复合材料机身曲板环向弯曲加载试验及失效机制

陈昊; 柴亚南; 迟坚; 陆磊

doi:10.13801/j.cnki.fhclxb.20191221.001

复合材料机身曲板环向弯曲加载试验及失效机制

doi: 10.13801/j.cnki.fhclxb.20191221.001

陈昊^1, ,,
柴亚南^1,,
迟坚^1,,
陆磊^2,

1.
中国飞机强度研究所，西安 710065
2.
上海飞机设计研究院，上海 201210

基金项目: 民机科研(MJ-2016-F-01)；航空基金(20180923002)

详细信息

通讯作者:
陈昊，硕士，工程师，研究方向为复合材料结构强度　E-mail：jjocking@126.com

中图分类号: TB33
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- 被引次数: 0
出版历程
- 收稿日期: 2019-11-06
- 录用日期: 2019-12-11
- 网络出版日期: 2019-12-23
- 刊出日期: 2020-10-15

Experiment and failure mechanism of composite fuselage curved panel under circumferential bending load

CHEN Hao^{1
, ,},
CHAI Ya’nan^1
,,
CHI Jian^1
,,
LU Lei^2
,

1.
Aircraft Strength Research Institute of China, Xi’an 710065, China
2.
Shanghai Aircraft Design and Research Institute, Shanghai 201210, China

摘要

摘要: 对复合材料机身曲板进行了环向弯曲加载试验，采用四点弯加载方式对考核段进行纯弯加载，设计一种加强连接方式避免加载段提前破坏，通过试验对机身曲板的环向稳定性和破坏模式进行了分析。同时，建立了基于内聚力单元的考虑长桁与蒙皮粘接界面损伤的有限元模型，分别使用Quads准则和Hashin准则作为界面和层合板的失效判据分析曲板结构的失效机制，计算结果与试验结果吻合较好。试验及有限元分析结果表明，长桁帽底蒙皮的局部屈曲引起长桁与蒙皮粘接的R区出现初始开裂，并最终扩展为长桁脱粘。随着蒙皮屈曲及长桁脱粘的扩大，蒙皮由局部屈曲变为整体失稳而失去承载能力，最终导致隔框承载过大而发生断裂。根据初始损伤模式，采取了长桁帽内全包工艺改进设计，改进后的曲板结构稳定性和承载能力分别提高了21.9%和16.8%。
- 机身曲板 /
- 环向弯曲 /
- 四点弯曲 /
- 失稳模式 /
- 失效机制
Abstract: The failure mechanism of the composite fuselage curved panel under circumferential bending load was studied by four-point bending test and finite element analysis (FEA). A set of strengthening and connecting fixture was designed to avoid undesirable failure. A finite element model was also established, in which cohesive element was used to simulate the interface between the hat stringer and skin. Quads criterion and Hashin criterion were used as failure criteria of the interface and the laminate respectively. The results obtained by experiment and FEA are in good agreement. It can be concluded the crack initiates from the R-zone of the bonding area between the hat stringer and skin, due to local buckling of the skin at the bottom of the hat stringer. Then it extends and causes debonding of the hat stringers. With the extension of skin buckling and stringer debonding, the skin shows global instability and loses its bearing capacity, which eventually leads to the failure of the frame due to excessive load. According to the initial damage mode, the full around bonding process between the hat stringer and skin is adopted to improve the circumferential stability of the curved composite panel. The bending test result shows that the bending loads of the initial buckling and overall failure increase 21.9%and 16.8%, respectively.
- composite fuselage curved panel /
- circumferential bending /
- four-point bending /
- instability mode /
- failure mechanism

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图 1 M21E碳纤维/IMA环氧树脂复合材料机身曲板

Figure 1. M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 2 M21E碳纤维/IMA环氧树脂复合材料机身曲板四点弯加载示意图

Figure 2. Four-point-bending test schematic of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 3 M21E碳纤维/IMA环氧树脂复合材料试件加载段加强连接示意图

Figure 3. Connection schematic of loading section of M21E carbon fiber/IMA epoxy resin composite specimen

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图 4 M21E碳纤维/IMA环氧树脂复合材料机身曲板四点弯加载试验

Figure 4. Four-point-bending test of IMA carbon fiber/M21E epoxy resin composite fuselage curved panel

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图 5 M21E碳纤维/IMA环氧树脂复合材料机身曲板试件应变片粘贴位置

Figure 5. Positions of strain gauges on M21E carbon fiber/IMA epoxy resin composite fuselage curved panel specimen

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图 6 M21E碳纤维/IMA环氧树脂复合材料机身曲板四点弯曲试验加载有限元模型

Figure 6. Finite element model of four-point-bending test for M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 7 M21E碳纤维/IMA环氧树脂复合材料机身曲板界面单元模型

Figure 7. Interface element model of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 8 Cohesive内聚力单元本构关系

Figure 8. Constitutive relation of cohesive element

K_NN —Normal stiffness; K_SS—Longitudinal shear stiff-ness; K_TT—Transvers shear stiffness; N_max—Tensile strength; S_max— Longitudinal shear strength; T_max—Transvers shear strength; $\delta _{\rm{n}}^{\rm{0}}$—Original relative displacement; $\delta _{\rm{n}}^{\rm{f}} $—Relative displacement at failure; G_c— Equivalent fracture toughness

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图 9 M21E碳纤维/IMA环氧树脂复合材料机身曲板的载荷-应变曲线

Figure 9. Load-strain curves of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 10 M21E碳纤维/IMA环氧树脂复合材料机身曲板四点弯曲试验和有限元分析的载荷-位移曲线

Figure 10. Load-displacement curves obtained by tests and FEA for four-point-bending test of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 11 M21E/IMA碳纤维/环氧树脂复合材料机身曲板蒙皮与长桁脱粘破坏

Figure 11. Debonding in R-zone between skin and stringer of M21E/IMA carbon fiber/epoxy resin composite fuselage curved panel

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图 12 M21E碳纤维/IMA环氧树脂复合材料机身曲板蒙皮与长桁大面积脱粘破坏

Figure 12. Complete debonding between skin and stringers of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

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图 13 M21E碳纤维/IMA环氧树脂复合材料机身曲板在环向纯弯载荷下的损伤演化规律

Figure 13. Damage evolution of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel under circumferential bending load

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图 14 M21E碳纤维/IMA环氧树脂复合材料帽底蒙皮载荷-应变曲线的有限元计算结果

Figure 14. Load-strain curves of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel skin under stringers obtained by FEA

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图 15 原试件长桁脱粘模式

Figure 15. Hat stringer debonding mode of original specimen

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图 16 帽型长桁全包示意图

Figure 16. Diagrammatic sketch of full around bonding hat stringer

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图 17 全包试件长桁脱粘模式

Figure 17. Hat stringer debonding mode of full around bonding specimen

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图 18 全包试件破坏模式

Figure 18. Failure mode of full around bonding specimen

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表 1 M21E碳纤维/IMA环氧树脂复合材料性能

Table 1. Material properties of M21E carbon fiber/IMA epoxy resin composite

Elastic property	Value	Strength property	Value
E₁₁/GPa	154	X_T/MPa	2 610
E₂₂/GPa	8.5	X_C/MPa	1 450
G₁₂/GPa	4.2	Y_T/MPa	55
ν₁₂	0.35	Y_C/MPa	285
		S₁₂/MPa	105
Notes: E₁₁—Longitudinal tensile modulus; E₂₂—Transverse tensile modulus; G₁₂—Shear modulus; ν₁₂—Poisson’s ratio; X_T—Longitudinal tensile strength; X_C—Longitudinal compressive strength; Y_T—Transverse tensile strength; Y_C—Transverse compressive strength; S₁₂—Shear strength.

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表 2 M21E碳纤维/IMA环氧树脂复合材料试件各部件铺层顺序

Table 2. Layup orientation of M21E carbon fiber/IMA epoxy resin composite parts

Part	Patch layup
Skin	[45/−45/−45/90/45/0]_s
Shear-clip	[45/−45/0/90/−45/45/0/0/45/−45/90/0/−45/45]
Stringer	[45/0/0/−45/90/−45/0/0/45]
Frame	[45/−45/0/90/45/−45/90/0/−45/45]

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表 3 M21E碳纤维/IMA环氧树脂复合材料机身曲板界面性能参数

Table 3. Material properties of adhesive of M21E carbon fiber/IMA epoxy resin composite fuselage curved panel

Parameter	Value
Interface strength /MPa	N_max	S_max	T_max
Interface strength /MPa	40	65	65
Fracture Toughness /(kJ·mm⁻²)	G_ⅠC	G_ⅡC	G_ⅢC
Fracture Toughness /(kJ·mm⁻²)	0.37	0.6	0.6
Elastic constant /(GPa·mm⁻¹)	K_NN	K_SS	K_TT
Elastic constant /(GPa·mm⁻¹)	4 324	7 027	7 027

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