other words, the material form is usually associated with a specific production technique. Moreover, different laminate fabrication procedures, such as bleeding and nonbleeding vacuum-bag procedures, may result in a difference in ply thickness. This results in a difference in the material constants because the percentage of fiber and matrix in a cured ply is changed. 3.2. Material

ply properties

The cured ply thickness in a structural part may be different from that of the test coupons that are used to generate the material properties. A change in ply thickness affects the material properties. Figure 7 shows a photomicrograph of a laminate crosssection of a tapered beam accomplished by ply termination. Ply terminations are used in rotor hub structures to control the hinge offset distance and to reduce the laminate bending stress at the attachment to the mast. As shown in the figure, the ply thickness near the tip of the epoxy proket area is up to 50% thicker than at other locations. If local stress behavior is desired, having accurate ply properties for use in the analysis is critical. Table 1 lists the ply properties for two different ply thicknesses. The data in the table were obtained by Chan and Ochoa (1989) using the rule-of-mixtures for the coupon ply properties and the matrix. Table 1. Variation of ply material properties AS4/3501

E, (Msi) E2 (Msi) G,, (Msi) VIZ

Matrix

i/h = 1.0

A/h = 1.25

h/h = 1.50

3501-6 resin

19.3 1.62 1.02 0.288

15.6 1.24 0.63 0.302

13.1 1.08 0.49 0.312

0.65 0.65 0.239 0.36

h = cured ply thickness in coupon, h = cured ply thickness in structure.

3.3. Characterization

of material allowables

Unlike isotropic materials, composites require two-dimensional allowables for static, dynamic, fatigue and facture toughness under different temperatures and moisture contents, i.e. those of the service environment. The introduction of fracture toughness or a primary material property for characterizing delamination is unique to flexure and thick laminates. 3.4. Marcells (ply fiber waviness) Marcells are deviations from the desired fiber orientation which are rather abrupt and of relatively short length, such as fiber waviness. They occur primarily in polarwound plies which are constructed from roving. Stiffness loss due to fiber waviness has been studied by Rai et al. (1991) and Chan and Chou (1992). Figure 8 shows an example of stiffness loss due to marcells in a given ply. The degree of marcell is quantified by the amplitude of waviness, A, divided by the length of waviness, L, i.e. L/A. As shown in the figure, the stiffness loss can be up to 15% of the laminate stiffness. Preventing the presence of marcells in composite hub fabrication has become a challenging task.

328

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s. CHAN

LAYUP [45,-45,0,90,0,90],

oZ:l 45-450

90 09090090 Location

-

AL=o.1

+

o-4545

of Waved

Ply

AJL=O.OS

*

NL=O.01

Fig. 8. Effective axial stiffness loss due to ply fiber waviness of a graphite/epoxy laminate. 4. CONCLUSION

The aim of presenting the composite issues described in this paper is to create an awareness so that the development of advanced rotor hubs can be undertaken with the minimum of risk. Although most of the issues have been under extensive research, composite designers often overlook or do not fully recognize the extent of these challenges in the development of a structurally efficient rotor hub. Acknowledgement-The sion and suggestions.

author thanks Mr C. Rogers of Bell Helicopter Textron, Inc. for his valuable discus-

REFERENCES Chan, W. S. (1991). Design approaches for edge delamination resistance in laminated composites. J. Compos. Technol. Res. 14(2), 91-96. Chan, W. S. and King, Y. F. (1991). Sensitivity analysis of delamination characterization in composite tapered beams. Proc. American Society for Composites, 6th Technical Conf., pp. 115-127. Chart, W. S. and Ochoa, 0. 0. (1989). Edge delamination resistance by a critical ply termination. In Key Engineering Materials (Edited by E. A. Armonios), Vol. 37, pp. 285-304. Traus Tech, Switzerland. Chan, W. S., Rehfield, L. W. and O’Brien, T. K. (1991). Analysis, prediction and prevention of edge delamination in rotor system structures. J. Am. Helicopter Sot. 36(2), 44-51. Chan, W. S., Rogers, C., Cronkkhite, J. D. and Martin, J. (1986). Delamination control of composite rotor hubs. J. Am. Helicoper Sot. 31(3), 60-69. Curry, J. M., Johnson, E. R. and Starnes, J. H. (1987). Effect of dropped plies on the strength of graphiteepoxy laminates. AIAA paper No. 87-0874, AIAA SDM Conf,, pp. 737-748. Fish, J. C. and Lee, S. W. (1988). Edge effects in tapered composite structures. Proc. 42nd AHS National Forum.

King, Y. F. and Chan, W. S. (1991). Delamination analysis of a composite tapered beam. Proc. 15th National Conf. on Theoretical and Applied Mechanics, Taiwan. Mayerjak, R. J. (1978). Design, fabrication and laboratory testing of a helicopter composite main rotor hub. USARTL-TR-78-16, Applied Technology Laboratory, U.S. Army Research and Technology Laboratories, Fort Eurtis, VA. O’Brien, T. K. (1990). Towards a damage tolerance philosophy for composite materials and structures. In Composite Materials.Testing and Design (ninth volume), ASTM STP 1059 (Edited by S. P. Arbo), pp. 7-13. ASTM, Philadelphia. Olster, E. E. (1980). Advanced composite rotor hub preliminary design. USARTL-TR-78-16, Applied Technology Laboratory, U.S. Army Research and Technology Laboratories, Fort Eurtis, VA. Pagano, N. J. (1989). Interlaminar Response of Composite Materials, Composite Materials Series, Vol. 5. Elsevier, Amsterdam. Phillips, N. B. and Merkey, D. J. (1990). BHTI’s technical assessment of advanced rotor and control concepts. Proc. American Helicopter Society Design Specialists’ Meeting on Vertical Lift Aircraft Design, San Francisco. Rai, H. G., Rogers, C. and Crane, D. A. (1991). Mechanics of curved fiber composites. Proc. 47th Annual Forum

of the American

Helicopter

Society.

Richardson, T. (1987). Composites: a Design Guide. Industrial Press Inc., New York. Salpekar, S. A., Raju, I. S. and O’Brien, T. K. (1991). Strain-energy-release rate analysis of delamination in a tapered laminate subjected to tension load. J. Compos. Mater. 25, 118-141.