WideEye 2Published on Monday, 27 June 2011 06:39 | Written by rcp ||| Hits: 2948

Critical part that analyzedStress analysis on critical part of the UAV is done by using MSC. NASTRAN and analyze it with MSC. NASTRAN. Distributed load is performed to find the stress on the critical part. The maximum stress found will be used to identify the structure whether the material yield stress is exceed or not when under the applied load. Critical part for this UAV is decided at the wing and the load is applied on the top surface of the wing.The analysis is done by using surface elements. Surface elements are two-dimensional elements (2-D) which used to represent a structure whose thickness is small compared to its other dimension such as plates, which are flat, or shells, which have single curvature or double curvature. For linear analysis, MSC.Nastran plate elements are assumed classical assumptions of thin plate behavior : A thin plate is one in which the thickness is much less than the next larger dimension. The deflection of the plates mid-surface is small compared with its thickness. The mid surface remains unstrained (neutral) during bending-this applies to lateral loads, not in-plane loads. The normal to the mid-surface remains normal during bending.Wing was made with Fiber Glass with Elastic Modulus of 76 000Mpa and Poisson Ratio of 0.22. Wing was tapped wing made of NACA4415 on the rib and NACA4412 on tip.MSC PATRAN Drawing and using MSC NASTRAN Analysis1. Airfoil Shape is transferred from CATIA to Patran. From CATIA file needed to save into step file first.2. Surface of the wing is drawn using the airfoil profile that is imported from the CATIA software.

3. The Surface is meshed. Output ID list: 1746 Nodes and 1217 Elements Element Shape: Quad Mesher: IsoMesh Topology: Quad4

4. Material & Properties are created. Wing is using Fibre Glass with Elastic Modulus of 76 000Mpa and Poisson Ratio of 0.22.5. Load and Constrain Condition are created. The wing is constrained on the root. Load is applied on each surface. Load calculated is discussed more detail on the Calculation of Load on Wing section below. Calculation of Load on WingTo determine distributed pressure at wing span wise, schrenks approximation method is used. From span wise distribution, pressure distribution on the surface can be obtained by dividing it to wing chord length.

Schreks loading approximation is a method commonly used to determine overall spanwise lift distribution, especially for low sweep and moderate to high aspect ratio wings. Load distribution is arithmetic mean of : load distribution representing actual plan form shape elliptical distribution of same span and areaFor elliptical load distribution equation below is used :

L = total wing lift = UAV weightL = 22.509 kNb = 13.15 my (m)0.4651.3952.3253.2554.1855.1156.045

w (kN/m)21.73921.29820.38618.93616.80913.6408.573

For trapezoidal distribution equation below is used :

y (m)0.4651.3952.3253.2554.1855.1156.045

w (kN/m)22.01620.40218.78817.17415.56013.94512.33

To get schrenks loading approximation :

y (m)0.4651.3952.3253.2554.1855.1156.045

w (kN/m)21.87820.8519.58718.05516.18513.79310.452

Then, w is divided with chord length of each y to get pressure distribution along wing chord. Chord length for each y is determined using Patran.

y (m)0.4651.3952.3253.2554.1855.1156.045

w (kN/m)21.87820.8519.58718.05516.18513.79310.452

c (m)2.1061.9361.7661.5961.4251.2531.082

P (kN/m2)10.38810.77011.09111.31311.35811.0089.660

Now pressure along wing chord should be distributed equally along the wing chord as a figure below :

So we assume pressure distribution along the wing chord should be like figure below, and its maximum pressure at 0.25c.

Calculation for y = 0.465 mArea under rectangular and area under the triangle must be equal. So,

Pressure at front surface :

Pressure at middle surface :

Pressure at back surface :

y(m)P(kN/m2)h(kN/m2)P1(kN/m2)P2(kN/m2)P3(kN/m2)

0.46510.38820.77610.38815.5404485.152448

1.39510.7721.5410.7716.111925.34192

2.32511.09122.18211.09116.5921365.501136

3.25511.31322.62611.31316.9242485.611248

4.18511.35822.71611.35816.9915685.633568

5.11511.00822.01611.00816.4679685.459968

6.0459.6619.329.6614.451364.79136

Applied load:

6. Analysis In Nastran is run after the MSC. Nastran Input is created.7. Result is created Into Patran & Fringe and Deformation Plot are created.Result1. Stress Tensor

Maximum stress found is 25.8Mpa at the wing root.2. Displacement and Translational

Maximum displacement is 0.00305mFrom the analysis, the wing of the UAV made by fiber glass is proved that the maximum Von Mises stress is 25.8MPa that is below the material yield stress for fiber glass that was 70MPa. So, the wing structure for this UAV is acceptable and satisfied the requirement to fly safely.