STATIC TESTING AND ANALYSIS OF COMPOSITE WING OF A TWO-SEATER AIRCRAFT
POWERED BY Li-ion BATTERY ELECTRIC PROPULSION
Kang Yang, Liguo Zhang*, Shude Ji, Yumei Yue, Wang Ji
(Shenyang Aerospace University, Liaoning General Aviation Academy, Shenyang, China, 110136) *Author to whom correspondence should be addressed: Tel: +86-024-88795091, E-mail:[email protected]
Received 2 November 2016; accepted 21 November 2016 ABSTRACT
The present paper studied the strength and deformation characteristics of a composite wing of a two-seater aircraft powered by Li-ion battery electric propulsion. In order to get the airworthiness certificate, the static testing is indispensable. The primary wing structure component includes upper and lower skins, leading edge, trailing edge, a root rib and main spar. The main purpose of static testing is to examine the bending strength of the wing. The testing results are in good agreement with the FE analysis results and the bending strength of wing is strong enough to support the limit loads (the maximum loads to the ex-pected in service) without detrimental and permanent deformation, and without failure for at least 3s under the ultimate loads (limit loads multiplied by prescribed factors of safety), which meets ASTM F2245-11(Standard Specification for Design and Performance of a Light Sport Airplane) enough.
Keywords: Static testing; composite wing; Li-ion battery; Electric propulsion; Two-seater aircraft; ASTM F2245-11
1. INTRODUCTION
With the constraint of traditional energy and the presen-tation of environmental problems, the pollution-free and renewable energy has attracted more and more attentions. In the field of aeronautical engineering application, the electric-driven aircraft is rapidly developed. And the fuel cell and Li-ion battery have been widely used [1-5]. Romeo et al. [1] presented the setting up and test flights of a fuel cell general aviation aircraft fuelled by hydro-gen. Romeo et al. [2] developed and validated the use of a fuel cell based power system for the propulsion of more/ all electric aircraft. Sylvie et al. [3] presented the fuel cell climatic tests designed for new configured aircraft applica-tion. Akira et al.[4] described the measurement results of a passive hybrid system consisting of fuel cell stacks, Li-ion battery packs, and two diodes. Kamen et al.[5] presents the PL55E cell developed and the performance results of the PL55E delivered by SAFT America. However, Li-ion battery is better to form the viewpoints of non-pollution. With the rapid development of Li-ion battery technologies which are more green energy sources, Li-ion battery have been used in the aerospace, marine, and civil engineering due to its superior properties. Engines that use the com-bustion of fuel produce a lot of carbon dioxide. Carbon dioxide directly or indirectly causes the greenhouse effect. So it is necessary for a general aviation aircraft powered by Li-ion battery.
Because of the fact that the power of electric engine is less than that of the oil engine, the aircraft structure is required to be lighter and have higher lift-drag ratio to improve the cruise time and distance. Therefore, the wing is very im-portant in the design of electric aircraft.
In order to ensure flight safety, the static structure testing of the wing structure of any aircraft should be performed, which can verify the strength and stiffness of the wing structure under the design of load provision. Smith et al.[6] described the static testing of an ultralight airplane and replaced with masses to fulfill centre-of-gravity require-ments before testing. Wong et al. [7] presented the design of the test rig and the results of the static test of the Lock-heed P-3 Orion Wing Leading Edge centre section struc-ture. Sullivan et al.[8] presented the results of the strength and stiffness characteristics of a carbon composite wing of an ultralight unmanned aerial vehicle. Sullivan et al.[9] described vibration testing of a full-scale carbon compos-ite ultralight unmanned aerial vehicle. These researches are only for the design and test of composite materials of unmanned aerial vehicles and do not involve the airworthi-ness static testing of the composite wing of a two-seater aircraft powered by Li-ion battery electric propulsion [6-9].
On the basis of a self-designed two-seater aircraft powered by Li-ion battery electric propulsion, the paper describes the static testing and finite element (FE) analysis results of the plane. The type of load is symmetrical wing load, which is specified regarding the limit loads (the maximum loads to be expected in service) and the ultimate loads , which is the limit loads multiplied by prescribed factors of safety. The ultimate load safety factor is 1.5. The prop-erty of material system, the result of finite element (FE) analysis, the static loading method and the testing result are described.
2. THE WING STRUCTURE
two-seater aircraft powered by Li-ion battery electric propul-sion, and the plane uses an all-composite structure which much lighter than traditional metal. The plane structural components are designed using ASTM F2245-11[10], which are the world’s first two-seater electric light sport aircraft conforming to the definition with ASTM 2245-11 and the China’s first electric light sport aircraft with com-pletely independent intellectual property rights. As the first electric light sport aircraft certificated by the Civil Aviation Administration of China, in order to verify the strength and deformation characteristics of the wing structural as-semblies, the wing structural static testing was performed. Composite materials have been widely utilized in aircraft due to their high specific strength as compared to conven-tional isotropic materials. The two-seater aircraft is pow-ered by Li-ion battery and made by composite materials. The aircraft uses electric motor as power plant and maxi-mum take-off weight is 480kg.
The wing span of this plane is 14.5m, the mean aerody-namic chord is 0.868m. The distance and maximum thick-ness from wing root to root rib are 1.17m and 47.4mm, respectively. As shown in Fig. 2, the primary wing struc-ture includes the upper and lower skins, a root rib, lead-ing edge, traillead-ing edge, main spar and spoiler houslead-ing, and these structures are glued by adhesive. Moreover, both the ailerons and spoiler are removed in finite element (FE) analysis and static testing.
The skins are made of a sandwich material. The spars and root rib are made of carbon composite. In order to pro-vide the wing bending stiffness and form the carry-through structure, the main spar extends inside the fuselage. Four steel pins on each fuselage side bear the shear load and tor-sion, and the root-rib/fuselage interconnection is formed by four pins. The poles prevent the top of the fuselage from deforming, as shown in Fig.3.
The wing structure components are made of carbon fibre (TENAX W-3161), glass fibre (SW110C-100A) and uni-directional prepreg fabric (UIN 46200), and the material properties are shown in Table 1. The upper and lower skins are made of sandwich construction with a 4mm foam core (H60 4 PSC). The laminate ply pattern in the wing region W for every member is given in Table 2. The steel lift pins are mounted on the root rib in Fig. 3.
Fig. 1: CATIA model of the plane
Fig. 2: The components of wing structural
Material
property carbon fibre
(TENAXW-3161)
glass fibre
(SW110C-100A) unidirectional prepreg fabric
(UIN 46200)
Foam core (H60 4 PSC)
E11,psi 5.0E10 2.38E10 1.26E11 1.233E4
E22,psi 5.0E10 2.38E10 0.23E8 --
G12,psi 0.95E10 0.71E10 0.23E8 --
V12 0.052 0.116 0.30 0.32
Description NO. of plies Ply pattern (a)
Lower/upper skin 7 s45/[t0]3/foam/t0/s0
Leading edge 2 [t0]2
Trailing edge 4 t0/[t45]2/t0
Root rib 14 [t0/t45]2/[t0]10/[t0/t45]2
Lower/upper flange 33 [u45]/[u(-45)]/[u0]29/[u(-45)]/[u45]
web 44 [t0]8/[foam4]2/[t0]20/[t45]13/foam
a. 45 is +-45 fabric,0 is 0/90 fabric, s is glass fibre, t is carbon fibre, u is unidirectional prepreg. Table 1: Properties of materials
Table 2: Laminate definitions of the wing region W
3. EXPERIMENT
In order to more actually reflect the wing structure load, a test frame was designed to serve as the support structure for static strength testing, as shown in Fig.4. In the ex-periment, the load distribution is simulated using iron sand pockets on the wing, which are commonly used for the structural static testing of ultralight aircraft.
Fig. 4: A universal test frame for static strength testing
NO. The first time (10%) The second time (20%) The third time (30%) The fourth time (40%) The fifth time (50%) The sixth time (60%) The seventh time (67%) The eighth time (75%) The ninth time (80%) The tenth time (90%) The eleventh time (100%) 0 13.7 27.5 41.2 54.9 68.7 82.4 92.0 103.0 109.9 123.6 137.4 1 14.4 28.7 43.1 57.4 71.8 86.1 96.2 107.6 114.8 129.2 143.5 2 12.7 25.5 38.2 51.0 63.7 76.5 85.4 95.6 102.0 114.7 127.5 3 12.5 24.9 37.4 49.8 62.3 74.7 83.4 93.4 99.6 112.1 124.5 4 13.5 27.1 40.6 54.2 67.7 81.3 90.8 101.6 108.4 121.9 135.5 5 12.6 25.1 37.7 50.2 62.8 75.3 84.1 94.2 100.5 113.0 125.6 6 11.4 22.7 34.1 45.4 56.8 68.1 76.1 85.2 90.9 102.2 113.6 7 9.9 19.8 29.8 39.7 49.6 59.5 66.5 74.4 79.4 89.3 99.2 8 8.1 16.2 24.3 32.4 40.5 48.6 54.3 60.8 64.9 73.0 81.1 9 5.1 10.2 15.4 20.5 25.6 30.7 34.3 38.4 41.0 46.1 51.2
Table 3: Force distribution at wing loading locations
3.1. Loading Method
Each wing has ten loading locations and the skin has been already described and the right wing is marked (from RS1 to RS10), as shown in Fig.5. The distance of each loading location away from the fuselage centreline (unit: mm) is shown in Fig.5. The loading at each loading location are under the limit load and ultimate load respectively. Table 3 shows the loading of the right wing and the total number of load cases. The left wing is symmetrical in the experi-ment. In order to ensure that the two wings are loaded at the same time, two cranes are used to prop the wing at each side before loading. After the iron sand pockets are loaded at the rational locations, the two cranes slowly put down the two wings at the same time, by which unsymmetrical wing loads are avoided.
3.2. Measurement Instrumentation
The right wing is fitted with eight strain gauges at four different wing locations, and all strain gauges are placed
along the wing span. On the upper skin, there are four strain gauges (from GSU1 to GSU4) located along the main spar and the strain gauges of the lower skin (from GSL1 to GSL4) are symmetrical, as shown in Fig.5. The strain gauges are used to measure the strain and the gen-eral-purpose strain gauges (BF350-5AA(11)N6-X) have a nominal gauge length of 6.4mm, resistance is 349.8±0.2 and gauge factor is 2.10±1. All strain gauges have been tested before bonding the upper/lower skin to complete the static testing.
In order to obtain the wing deformation a ruler near each wing tip is placed during the static testing process. The ruler readings of every load case are recorded and then the deformations of the two wing tips are got. In order to collect strain data, a computer data acquisition system is used. The first and last record of the wings are measured under the condition of unloaded. The difference between ultimate load and limited load is used to describe the tip deformation of δL and δU, and the values are shown in Table 4.
Fig. 5: Locations of loading locations, strain-gauge locations.
Measure location δL mm δU mm δL/S δU/S
Left wing tip 880 1275 0.121 0.176
Right wing tip 895 1295 0.125 0.178
Table 4: Measured wing deformations and strength
4. FINITE ELEMENT ANALYSIS 4.1. The model of finite element
In order to predict the wing structural deformation and strains under the critical loading conditions, MSC.Patran/ Nastran is used to establish FE model. The model is es-tablished according to the actual wing structure. The con-tinuum elements are used to the adhesive layer between structural elements (spar, root-rib/skin/spar, and skin/skin interconnections). Moreover, in order to simulate bolt con-nections, appropriate multiple-point constraints (MPCs) are used to the two main spars. And all structural elements adopt Tri/Quad element, as shown in Fig.6. The FE model contains 1,226,601 elements and 2,539,067 nodes.
Fig. 6: Finite element mesh with boundary conditions 4.2. Boundary Conditions and Loading
The boundary conditions agree with the static structural testing as identified in Fig. 6.The aft, outboard, and up di-rections define as the X, Y and Z coordinates in Fig. 6, respectively. In the model in Fig.6, three nodes are used to describe the landing gear(include main/nose landing gear)/ fuselage fixture interconnection, while the displacements of nodes are fixed (UX=UY=UZ=0). Uniform normal pressures are considered as the resultant force at each load location, which is applied to the lower skin.
4.3 Static Analyses Method
To predict wing deformation as well as laminate strains under incremental loading to design ultimate load, geo-metric linear static FE analyses are performed. In order to identify the potential failure zones of the wing structure, the deformation and strain distributions are used.
5. PRESENTATION OF RESULTS
Table 4 shows under ultimate loads the values for the per-cent relative tip deformation, namely the right wing tip (δR/S) and the left wing tip (δL/S). Both two wings do not appear detrimental and permanent deformation under limit loads, while the right-wing tip and left-wing tip
de-formations are 895 mm (0.125S) and 880 mm (0.121S), respectively. When the loads reach 75 of the ultimate load, the upper skin near the spoiler housing occurs sunken and hump, as shown in Fig.7. Under the ultimate load, there are no other failures to appear the wing, while the right-wing tip and left-wing tip deformations are 1295 mm (0.178S) and 1275 mm (0.176S), respectively. So there is no shear deformation of the two wings and the static testing of sym-metrical wing loads is reasonable. It is conclude that the bending strength of wing is strong enough to support ulti-mate loads without failure for at least 3s.
Fig. 7: Photographs of the upper skin near the spoiler housing occur sunken and hump.
Fig. 8: Photographs of a) and b) shows the measured and pre-dicted strain of the upper/lower skin at limit load and ultimate
load.
Fig.8 shows the measured and simulated strain of the upper skin of the right wing under limit loads and ultimate loads. At these locations, the FE predictions result very well with the experimentally measured values. The present analysis method of composite wing has the good applicable value. 6. CONCLUSIONS
Static testing of composite wing of a two-seater aircraft a) b) -6000 -4000 -2000 0 2000 4000 6000
GSU1/GSL1 GSU2/GSL2 GSU3/GSL3 GSU4/GSL4
st rai n( uε ) limit load upper measured upper predicted lower measured lower predicted -20000 -15000 -10000 -5000 0 5000 10000 15000 20000
GSU1/GSL1 GSU2/GSL2 GSU3/GSL3 GSU4/GSL4
st rai n( uε ) ultimate load upper measured upper predicted lower measured lower predicted
powered by Li-ion battery electric propulsion was pre-sented. A test frame was designed to serve as the support structure for the static testing. The deformation of wing tip and measured strain data were collected by computer. Af-ter unloading from the limit loads (the maximum loads to the expected in service) and the ultimate loads (limit loads multiplied by prescribed factors of safety), the failure loca-tion and the all wing structures are recovered approximate-ly. In our opinion, the wings can support the limit loads and ultimate loads, and after unloading without appearing of permanent deformation and structural failure, so the wings can support ultimate loads condition at 6g without failure and content the item of the ASTM F2245-11.
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